ASME Conference Presenter Attendance Policy and Archival Proceedings

2017;():V001T00A001. doi:10.1115/GTINDIA2017-NS1.

This online compilation of papers from the ASME 2017 Gas Turbine India Conference (GTINDIA2017) represents the archival version of the Conference Proceedings. According to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference by an author of the paper, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.

Commentary by Dr. Valentin Fuster

Compressors, Fans and Pumps

2017;():V001T01A001. doi:10.1115/GTINDIA2017-4516.

A novel infill criterion for so-called ensemble of surrogates-based optimization is proposed and applied in practice for an aerodynamic compressor rotor design optimization. The ensemble uses a combined approach based on different radial basis functions and aims to reduce prediction errors through weighted linear combinations of radial basis functions. The update strategy uses a new hybrid custom metric termed α, which incorporates information about each surrogate’s local agreement through correlation coefficients and also information about the global accuracy of each ensemble combination through the root-mean-square error. Surrogate models are searched using a hybrid optimizer, i.e., with a genetic algorithm and sequential quadratic programming, and proposed update points are evaluated using the high-fidelity black box function. The results are compared with established optimization approaches and the best design is analyzed further in terms of the flow physics. Results show that α-based ensemble of surrogates approaches are particularly efficient for large-scale cases, where other types of surrogates such as Kriging models are onerous to construct.

Commentary by Dr. Valentin Fuster
2017;():V001T01A002. doi:10.1115/GTINDIA2017-4518.

Performance monitoring of wet gas compressors is challenging due to the liquid phase impact on performance. Introduction of a liquid phase alters both the thermodynamics as well as the fluid dynamics of the compression process. Hence, understanding the flow interaction between the impeller, diffuser and volute is pivotal. Previous investigations have detected occurrence of compressor hysteresis at certain wet gas operating conditions, resulting in temporary deviations from the steady state compressor characteristics. This kind of behavior influences both the compressor stability and performance. Thus, being able to understand the onset of hysteresis and its impact on the compressor is paramount.

An experimental test campaign has been performed at the Norwegian University of Science and Technology (NTNU). The test facility is an open loop configuration consisting of a shrouded centrifugal impeller, a vaneless diffuser and a circular volute. The current investigation document the compressor performance shift and the occurrence of compressor hysteresis when gradually increasing the liquid load on a centrifugal compressor. Emphasis was put on the compressor performance and its correlation to the diffuser multiphase flow regime. The investigation revealed that there is a clear dependence between the diffuser multiphase flow characteristics and the compressor performance.

Topics: Gas compressors
Commentary by Dr. Valentin Fuster
2017;():V001T01A003. doi:10.1115/GTINDIA2017-4528.

External gear pumps are typically used in aero-engines for the fuel and lubrication system due to its simplicity in construction. The design of the gear pump has been considerably improved over several years by including design features to improve its overall performance and reliability. In this paper, three-dimensional numerical analysis of an external gear was carried out by including design features such as scallops at the inlet and outlet, radial and axial clearances, journal bearing clearances and the axial tilt of the supporting bushes. The Immersed Solid Method (ISM) is used to analyze the gear pump at different operating conditions. The applicability of different turbulence models to the Immersed solid method is discussed. The internal flow features are discussed and compared with the results available in the literature. The Pump characteristics curve developed from the numerical analysis using the Immersed solid method (ISM) is compared with the experimental test results.

Commentary by Dr. Valentin Fuster
2017;():V001T01A004. doi:10.1115/GTINDIA2017-4529.

The world’s energy demand is increasing, asking for new and cost-efficient ways to extract oil and gas. With traditional technologies, oil and gas production relies on a sufficiently high well head pressure for transportation to nearby process facilities. Utilization of subsea wet gas compression systems enables production at significantly lower pressures and is a favourable solution concerning production in remote regions.

Wet gas compressors are particularly useful when handling multiphase mixtures consisting of 95%–100% gas, on a volumetric basis. The remaining content is water and liquid condensate, which introduces flow mechanisms such as droplet deposition, liquid film formation and momentum transfer, which influence the fundamental flow behavior through the compressor. Previous tests have documented the occurrence of compressor hysteresis at low compressor flow rates. Recent findings have revealed the flow interaction between the diffuser and the volute is a governing factor concerning the documented hysteresis. This kind of behaviour induces challenges with regard to compressor performance prediction and securing stable operation.

An experimental test campaign has been performed at the Norwegian University of Science and Technology (NTNU). The test facility is an open loop configuration consisting of a shrouded centrifugal impeller, a vaneless diffuser and a circular volute. The test was performed by establishing the compressor characteristics while monitoring the diffuser/volute flow regime. Emphasis was put on the volute flow characteristics and the correlation with the compressor performance.

The investigation reveals that the volute flow characteristics and the interaction with the diffuser has a distinct impact on the compressor performance, particularly at lower gas mass fractions. Furthermore, the test reveals that the diffuser design is a key factor concerning the performance impact.

Commentary by Dr. Valentin Fuster
2017;():V001T01A005. doi:10.1115/GTINDIA2017-4531.

This paper describes an improved throughflow calculation method on S2m based on streamline curvature method for predicting the performance of centrifugal compressor. A general method of specifying the empirical data provides separate treatment of blockage, deviation and losses. The spanwise and streamwise distribution laws of losses are described. The paper describes a new aspect of method about the mixing loss. Two-zone model considering the “jet and wake” can obtain the secondary flow width. For this reason, the improved prediction method combined with two-zone model is proposed to correct the mixing loss. Due to the average static pressure at outlet unknown, the secondary flow width is obtained by iterations.

This performance prediction method is validated with experimental and CFD data of three cases, including impeller(A), impeller(B) and impeller(C). The results show that the improved throughflow calculation method predicts the performance of centrifugal compressor more accurately than conventional throughflow calculation, with increased the accuracy of total pressure ratio and isentropic efficiency by about 3.18% and 1.30%.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2017;():V001T01A006. doi:10.1115/GTINDIA2017-4587.

Usually at high angle of attack, aerofoil stalls due to flow separation on suction surface of aerofoil. To delay the flow separation, pulsating jet arrangement, known as Synthetic jet is used in aerofoil. It is produced by periodic suction and ejection of fluid from an orifice. This condition can be achieved by inducing movement to diaphragm or by giving a zero mass flux sinusoidal boundary condition to the jet. This allows the reattachment of boundary layer which improves the lift and drag performance and angle at also delays stalling angle. In present study, CFD analysis on NACA0015 aerofoil is performed for different angles of attack and the Co-efficients of Drag (Cd) and Lift (C1) are validated with the experimental results of Gilarranz et al. [1]. The flow is simulated by solving Unsteady RANS coupled with k-ε realizable turbulence model with enhanced wall treatment. Synthetic jet is placed in NACA0015 airfoil at 12% of the chord length with width as 0.53% of chord and is studied for a Reynolds number Re = 8.96 × 105 and for angle of attack from 12 to 20 degrees [2]. The jet is almost tangential to the wall at an angle, αjet = 10° and chord length is considered as 0.375m for the study. Further, parametric analyses are conducted on NACA 0015 aerofoil to investigate effect of parameters (frequency, jet angle, jet velocity). It is observed that aerofoil’s performance is improved significantly for jet angle (30°–40°), jet frequency (100 Hz) and non dimensional jet velocity (1.8–2.0). A maximum increase of approximately 26% in Lift was observed at AOA 20°.

Commentary by Dr. Valentin Fuster
2017;():V001T01A007. doi:10.1115/GTINDIA2017-4592.

A transonic axial flow compressor undergoes severe vibrations due to instabilities like stall and surge when it operates at lower mass flow rate in the absence of any control devices. In present study, the attempt was made to understand the combine impact of circumferential casing grooves (CCG) of constant aspect ratio and different axial spacing between rotor and stator on the operating stability of single stage transonic axial compressor and that of rotor alone using numerical simulation. The optimum rotor-stator gap in the presence of grooved casing treatment was identified. The steady state numerical analysis was performed by using three-dimensional Reynolds Average Navier-Stokes equation adapting shear stress transport (SST) k-ω turbulence model. The study is reported in two sections. First section includes the detailed numerical study on baseline case having smooth casing wall (SCW). The computational results were validated with the experimental results available at Propulsion Division of CSIR-NAL, Bangalore. The computational study shows good agreement with experimental results. The second section comprises the effects of optimum designs of CCG and various axial spacing on the stall margin improvement of transonic compressor. Current computational study shows that the axial spacing between rotor and stator is an important parameter for improvement in stall margin not only for SCW but also for CCG. Therefore, the highest stall margin improvement of 9% has achieved for 75% axial spacing.

Commentary by Dr. Valentin Fuster
2017;():V001T01A008. doi:10.1115/GTINDIA2017-4594.

Application of surface roughness to rotating mechanical bodies will result into performance degradation. In Aviation Industry, one of the most affecting causes for performance or efficiency degradation of gas turbine engine is the blade surface roughness. The aerosols which are very small particles in the atmosphere having diameters in the microns, impinges to the compressor blade inside the aircraft engine at higher altitudes. The aerosols damages surfaces of the compressor blades. Despite of having small dimensions, due to higher velocity of the aircraft, aerosol’s impinging creates roughened surfaces and fouling. This paper is an attempt to numerically evaluate the performance degradation of the single stage transonic axial flow compressor due to uniform roughness created by the aerosols. Various cases with different roughness on various sections of the blades are analyzed to study and identify which section of the blade is more influenced by roughness. The transonic axial flow compressor has a capability of producing 1.36 stage total pressure ratio, swallowing air mass flow rate of 23 kg/s at rated design speed of 12930 rpm is used for the steady state numerical analysis. A systematic steady state 3-dimensional numerical study using solver with SST k-ω turbulence model has been carried out to evaluate the impact of blade surface roughness on the performance of compressor stage. Moreover, cases with the aerosols having different dimensions and their resulting effect is also studied to find out how performance varies when the aircraft enters into atmosphere having big aerosols from the atmosphere having smaller one and vice-e-versa.

Commentary by Dr. Valentin Fuster
2017;():V001T01A009. doi:10.1115/GTINDIA2017-4625.

Efficiency of the centrifugal compressor is affected by non-uniform flow at the exit of the impeller and the losses in the diffuser. This causes a significant loss of total pressure and drop in the performance of a centrifugal compressor. By rotating some portion of stationary vaneless diffuser walls with the speed of the impeller, the shear forces between the flow and diffuser walls are greatly reduced. Thereby improvement in the operating range and performance is achieved. This paper presents CFD studies on the effect of different single wall rotations i.e. hub rotation and shroud rotation of the vaneless diffusers on the overall performance at 10% and 15% extension of impeller walls. It is observed that the performance characteristics of compressors with all RVD models offer higher static pressure recovery and also higher rate of diffusion compared to the base compressor with SVD. It is clear that as extended radius increases from 10% to 15%, substantial improvement of efficiency and reduction of losses are observed for both type of models. Out of two single wall rotation models, SRVD model is able to better mix the jet-wake type of impeller exit flows and minimizes the losses therein and improve the performance of the centrifugal compressor.

Commentary by Dr. Valentin Fuster
2017;():V001T01A010. doi:10.1115/GTINDIA2017-4634.

The introduction of wet gas compression provides the opportunity for future cost-effective production of oil and gas. A wet gas compressor consists of a robust unit able to increase the pressure of untreated natural gas. This permits longer transport of hydrocarbons without topside facilities if installed at the well head. Obvious benefits include prolonging the life of existing wells and the possibility of exploiting smaller hydrocarbon sources otherwise considered non-commercial.

Successful development of robust wet gas compressors requires further understanding of the phenomena which occur when liquid is present in the gas stream. Understanding the way the presence of liquid affects the velocity triangle and slip factor is essential for the design of wet gas compressors and for comprehending their response to varying levels of liquid content in the inlet stream.

An experimental study has been performed with various levels of liquid fractions and inlet swirl angles. Impeller-exit velocity components and shift in slip factors are presented within the experimental test boundary.

A shift in velocity components and slip factor is experienced with increasing liquid content and inlet guide vane (IGV) setting angle. Consequently, existing slip factor correlations not utilizing inlet flow characteristics are not valid for wet gas flow or with impeller inlet swirl.

Commentary by Dr. Valentin Fuster
2017;():V001T01A011. doi:10.1115/GTINDIA2017-4686.

Winglets are plane surfaces with certain thickness and different shapes. Winglets are used in aircraft to reduce wing tip vortex which is created due to differential pressure in between pressure surface and suction surface. In transonic axial compressor, rotor tip leakage vortex interaction with shock layer and shroud boundary layer leads to total pressure loss and initiation of stall phenomenon. Effect of tip winglets are investigated in compressor rotor cascade. Cascade investigation shows that rotor tip winglet are able to reduce total pressure loss due to tip leakage flow and blade passage secondary flow. Cascade studies are performed with winglet on blade suction side, pressure side and combination of both. From cascade studies it is revealed that suction side winglet are aerodynamically better than pressure side and combined winglets. Owing to favorable results of tip winglet on compressor cascade performance, it was assumed that tip winglets would enhance overall performance of transonic compressor stage with rotating rotor. Results of present CFD simulations have predicted both positive and negative effects of winglets. Effect of different winglet configurations on pressure side and suction side of rotor blade tip are investigated to analyze the compressor stage overall performance. Rotor tip winglets are able to increase stage total pressure ratio compare to the baseline stage without winglet. Stage with winglets have shown better performance in choke region. Winglets are able to vary rotor blade loading from hub to tip region. Presence of winglet has shown ability to reduce to total pressure loss in trailing edge wake region. Stall margin is decreased in compressor stage with winglets due to more blockage towards trailing edge in tip region.

Commentary by Dr. Valentin Fuster
2017;():V001T01A012. doi:10.1115/GTINDIA2017-4704.

Numerical investigation is carried out on a low-pressure ratio centrifugal compressor stage to study the effects of the rotational speed of a rotating vaneless diffuser on flow diffusion using various flow parameters and performance characteristics parameters. The results obtained are compared with a stage having conventional stationary vaneless diffuser. Rotational speed of the rotating vaneless diffuser plays a major role in determining the extent of net gain in energy level of the fluid and drop in stagnation pressure losses. The net gain in energy level result as rise in kinetic energy level of the fluid. By an effective diffusion process, this results into an improved static pressure and stagnation pressure distribution at stage exit and FreeRVDSR0.75 undergoes a comparatively better diffusion process. Based on this study, it can be concluded that the diffusion process efficacy of a compressor stage with rotating vaneless diffuser is better in the free type at diffuser’s rotational speed above 0.50 times the impeller’s rotational speed.

Commentary by Dr. Valentin Fuster
2017;():V001T01A013. doi:10.1115/GTINDIA2017-4753.

High speed centrifugal compressors are used in turbochargers and in small gas turbine engines that typically power cruise missiles, helicopters and auxiliary power units (APU). Centrifugal compressors have wider operating range compared to axial compressor and are compact. Though centrifugal compressors having a pressure ratio of the order of 12:1 per stage have been demonstrated with reasonably good isentropic efficiencies, achieving a wider operating range has always been a challenge. A Turbocharger that needs to be designed to function both at sea-level and 5 km altitude conditions, requires a wider compressor map to accommodate the diesel engine operating line. A wider compressor map can be achieved by various techniques. The approaches used in the present study include providing pinch in the diffuser entry region and ported shroud arrangement in the compressor casing. A parametric study has been carried out by varying geometric parameters and an appropriate configuration that offers lower total pressure loss and better diffuser pressure recovery is chosen. The flow mechanisms responsible for better performance is investigated numerically for various configurations with diffuser pinch. To further enhance the operating range, a ported shroud configuration in the compressor housing is designed and analysed with the finalized diffuser pinch. Results of computational analysis for different ported shroud slot geometries have been studied numerically and are presented. Two configurations have been tested in a turbo-drive based test rig. The first configuration is only with diffuser pinch and the second configuration is with diffuser pinch and ported shroud. The extent of map width enhancement obtained by each technique is presented and compared with numerical analysis. The test results show good match with the predicted trend and confirms that diffuser pinch and ported shroud configurations offer significant enhancement in achieving a wider operating range. The flow mechanisms responsible are discussed in detail in the paper.

Commentary by Dr. Valentin Fuster
2017;():V001T01A014. doi:10.1115/GTINDIA2017-4762.

This paper describes a methodology for obtaining correct blade geometry of high aspect ratio axial compressor blades during running condition taking into account of blade untwist and bending. It discusses the detailed approach for generating cold blade geometry for axial compressor rotor blades from the design blade geometry using fluid structure interaction technique. Cold blade geometry represents the rotor blade shape at rest, which under running condition deflects and takes a new operating blade shape under centrifugal and aerodynamic loads. Aerodynamic performance of compressor primarily depends on this operating rotor blade shape. At design point it is expected to have the operating blade shape same as the intended design blade geometry and a slight mismatch will result in severe performance deterioration. Starting from design blade profile, an appropriate cold blade profile is generated by applying proper lean and pre-twist calculated using this methodology. Further improvements were carried out to arrive at the cold blade profile to match the stagger of design profile at design operating conditions with lower deflection and stress for first stage rotor blade. In rear stages, thermal effects will contribute more towards blade deflection values. But due to short blade span, deflection and untwist values will be of lower values. Hence difference between cold blade and design blade profile would be small. This methodology can especially be used for front stage compressor rotor blades for which aspect ratio is higher and deflections are large.

Commentary by Dr. Valentin Fuster
2017;():V001T01A015. doi:10.1115/GTINDIA2017-4767.

This study discusses in detail the aeroelastic flutter investigation of a transonic axial compressor rotor using computational methods. Fluid structure interaction approach is used in this method to evaluate the unsteady aerodynamic force and work done of a vibrating blade in CFD domain. Energy method and work per cycle approach is adapted for this flutter prediction. A framework has been developed to estimate the work per cycle and aerodynamic damping ratio. Based on the aerodynamic damping ratio, occurrence of flutter is estimated for different inter blade phase angles. Initially, the baseline rotor blade design was having negative aerodynamic damping at part speed conditions. The main cause for this flutter occurrence was identified as large flow separation near blade tip region due to high incidence angles. The unsteadiness in the flow was leading to aerodynamic force fluctuation matching with natural frequency of blade, resulting in excitation of the blades. Hence axially skewed slot casing treatment was implemented to reduce the flow separation at blade tip region to alleviate the onset of flutter. By this method, the stall margin and aerodynamic damping of the test compressor was improved and flutter was avoided.

Commentary by Dr. Valentin Fuster
2017;():V001T01A016. doi:10.1115/GTINDIA2017-4773.

Study of aerodynamic flow and aeroelastic stability in vibrating blades of cascade is the main objective of this study. Standard test configuration (STC-5) was chosen for this study as it involves transonic flow regime in compressor blade cascades. CFD analysis were carried out for 11 test cases of STC-5 configuration and pressure coefficient values were compared with test data. The range of incidence angles vary from 2° to 10° and reduced frequency varies from 0.14 to 1.02. Inflow Mach number was fixed at 0.5 and Reynolds number was fixed at 1.4 × 106. Analysis of vibrating blades and comparison of test data results of axial compressor with linear cascade stator blades of fifth standard configuration at high subsonic speed is compared with CFD results. While doing this vibration of only the center blade is concerned when all the other blades in the cascade are fixed. Fluid structure interaction approach is used here to evaluate the unsteady aerodynamic force and work done for a vibrating blade in CFD domain. Energy method and work per cycle approach is adapted for aerodynamic damping prediction. A framework has been developed to estimate the work per cycle and aerodynamic damping ratio. Final sensitivity study was carried out to evaluate the influence of blade incidence and frequency on blade damping values.

Commentary by Dr. Valentin Fuster
2017;():V001T01A017. doi:10.1115/GTINDIA2017-4848.

An electric submersible pump (ESP) is a multistage centrifugal pump widely used in the petroleum industry to transport wellbore fluids to the surface. For a turbulent flow through a pump, the surface roughness plays a vital role as it causes flow separation and increases boundary layer momentum loss. In the present work, a 3D numerical analysis by solving a Reynolds-averaged Navier-Stokes equation with k-ω SST turbulence model for wall bounded steady incompressible flow through a pump was carried out. The geometry was meshed and validated numerically with the experimental data available in the literature. At the design and off-design conditions, the simulations were conducted to study the effect of roughness and its dependence on Reynolds number of the flow. The performance of the pump was compared for a nondimensional roughness factor (K). A hydraulically smooth surface gives a maximum head. A drop in head observed up to the critical surface roughness (K = 0.1), and the head further rises for K > 0.1. The effect of roughness factor increases with increase in Reynolds number.

Commentary by Dr. Valentin Fuster
2017;():V001T01A018. doi:10.1115/GTINDIA2017-4849.

A methodology for large eddy simulation (LES) of a turbomachine stage is presented. Computations of mean fields (RANS) of stages may be performed separately of rotor and stator rows by providing an averaged solution as input to the down-stream row. In unsteady simulations, unsteady field information must be exchanged in both directions after every time step. Here a procedure for linear cascade simulations of a stage has been implemented in a high-resolution compressible flow solver for LES. The LES uses an explicit filtering method for sub-grid-scale modelling. Grids overlap at the interface between blade rows. Field data is transferred in both directions. Rotor velocity is added or subtracted as needed to tangential velocity component during this transfer. The relative movement of the rotor and stator grids is accounted for by suitable periodic tangential shifting of the paired grid points in the overlap for the transfer. The method has been tested against a published DNS of a statorrotor stage. The Reynolds number based on blade chord and mean axial velocity at inflow was 40000. Solution fields show the wake vortex street of the upstream blade row impinging on downstream blades and being convected through the downstream blade passage. The LES captured transition on rotor blade surface boundary layers. Blade surface pressure distributions agree closely on pressure surfaces. Separation and transition on downstream blade suction surface is delayed slightly at the present resolution, but this will improve with grid refinement, monotonically, for this LES method.

Commentary by Dr. Valentin Fuster
2017;():V001T01A019. doi:10.1115/GTINDIA2017-4871.

Modern Internal combustion engines require high pressure ratios to perform efficiently as well as to reduce emissions. For such applications, a centrifugal compressor with high pressure ratio and broader operating range may be employed. The impeller design of such compressors plays a vital role in producing the efficient operation and hence today’s research focus rigorously over its design.

The objective of the current investigation is to study the performance of centrifugal compressors based on variation in exit width, eye tip radius and shroud extension. Shroud extension have previously found to generate higher pressure rise at the same time have the least amount of losses. The pressure ratio, power and torque requirement, isentropic efficiency and thermal aspects were the main considerations for the study.

Commentary by Dr. Valentin Fuster
2017;():V001T01A020. doi:10.1115/GTINDIA2017-4880.

A large diameter, low speed axial flow fan has been designed to drive the Centre for Railway Research (CRR) wind tunnel at IIT Kharagpur. Total pressure rise of 1400 Pa is required from the fan at the design mass flow rate of 420 kg/s. The outer diameter of the fan is 4 m with 0.6 hub to tip ratio. The present paper discusses the design methodology and computational analysis of wind tunnel fan. The design methodology includes a preliminary mean line analysis using the fundamental governing equations and CFX inbuilt design module. Subsequently, detailed computational analysis of the rotors was carried out using ANSYS CFX®. Based on the results obtained from the CFX analysis, the mean line analysis was modified to achieve the desired performance of the rotor. Parametric study was undertaken with different blade number combination, axial spacing, blade chord length, location of maximum thickness, leading edge profile, etc. The selection of number of blades for rotor and stator is an important aspect for such fan design in terms of cost, utility installation and noise. For low background noise in wind tunnel higher number of blades with suitable combination was selected for the fan ensuring the tone noise (produced by interaction of rotor and stator) is generated at attenuable frequencies. The blade angles, chord lengths, blade camber were optimized for good aerodynamic performance in terms of blade loading, wake thickness, hub-tip pressure profile, overall efficiency and pressure rise. It is strongly believed that discussed design aspects in terms of selection of various geometrical parameters for wind tunnel fan will give useful guidelines for future design of a low speed, large diameter fan of such capacity.

Commentary by Dr. Valentin Fuster


2017;():V001T02A001. doi:10.1115/GTINDIA2017-4502.

To obtain high specific work output with small mass flow rate, high-pressure ratios across the turbines are required in liquid rocket engine turbopumps. An impulse-type supersonic turbine can achieve this. To prevent losses due to low blade aspect ratio and issues related to manufacturing and industrial problems, partial admission configuration is adopted. Partial entry in a turbine is achieved by adjusting the extent of the nozzle arc of admission, leading to a strong unsteady circumferential asymmetry of flow parameters in the rotor passage, and degradation in efficiency. The pressing need of aerodynamic design of supersonic partial admission turbines to improve their efficiency demands an investigation of the viscous fluid dynamic of the turbine flow field. This work reports the aerothermodynamic steady state CFD analysis to obtain the performance parameters of a three-dimensional partial admission turbine for LOX booster turbopump in a semicryogenic engine using ANSYS® CFX. The areas of steady loss have been identified through entropy generation contours, and the effects associated with aerodynamic loss structures like secondary flow, shock location, recirculation with additional pumping and mixing losses have been investigated for designed operating condition corresponding to 100% nominal thrust.

Commentary by Dr. Valentin Fuster
2017;():V001T02A002. doi:10.1115/GTINDIA2017-4542.

In steam turbine power plants, the appropriate design of the last stage blades is critical in determining the plant efficiency and reliability. The development of LP module for desert applications is finding applications for a number of industrial steam turbine operating with air cooled condensers. The conventional LP Module for water cooled condenser operates at low back pressure (Pexit = 0.09 bar) and are generally not suitable for high back pressure application. This paper focuses on the aerodynamic design & optimization of last stages of LP blade module for high back pressure application and validation through 3D CFD. The guide and moving blade are designed with seven equally-spaced profiles section from hub to shroud through Axstream S/w. The profile and incidence losses are minimized for the design and off-design conditions.

Aeromechanical design of LP blade module consisting of 2 stages for 0.2 bar back pressure, 1.1 bar inlet static pressure and a mass flow of 61.2 kg/s is carried out. An optimization process through a streamline curvature code and design optimization software using Optimus is established and flow path contours is optimized thoroughly, a total to total efficiency of 81.4% is achieved for the rated condition. The off-design performance is investigated for a wide range of operating conditions, especially at low volume flow rate of steam condition.

Topics: Pressure , Design , Blades
Commentary by Dr. Valentin Fuster
2017;():V001T02A003. doi:10.1115/GTINDIA2017-4553.

An iterative inverse design methodology is used to design gas turbine blades for the prescribed flow conditions. The input flow parameter considered here is the pressure distribution along the suction and pressure surfaces of the blade. The flow is regarded as inviscid. A guess blade is presumed and the flow analysis over the blade is determined using the existing commercial software. In case of mismatch of the flow parameters, the guessed profile surface is considered as a permeable membrane and the normal velocity on the blade surface is computed by conservation of momentum flux approach. The computed normal velocity is used to revise the blade geometry by mass conservation principle till the flow parameters converge. A few geometric constraints are enforced on the model to avoid quixotic blade model. The validation of the above method is being done using NACA profiles. The robustness of the method is verified by using various combinations of NACA blade profiles, where different initial guessed profiles are taken for the same prescribed pressure distribution.

This approach can be extended to three dimensional cases. To incorporate the complications attached with the three dimensional flows, three two dimensional sections can be considered on the blade geometry namely at hub, mid span and tip.

Commentary by Dr. Valentin Fuster
2017;():V001T02A004. doi:10.1115/GTINDIA2017-4567.

The extraction of wave energy through self-rectifying air turbine is one of the emerging technologies for oscillating water column (OWC) based wave energy devices. In the present effort, a bi-directional impulse (BDI) turbine is designed and the performance parameters were found numerically and compared with an existing unidirectional impulse (UDI) turbine. A brief analytical formulation through similarity laws, to find a dynamically similar BDI turbine using pressure drop vs flow characteristics, gives the approximate diameter range equivalent to the reference UDI turbine. The results are used to reduce the range of diameters and it is found that the characteristics are matching with the reference UDI turbine. The maximum and minimum diameters among the selected range are considered for detailed computational fluid dynamics (CFD) analysis. These two BDI turbines are modeled and meshed in ICEM CFD 14.5. The commercial CFD code CFX 14.5 is used for the numerical simulations. The Reynolds-averaged Navier-Stokes (RANS) equations with the standard k-ϵ scalable wall function model are solved to obtain the performance parameters. A detailed flow physics of the BDI turbines has also been included.

Commentary by Dr. Valentin Fuster
2017;():V001T02A005. doi:10.1115/GTINDIA2017-4583.

The aerodynamic design of a turbine stage requires the accurate prediction of radial profiles of pressure, temperature and velocity at various axial locations within the turbine stage. In the case of hot gas path components like the High Pressure Turbine (HPT), which is located downstream of the combustor, the location of the hot spot and its migration through the stage is critical in arriving at an appropriate aerofoil cooling flow requirement and distribution. In addition, the migration of the flow and the evolution of the temperature traverse through the stage impacts the aerodynamic efficiency of the stage. This is predicted using CFD techniques and has been an inevitable part of the design process.

Typically, the fidelity of the computational model evolves with the component design. During early design phases, simplistic geometry is used for the simulations and progressively the fidelity is increased to resolve the geometrical features of interest, like that of the end wall film cooling and rim seal cavity geometries. The present paper provides an improved understanding of the temperature evolution in a HP turbine stage, particularly with respect to the geometry fidelity and the choice of turbulence models. Computational analyses are carried out using the Rolls-Royce in-house CFD solver, HYDRA.

The geometry fidelity comparisons dealt with are discrete endwall cooling holes vs. equivalent slot and explicit cavity resolution vs. patch surface techniques. In addition, comparisons of traverses predicted using the k-Epsilon realizable turbulence model and SST k-Omega model are presented and debated. The influence of the geometry fidelity and turbulence model on the evolution of radial distribution through the stage is presented along with supporting flow field interpretations. It is concluded that the slot representation of platform cooling flow is satisfactory to replicate the overall traverse at the exit of the High Pressure Nozzle during early stages of design. The near wall temperature gradient would be lower and in the present case the Horse Shoe Vortex (HSV) at the endwalls are not observed with discrete cooling flow modelling which indicates probable aerodynamic impact. The choice of turbulence modelling could have significant impact on the traverse prediction in comparison to the geometry approximations.

Commentary by Dr. Valentin Fuster
2017;():V001T02A006. doi:10.1115/GTINDIA2017-4617.

Central-station power plants (CSPP) are the main provider of energy today. In the process of power generation at central-power stations, about 67% of primary energy is wasted. Distributed cogeneration or combined heat and power (CHP) systems are an alternative to central-station power plants. In these systems, an electrical generation system located in a residence or at a commercial site consumes natural gas to generate electricity locally and then the exhaust heat is utilized for local heating needs (in contrast to being wasted at central-stations). Microturbines offer a number of potential advantages compared to other technologies for small-scale power generation. For example, compact size and low-weight leading to reduced civil engineering costs, a small number of moving parts, lower noise and vibration, multi-fuel capabilities, low maintenance cost as well as opportunities for lower emissions. Inverter generators allow using micro-turbines of different shaft rotation speed that opens opportunities to unit optimization at off-design modes. The common approach to predict the off-design performance of gas turbine unit is the mapping of the compressor and the turbine separately and the consequent matching of common operation points. However, the above-mentioned approach might be rather inaccurate if the unit has some secondary flows. In this article an alternative approach for predicting off-design performance without using component maps is presented. Here the off-design performance is done by direct calculation of the components performances. On each off-design mode, the recalculation of the characteristic of all scheme components, including a compressor, gas turbine, combustor, recuperator and secondary flow system is performed. The different approaches for obtaining the performance at off-design modes considering the peculiarities of the gas turbine engine are presented in this paper.

Commentary by Dr. Valentin Fuster
2017;():V001T02A007. doi:10.1115/GTINDIA2017-4638.

The efficiency of a turbine stage is impacted by a number of factors such as the component design philosophy, operating environment, leakage flow and its interaction with the main gas flow path. When looking at improving a turbine stage performance, there is a natural tendency amidst the designers to look into the factors listed above. Every engine manufacture has a unique style of component design philosophy and hence there are fewer opportunities to radically change the design. On the other hand, the operating environment or operating conditions are usually becoming more challenging. Hence, component designers typically look for opportunities to reduce the leakage or to reduce the losses due to interactive effect of the leakage with the gas path. The rim seal flow and its interaction with the gas path has been of interest for the past few decades and many studies have been carried out to understand the impact of cavity geometry, leakage flows and the ingestion of the hot gas into the rim seal cavities.

The rim seal cavities functionally act as a buffer cavity to dilute and dampen the effect of the hot gas ingested into the secondary air flow path and to prevent the discs from being exposed to ingested hot gas. The successful function of the rim seal cavity depends on multiple factors like rotor-stator axial clearance, cavity volume, cavity shape, cavity approach to the gas path and its interface, in addition to the leakage flow into the main flow path.

The present paper aims at providing a review of a typical rim seal cavity used in the High Pressure Turbine based on systematic CFD studies of the rim seal cavities. While the paper does not present validation data for the approach, the authors attempt to provide references to specific design aspects that are already available in the literature, which are usually less noticed.

Commentary by Dr. Valentin Fuster
2017;():V001T02A008. doi:10.1115/GTINDIA2017-4644.

Improvement in performance and size of gas turbine engine pave the way to the design of high blade loading, low aspect ratio and the small axial gap in the initial stages of turbine. This gives rise to the substantial amount of losses especially secondary flow losses and at the same time, endwall contouring showed its effective part in reduction of secondary flow losses. The intended purpose of the present numerical investigation is to compute the secondary flow mechanisms occurring inside the nozzle guide vane of high pressure turbine stage, a typical representative of modern aero engine design and to optimize its endwall to reduce these losses. Axisymmetric variation in endwall profiles are achieved by functional approximation and genetic algorithm based numerical optimization method. Initially, nozzle guide vane endwalls are parameterized with control points of Bezier curve. A subset of control points are considered as the design variables having a constraint to move in radial coordinate. Statistical tool latin hypercube sampling is adopted to explore the design space. Based on the values of objective functions obtained from the numerical CFD calculations, functional approximation model is construed using the artificial neural network. Finally, a genetic algorithm is used to obtain the optimum solution and analyze the effect of the axisymmetric endwalls on secondary flow losses. Design of endwall profile is studied in three different approaches by confining the axisymmetric variation near hub, shroud and both endwalls. Based on the obtained optimized profiles, secondary flow mechanisms occurring inside the NGV passage are investigated. Reduction in total pressure loss coefficient and improvement in isentropic total-total efficiency are observed for the optimized profiles.

Commentary by Dr. Valentin Fuster
2017;():V001T02A009. doi:10.1115/GTINDIA2017-4669.

A well-formulated design space parametrisation is the key to the success of design optimisation. Most parametrisation methods require manual set-up which typically results in a restricted design space and impedes the generation of superior designs which may be found outside this restricted envelope. In this work, we adopt a NURBS-based automatic and adaptive parametrisation approach where the optimisation begins in a coarser design space and adapts to finer parametrisation during the optimisation. Our approach takes CAD descriptions as input and to alter the shape perturbs the control points of the NURBS patches that form the boundary representation. Driven by adjoint sensitivity information the control net is adaptively enriched using knot insertion. The sensitivity-driven parametrisation method is applied here to reduce the pressure loss of a U-bend passage of a turbine blade serpentine cooling channel.

Commentary by Dr. Valentin Fuster
2017;():V001T02A010. doi:10.1115/GTINDIA2017-4752.

Mean-line modelling approach which has generally been applied to fixed geometry turbocharger turbines has been extended to predict the performance of the variable geometry turbine for different inlet blade angles. The model uses an initial assumption of turbine inlet pressure which was iteratively corrected based on outlet pressure boundary condition. The model was implemented in MATLAB and stable and convergent solutions were obtained using relaxation techniques for different operating conditions. Experiments were done on a state of the art transient diesel engine test bed using the same VGT turbine in the turbocharger at different engine torques and speeds. Using experimental data the model was calibrated for the aerodynamic blockage in the fixed nozzle and rotor blade passages. Results revealed that turbine overall pressure ratio can be predicted accurately if a blockage factor varying with nozzle blade orientation is used in the model.

Commentary by Dr. Valentin Fuster
2017;():V001T02A011. doi:10.1115/GTINDIA2017-4860.

Turbochargers are used in internal combustion engines to increase their volumetric efficiency and power. Turbochargers consist of a centrifugal compressor driven by a radial turbine. Radial turbines convert the excess kinetic energy in the exhaust gases to power. Vane less radial turbine consists of a volute and a turbine wheel. It is preferred because of its low cost, robustness and good off-design performance. In this study, a radial turbine wheel and volute are designed to meet the power and efficiency requirements. A number of trials are carried out, and the design, which gives the necessary performance and meets the customer requirements, is chosen. The design is analyzed using a validated 3D Navier-Stokes (NS) solver, viz. ANSYS-CFX software at both design and off-design conditions and turbine characteristics are generated.

Commentary by Dr. Valentin Fuster
2017;():V001T02A012. doi:10.1115/GTINDIA2017-4868.

The tip leakage flow which is characterized by the blade tip vortex is a highly complex and unsteady phenomenon. It is of great interest to understand the unsteady pressure characteristics of tip leakage flow. The present study aims to focus on the unsteady pressure behavior caused by tip leakage flow of a single stage axial flow turbine. A series of experiments were carried out with a scaled up high work axial flow turbine in a Large Scale Rotating Rig to understand this unsteady pressure behavior. The pressure fluctuations were mapped at five axial locations on the casing across the rotor tip. Unsteady pressures were measured using calibrated fast response transducers. Measurements were phase-locked with the rotor. The pressure data was ensemble averaged and represented gross unsteadiness in the flow. Experiments were carried out over a range of flow coefficients (Cx/U) like 0.30, 0.32, 0.34 and 0.36. The blade pass frequency was the predominant one seen during the measurements. Mild influence of vane exit wake turbulence was also seen during the measurements. Constant speed flow coefficient variation did not seem to cause additional unsteadiness to the tip leakage flow.

Commentary by Dr. Valentin Fuster
2017;():V001T02A013. doi:10.1115/GTINDIA2017-4874.

The curvature of a turbine blade airfoil downstream of the throat location significantly affects its aerodynamic performance, specifically at Mach number close to unity. In the present work, a low aspect ratio (0.64), highly curved back airfoil corresponding to stator blade ‘mean’ section of a high-pressure (HP) turbine stage is studied. The details of the blade parameters, experimental test setup, CFD solver and numerical setup are explained in the paper. Its aerodynamic characteristics are obtained numerically using a commercial CFD solver and are compared to those from experimental cascade test results. For numerical assessment, CFD simulations are carried out on three configurations viz. (i) Full turbine stage (stator and rotor) domain (ii) Isolated turbine stator row domain (iii) Stator mean section airfoil cascade domain. The loss predictions obtained through CFD are also compared against the loss estimates calculated using two loss models. The experimental cascade pressure loss across the blade row at design point Mach number 0.996 increases to 250% of that at lower Mach numbers. This drastic increase is not desirable. But the airfoil performs appreciably well in a ‘stage’ setup i.e. with downstream rotor. Therefore, the present study brings out the behaviour of the stator airfoil performance in a linear cascade, annular cascade and stage environments.

Topics: Airfoils
Commentary by Dr. Valentin Fuster
2017;():V001T02A014. doi:10.1115/GTINDIA2017-4912.

Ocean stores a huge amount of energy and ocean current energy can be a viable source in future. In this article, an axial marine current turbine has been optimized to enhance its power coefficient through numerical modeling. The blade pitch-angle and number of blades are the design parameters chosen for the analysis to find the optimal design. A commercial code for CFD simulations with in-house optimization code was used for the analysis. It was found that, changing the blade pitch-angle and reducing the number of blades can improve the turbine’s coefficient of power. This is due to increase in lift and reduction of losses caused by turbulence near the downstream of the turbine. The article presents flow-simulation difficulties and characteristic curves to identify the differences between the actual and optimized turbine. The detailed flow physics is discussed and pictured in the post processed plots.

Topics: Design , Turbines
Commentary by Dr. Valentin Fuster

Heat Transfer

2017;():V001T03A001. doi:10.1115/GTINDIA2017-4525.

Turbine inlet air cooling (TIAC) has long been the most commonly used method to improve the performance of gas turbine based power plants. It is particularly effective in regions with high ambient temperatures. With growing energy demands and higher ambient temperatures around the globe, it is important to look beyond cooling cycles like vapor-absorption and vapor-compression which have certain limitations. It is prudent to use a vapor-adsorption cycle for TIAC since the waste exhaust heat can be utilized as the power source for adsorption compressor, resulting increase in thermal efficiency of the power plant. Also, the scalability of adsorption cooling from mere Watts to hundreds of kW and its ability to function using lower temperature heat sources (as low as 60 °C) render it highly suitable for TIAC. In this paper, a gas turbine power plant and a TIAC system running on vapor-adsorption cycle are mathematically modelled and thermal analysis involving comparison of performance of the power plant with and without inlet air cooling at various ambient and desorption temperatures is presented. Performance parameters analyzed include net power output and thermal efficiency of the power plant and the COP of the chiller. The results show that vapor-adsorption system has huge potential to be integrated with gas turbine power plant for inlet air cooling.

Commentary by Dr. Valentin Fuster
2017;():V001T03A002. doi:10.1115/GTINDIA2017-4540.

Vapor compression refrigeration plant power consumption can be heavily reduced for a definite cooling power by the adoption of a power regeneration process internal to the simple cycle. Such a process takes benefit from the adoption of turbomachinery based vapor pressure amplifier.

The selection of the wheel sizes, their machining to be adapted to the process and the addition of stator blades in the expander and a diffuser in the compressor path is discussed and tests results are presented.

Improvements of the power consumption and of the COP of about 23–24% for a specified cooling power have been demonstrated.

Commentary by Dr. Valentin Fuster
2017;():V001T03A003. doi:10.1115/GTINDIA2017-4549.

The inlet air temperature of turbines are significantly high which may result in the damage of the blade material. As a consequence, it is required to cool the turbine blades and a number of cooling techniques are introduced in the past. The cooling technique involving evaporation of water droplets is the motivation and focus of this work. When the water droplets are injected along with the dry air, the existing concentration difference between the water droplets and dry air results in the evaporation leading in the drop of the coolant temperature. This phenomena does not occur when the coolant is dry air and this part is addressed for the first time in this study. The cooling performance is investigated on wall of straight channel in the presence of a film cooling hole. Dry air along with water droplets are supplied through the film cooling hole and the computations are carried out for different droplet diameters and mist concentrations. Results showed that, the overall cooling effectiveness is always larger for the mist-air case compared to that for the dry air case. Also, the cooling effectiveness increases with the percentage of mist in the mist-air system for a specific droplet diameter. At a specific mass flow rate and specific mist percentage, the increase in droplet diameter results in the decrease of cooling effectiveness.

Commentary by Dr. Valentin Fuster
2017;():V001T03A004. doi:10.1115/GTINDIA2017-4557.

Gas/steam combined cycle power plants are extensively used for power generation across the world. Today’s power plant operators are persistently requesting enhancement in performance. As a result, the rigour of thermodynamic design and optimization has grown tremendously. To enhance the gas turbine thermal efficiency and specific power output, the research and development work has centered on improving firing temperature, cycle pressure ratio, adopting improved component design, cooling and combustion technologies, and advanced materials and employing integrated system (e.g. combined cycles, intercooling, recuperation, reheat, chemical recuperation). In this paper a study is conducted for combining three systems namely inlet fogging, steam injection in combustor, and film cooling of gas turbine blade for performance enhancement of gas/steam combined cycle power plant. The evaluation of the integrated effect of inlet fogging, steam injection and film cooling on the gas turbine cycle performance is undertaken here. Study involves thermodynamic modeling of gas/steam combined cycle system based on the first law of thermodynamics. The results obtained based on modeling have been presented and analyzed through graphical depiction of variations in efficiency, specific work output, cycle pressure ratio, inlet air temperature & density variation, turbine inlet temperature, specific fuel consumption etc.

Commentary by Dr. Valentin Fuster
2017;():V001T03A005. doi:10.1115/GTINDIA2017-4568.

Film cooling is one of the preferred methods for effective cooling of a gas turbine that forms a protective layer between hot flue gases and blade surface. This paper investigates the interaction of mist in the secondary flow and physics indicating the upper limit of mist concentration. Numerical simulations are performed on a flat plate having a series of discrete holes with 35 degree streamwise orientation and the holes are connected to a common delivery plenum chamber. The blowing ratio, density ratio and Reynolds number based on freestream and hole diameter (D) are 0.5, 1.2 and 15885 respectively. A two-phase mist consisting of finely dispersed water droplets of 10 micron in an airstream is introduced as the coolant from these holes. The latent heat absorbed by the evaporating droplets significantly reduces the sensible heat of the main stream, providing heat sinks that result in enhanced cooling effectiveness. The coupling between the two-phases is modelled through the interaction terms in the transport equations. Computations are performed by ANSYS Fluent 15.0 using k-ε realizable model.

The results illustrate insight of complex transport phenomena associated with the mist of varying concentration from 2% to 7%. It has been observed that the maximum enhancement of cooling effectiveness reaches 43% at X/D = 10 for 2% mist by mass with an average enhancement of 26.5%. For 3% mist, the maximum enhancement becomes 80% at X/D = 16 with the average cooling enhancement of 43%. Mist concentrations 5% and beyond trend to increase average cooling because of more absorption of latent heat by droplets, but its trajectories shift towards wall, detrimental to the blade due to corrosion effect and thermal stresses.

Commentary by Dr. Valentin Fuster
2017;():V001T03A006. doi:10.1115/GTINDIA2017-4580.

The work attempts to address excessive heating and related risks in most of the engineering systems. Perforations are well known to boost heat transfer. Present work is an attempt to pact optimization of perforated enclosures for internal natural convection heat transfer. Heat dissipation effect is experimented over a flat plate and implications are understood with variation in convective heat transfer coefficient. Controlling parameters viz., plate orientation, perforation shape and size, enclosures in diverse configurations are varied systematically aiming enhanced heat transfer. Results confirm fact that perforated enclosures significantly affect the heat transportation. Enclosures with varying perforations are found to yield distinct heat sink characteristics. For varying perforation shape, size and plate orientation, the heat transfer rate variation is owed by the resultant flow behavior which governs the energy transference. The complied results are noted as excellent physical insight and optimized to propose a novel design for operations under varying heat transfer requirements for wide-ranging applications.

Commentary by Dr. Valentin Fuster
2017;():V001T03A007. doi:10.1115/GTINDIA2017-4586.

Gas turbine operating temperatures are projected to continue to increase and this leads to drawing more cooling air to keep the metals below their operational temperatures. This cooling air is chargeable as it has gone through several stages of compressor work. In this paper a surrogate based design optimization approach is used to reduce cooling mass flow on combustor tiles to attain pre-defined maximum metal surface temperatures dictated by different service life requirements.

A series of Kriging based surrogate models are constructed using an efficient GPU based particle swarm algorithm. Various mechanical and manufacturing constraints such as hole ligament size, encroachment of holes onto other features like side rails, pedestals, dilution ports and retention pins etc. are built into the models and these models are trained using a number of high fidelity simulations. Furthermore these simulations employ the proprietary Rolls-Royce Finite Element Analysis (FEA) package SCO3 to run thermal analysis predicting surface heat transfer coefficients, fluid temperatures and finally metal surface temperatures.

These temperature predictions are compared against the pre-defined surface temperature limits for a given service life and fed back to the surrogate model to run for new hole configuration. This way the loop continues until an optimized hole configuration is attained. Results demonstrate the potential of this optimization technique to improve the life of combustor tile by reducing tile temperature and also to reduce the amount of cooling air required.

Commentary by Dr. Valentin Fuster
2017;():V001T03A008. doi:10.1115/GTINDIA2017-4608.

Multi-objective optimization of the film cooled holes for the coupled impingement-film cooled nozzle guide vane is conducted. Two objectives are considered to be minimized: coolant jet exit total temperature and static pressure drop, to assess the trade-off between them. Three-dimensional computationally using SIMPLE algorithm analysis and a k-ω SST turbulence model are used for generating a data base. The plenum mass flow rates and mainstream velocity are considered as the two design variables. The second order polynomial response surface method is chosen to develop the objective function approximation. The multi-objective optimization has been carried out with help of a genetic algorithm and sequential quadratic programing (fminicon) in MATLAB 7.11.0 (R2010b). The Pareto-optimal design points are obtained as the plenum coolant mass flow rate of 0.004kg/s and mainstream velocity of 10m/s. Based on these results, the global minimum coolant jet static pressure drop of 32.5 Pa and global minimum jet exit total temperature of 312K are observed for the film holes of the NGV surface. At these operating conditions, the coupled impingement -film cooled NGV is subjected to its higher safety and durability. This happened due to without cause of hot gas ingestion into the film cooled jets of the typically cooled NGV.

Commentary by Dr. Valentin Fuster
2017;():V001T03A009. doi:10.1115/GTINDIA2017-4632.

The objective of this present work is to investigate numerically the effect of converging conical hole on blade cooling at leading edge. Diameter ratio of the converging conical holes is maintained as two with a converging angle of 20°. Cylindrical geometry used by Lee et al. [5] for the experimental investigation is taken as base reference for this present numerical investigation. Turbulence model study is carried out with three different models and kω-SST is found to give closer result with the experimental data. Subsequent investigations are carried out using kω-SST turbulence model. The target surface for the present study is 19.05 cm in radius and the diameter of the impingement hole is 1.30 cm, 2.15 cm 3.40 cm. The jet hole to the target surface spacing is varied as R/4, R/2 and 3R/4. The steady-state Reynolds Averaged Navier Stokes equations are solved for different impingement hole diameters at Reynolds number of 11000, 23000 and 50000. The target surface is maintained at constant heat flux of 10000 W/m2. Numerically computed Nusselt number and temperature distribution for the convergent conical hole and cylindrical hole are compared. Around 186% increase in Nu and 13% decrease in surface temperature is observed at the stagnation point for the optimum case in this present study i.e, jet spacing R/2, converging conical hole diameter 2.15 cm and Re 23000. The converging conical hole configuration increases the fluid velocity in the potential core region and enhances the heat transfer. It is followed by 185% increase in Nu and 15.58% reduction in target surface temperature for Re 11000 and 170% increase in Nu and 7.28% reduction in temperature for Re 50000.

Commentary by Dr. Valentin Fuster
2017;():V001T03A010. doi:10.1115/GTINDIA2017-4651.

Rib turbulator is the most effective, economically feasible, and rigorously studied tool to increase thermal performance because of its fundamental nature and due to the vast field of industrial applications. The rib turbulator results in heat transfer enhancement with additional pressure penalties, and thus encourages the researcher and designers towards selecting an efficacious rib configuration. The present work is a study towards detailed heat transfer and flow field characteristics inside a rectangular duct roughened by solid as well as ventilated pentagonal ribs placed transversely on the bottom wall. The rib height-to-hydraulic diameter ratio, the rib pitch-to-height ratio, the open area ratio, and the Reynolds number based on duct hydraulic diameter fixed during experiments are 0.125, 12, 25%, and 42500, respectively. The heat transfer coefficient (HTC) distribution was mapped by using transient Liquid Crystal Thermography (LCT) technique, while detailed flow measurements were made by using Particle Image Velocimetry (PIV) technique. The investigation focuses towards assessing the influence of three different rib configurations named as solid pentagonal ribs, pentagonal rib with parallel slit, and pentagonal rib with inclined slit, on the local heat transfer fields as well as flow characteristics. The flow mechanisms responsible for high or low heat transfer regions as well as for hot-spot formation in the wake of the ribs are identified and explained. The overall heat transfer and friction factor measurements are observed along with the thermohydraulic performance. Results show that the solid pentagonal ribs are superior to slitted ribs from both heat transfer augmentation and thermo-hydraulic performance perspective. Additionally, the slitted pentagonal ribs significantly control the small-scale vortices present at the leeward corner of the solid pentagonal ribs and eventually facilitates in preventing the hot spots formation with reduced pressure penalty.

Commentary by Dr. Valentin Fuster
2017;():V001T03A011. doi:10.1115/GTINDIA2017-4743.

Shock Vector Controlling (SVC) nozzle is one of the most effective ways to obtain the thrust vector for the advanced engine. The SVC nozzle has the advantages of simple structure, light weight and fast response. Numerous analytical, numerical and experimental studies have been carried out on SVC nozzle in recent years. While these studies mainly focus on the nozzle and ignore the influence between SVC nozzle and engine. In fact, the secondary flows of fluidic thrust vectoring nozzle are all extracted from the engine fan or high pressure compressor, no matter it is based on SVC or others. The extraction and injection of the secondary flow both influence the components matching and engine performance. And the mass flow and pressure of secondary flow are constrained by engine working condition in turn. So the SVC nozzle should be installed in the engine environment for evaluation.

The influence of SVC nozzle secondary flow on engine thermodynamic cycle and performance is in two major ways. First, when the secondary flow extracts from the fan or HPC, the operating point of the engine will change. It results in the thermodynamic parameters at each engine stations and the performance of engine change. The reason for this change is that the bleeding of secondary flow broke the mass flow balance equations between the engine components. Second, the injection of secondary flow in the expansion section of nozzle also leads to engine performance change, while the engine operating point remains normal.

The characteristic of certain SVC nozzle is obtained by computational fluid dynamics method. Then the Response Surface Method (RSM) model based on the design of experiment is employed to build the relations of performance parameters and design variables of SVC nozzle. Combining basic engine performance simulation model with SVC nozzle RSM model, the model for engine with SVC nozzle is built. To evaluate the performance of the engine with SVC nozzle, the conceptual design of double bypass variable cycle engine with SVC nozzle is given. The influence of the secondary flow extracting from the fan outlet, CDFS outlet and the first stage of HPC is discussed. And its performances such as vector angle, thrust coefficient, thrust and specific fuel consumption are obtained by employing the engine simulation model built by this paper.

Commentary by Dr. Valentin Fuster
2017;():V001T03A012. doi:10.1115/GTINDIA2017-4746.

The power output of a gas turbine plant decreases with the increase in ambient temperature. Moreover, the ambient temperature fluctuates about 15–20°C in a day. Hence, cooling of intake air makes a noticeable improvement to the gas turbine performance. In this regard, various active cooling techniques such as vapor compression refrigeration, vapor absorption refrigeration, vapor adsorption refrigeration and evaporative cooling are applied for the cooling of intake air. This paper presents a new passive cooling technique where the intake air temperature is reduced by incorporating phase change material (PCM) based heat exchanger parallel to conventional air intake line. During the daytime, the air is passed through the PCM which has melting temperature lower than the peak ambient temperature. This will reduce the ambient air temperature before taking to the compressor. Once the PCM melts completely, the required ambient air would be drawn from the ambient through conventional air intake arrangement. During the night, when there is lower ambient temperature, PCM converts from liquid to solid. The selected PCM has a melting temperature less than the peak ambient temperature and higher than the minimum ambient temperature. It is observed from the numerical modeling of the PCM that about four hours are required for the melting of PCM and within this time, the intake air can also be cooled by 5°C. The thermodynamic analysis of the results showed about 5.2% and 5.2% improvement in net power output and thermal efficiency, respectively for four hours at an ambient temperature of 45°C.

Commentary by Dr. Valentin Fuster
2017;():V001T03A013. doi:10.1115/GTINDIA2017-4771.

In this paper, a double bypass variable cycle engine with FLADE (Fan on Blade) is considered. The FLADE VCE is one of the research hotspots for future military and civil aircraft power device, which shows outstanding performance advantages. Compared to the mixed-flow turbofan, FLADE VCE is more complex than conventional aero-engine for its multi-components and multi-variable parts, which make it difficult to modeling and optimization.

For getting the performance of FLADE VCE, the model for engine performance simulation is researched. The method for FLADE performance simulation and the steady-state performance simulation model for FLADE VCE are developed. And a component-based engine performance simulation system is established based on object-oriented modeling method.

For obtaining the optimal integrated performance of FLADE VCE, suitable optimization method is required. Unfortunately, the optimization of FLADE VCE is a non-linear non-differentiable problem, which makes it difficult to solve by conventional deterministic optimization method. In order to solve this problem, the differential evolution (DE) algorithm is considered. To overcome the limitations of original DE algorithm, an improved DE algorithm with modifying mutation operator is proposed by this paper. The FLADE VCE optimization problem is solved by employing the improved DE algorithm.

Commentary by Dr. Valentin Fuster
2017;():V001T03A014. doi:10.1115/GTINDIA2017-4776.

The Gas turbine combustion chamber is the highest thermally loaded component where the temperature of the combustion gases is higher than the melting point of the liner that confines the gases. Combustor liner temperatures have to be evaluated at all the operating conditions in the operating envelope to ensure a satisfactory liner life and structural integrity. On experimental side the combustion chamber rig testing involves a lot of time and is very expensive, while the numerical computations and simulations has to be validated with the experimental results. This paper is mainly based on the work carried out in validating the liner temperatures of a straight flow annular combustion chamber for an aero engine application. Limited experiments have been carried out by measuring the liner wall temperatures using k-type thermocouples along the liner axial length. The experiments on the combustion chamber testing are carried out at the engine level testing. The liner temperature which is numerically computed by CHT investigations using CFX code is verified with the experimental data. This helped in better understanding the flow characterization around and along the liner wall. The main flow variables used are the mass flow rate, temperature and the pressure at the combustor inlet. Initially, the fuel air ratio is used accordingly to maintain the same T4/T3 ratio. The effect of liner temperature with T3 is studied. Since T4 is constant, the liner temperature is only dependent on T3 and follows a specific temperature distribution for the given combustor geometry. Hence this approach will be very useful in estimating the liner temperatures at any given T3 for a given combustor geometry. Further the liner temperature is also estimated at other fuel air ratios (different T4/T3 ratios) by using the verified CHT numerical computations and found that TL/T3 remains almost constant for any air fuel ratio that is encountered in the operating envelope of the aero engine.

Commentary by Dr. Valentin Fuster
2017;():V001T03A015. doi:10.1115/GTINDIA2017-4811.

This paper describes five methods to achieve effective heat transfer and higher plant efficiency when condensate return temperature is high and treatment is needed to improve water quality in the water treatment plant before sending it to the deaerator for a combined heat and power plant.

Commentary by Dr. Valentin Fuster
2017;():V001T03A016. doi:10.1115/GTINDIA2017-4881.

In the present work, transient liquid crystal thermography (LCT) has been used for capturing the temperature field as well as the local heat transfer distribution inside a rectangular duct. Experiments have been carried out in an open loop airflow system at a Reynolds number (based on the channel hydraulic diameter) of 58850 and for rib height to channel hydraulic diameter ratio of 0.125. This investigation emphases headed for assessing the potential impact of design parameters such as chamfering angle and rib pitch to height ratio of the trapezium ribbed rectangular duct on the thermo-hydraulic performances, which forms the basis of analysis while using response surface methodology (RSM). The chamfering angle has been varied from 0 to 20° in a step of 5°, while the rib pitch to height ratio is varied from 8 to 12 in a step of 2. The quadratic model generated by RSM is used to predict the optimal performance parameters. The results show that different geometrical parameters have to be considered simultaneously in order to improve the performance of ribbed-duct. Eventually, based on this analysis, the optimum levels of design parameters for trapezium rib corresponding to the highest augmentation Nusselt number, the lowest friction factor, and the highest thermo-hydraulic performance have been determined. Finally, the desired correlations for all performance parameters have been developed using RSM. The comparison of predicted values with the experimental values has been carried out, which is found to be in harmony with the experimental values in the uncertainty range of ±5%., which are found to predict the performance parameters with reasonably good accuracy.

Commentary by Dr. Valentin Fuster

Combustion, Fuels and Emissions

2017;():V001T04A001. doi:10.1115/GTINDIA2017-4512.

It is a well-known fact that NO2 has far more harmful effects as compared to NO. NO2 creates ozone, which causes eye irritation and exacerbates respiratory conditions. This leads to an increased emergency departments’ visits and hospital admissions for respiratory issues, especially asthma. Under current situation, majority of regulations deal with total NOx emissions, without looking at the break-up of NO2 and NO. However, there is a feeling in emissions regulation community to implement regulations on NO2 emissions. There are standards to measure total NOx emissions. However, these standards are not equipped enough to measure NO2 emissions accurately. The effect of sample line length on NO2 emissions is not fully understood to date. Also, the standards only suggests maximum of 10 seconds residence time regardless of what the line length is. In this study, a systematic experimental test campaign has been conducted to understand the effect of sample line length on NO2, NO distribution. The residence time was maintained below 10 seconds in accordance with the SAE ARP1256D standards. A Rolls-Royce gas turbine combustor and different calibration cylinders have been used to study the effect of sample line length. A numerical study has also been done to predict the conversion of NO2 to NO. It has been found that with increasing sample line length, more NO2 gets converted to NO and overall NO2 emissions show a reduction, whereas this would not be the case at engine exhaust. This effect of sample line length can be used as a loophole in giving lower NO2 emissions readings.

Commentary by Dr. Valentin Fuster
2017;():V001T04A002. doi:10.1115/GTINDIA2017-4600.

The current work is aimed towards development of high thermal intensity, low emission combustor for gas turbine engines. Employing discrete and direct injection of air and fuel in a combustion chamber and has been demonstrated to result in low pollutant emissions (NOx, CO, UHC). From our previous investigations, we found that the reverse-cross flow configuration, where air is injected from the exit end and fuel is injected in the cross flow of the injected air, results in favorable combustion and emission characteristics. Though the air jet is the dominant jet, the fuel jet can also influence the flow field, mixing and the combustion behavior inside the combustor, which is the subject of the current investigation. Here we investigate a high thermal intensity combustor relevant to gas turbine engines (at equivalence ratio of 0.8, the combustor operates at thermal intensity of 39 MW/m3-atm and heat load of 6.25 kW). Natural gas is used as the fuel and two different fuel injection diameters of 1 mm and 2 mm are investigated. This result in significantly higher (four times) fuel jet momentum from the smaller fuel injection port as compared to the larger port. From computational fluid dynamics (CFD) studies, it is observed that for the case with higher fuel jet momentum, the fuel jet deflects the air jet such that the flow pattern is significantly altered as compared to the case with lower fuel jet momentum. OH* chemiluminescece images show that the reaction zone location is significantly affected with high momentum fuel jet. NOx is reduced whereas CO is increased with higher momentum fuel jet.

Commentary by Dr. Valentin Fuster
2017;():V001T04A003. doi:10.1115/GTINDIA2017-4601.

This study deals with an experimental investigation of a low emission liquid fuelled (ethanol) reverse cross-flow combustor. This investigation is carried out to cater to the need of burning liquid fuels (including alternative fuels) with minimum emissions in gas turbine engines used for both aircraft and land based power generation applications using modern combustion technologies. In the present combustor design, the air inlet and the exhaust ports are located on the same side (and hence the name reverse-flow) whereas the liquid fuel is injected directly into the strong cross-flow of the air using a small diameter round tube to aid fuel atomization. Hence, a conventional atomization system is absent in the investigated combustor. The reverse-flow configuration allows effective internal product gas recirculation to facilitate the preheating and dilution of the oxidizer stream and stabilization of a distributed reaction zone. This apparently suppresses near stoichiometric reactions and hot spot regions resulting in low pollutant (NOx and CO) emissions. In the present case, the heat load is varied (keeping a constant air flow rate) from 3.125 kW to 6.25 kW which results in the thermal intensity variation from 19 MW/m3-atm to 39 MW/m3-atm. Two different tubes with internal diameters (dfuel) of 0.5 mm and 0.8 mm are used for injection of liquid fuel into the cross flow of air. The combustor was also tested in premixed-prevaporized (PP) mode with ethanol for benchmarking. The combustion process was found to be stable with NOx emissions of 1.6 ppm (premixed-prevaporized), 8 ppm (dfuel = 0.5 mm), 9 ppm (dfuel = 0.8 mm). The CO emissions were 5 ppm (premixed-prevaporized), ∼100 ppm (dfuel = 0.5, 0.8 mm), at atmospheric pressure operation (corrected to 15% O2) and ϕ = 0.7, Tadiabatic ∼1830 K. Reaction zone positioning inside the combustor was investigated using OH* chemiluminescence imaging and global flame pictures, and the same was found to be located in the vicinity of the air jet.

Commentary by Dr. Valentin Fuster
2017;():V001T04A004. doi:10.1115/GTINDIA2017-4631.

A gas turbine afterburner is required to operate under severe conditions of pressure and temperature to meet the design requirements of next generation gas turbine engines. This fact, coupled with the current trends towards higher turbine discharge temperature and the requirement for satisfactory operation over extended fuel/air ratios and flight maps call for greater understanding of the internal aerodynamics for improving thrust developed by the afterburner. The present work focuses on prediction of performance of a practical afterburner for different altitude conditions and reheat strengths (i.e., fuel-air ratios) using Computational Fluid Dynamics (CFD) simulations. Combustion efficiency and thrust boost at these conditions have been predicted. The reacting flow field has been analyzed and changes suggested for improving thrust at low performance points.

Commentary by Dr. Valentin Fuster
2017;():V001T04A005. doi:10.1115/GTINDIA2017-4635.

The present work describes the state-of-the-art technology for a Sideway Faced Porous Radiant Burner (SFPRB) of 10–15 kW capacity, operated by liquefied petroleum gas (LPG) applicable for industrial furnace and incinerator. The newly developed SFPRB is a two layer burner, consisting of a reaction zone and a preheat zone. The combustion zone is of reticulated SiC ceramic matrix of porosity 90%, diameter 120 mm and thickness 20 mm and the preheat zone is of Al2O3 ceramic having 463 through holes (diameter 1.5 mm), with 15 mm thickness and 120 mm diameter. The work presents the effect of geometrical parameters (length of mixing pipe and diameter of orifice) on the radial temperature distribution of burner surface. Experimentation has been done in 15 kW input power to study the behavior of air-fuel mixture entering the burner. Ultimately, it is focused for uniform temperature distribution on the burner surface with a suitable arrangement. The work also presents a detailed account of the temperature distribution along the two main burner axes and the emission measurements (CO and NOx) for the suitable SFPRB. Investigation was done for an input power range of 10–15 kW with an equivalence ratio of 0.5.

Commentary by Dr. Valentin Fuster
2017;():V001T04A006. doi:10.1115/GTINDIA2017-4642.

The latest trend in aero engine industry is to reduce the emission levels to a maximum extent as well as to improve the combustion efficiency. One of the ways to reduce the emission levels and improve the combustion efficiency in gas turbine combustors is to operate with lean premixed conditions. However, design modifications are required in conventional combustors to operate at lean premixed conditions.

A new combustor configuration was studied for lean premixed pre-vaporized (LPP) combustion. To arrive at LPP configuration a simple gas turbine Can-combustor geometry is modified by attaching a fuel air premixer chamber at upstream of the flame tube of basic diffusion combustor [1]. Reactive flow analysis was performed with both the LPP and conventional diffusion combustor using commercial CFD tool Fluent. Significant improvements were obtained from LPP combustor in terms of emission reduction. However, there are limitations for lean burning of fuel such as flashback and lean blow out. Present work is focused on studying these issues on the LPP combustor. These methodologies can be extended for other type of combustors after further studies.

Commentary by Dr. Valentin Fuster
2017;():V001T04A007. doi:10.1115/GTINDIA2017-4655.

The main objective of this computational analysis is to investigate the effect of increase in Weber number at constant momentum flux ratio on the primary breakup process and deformation of kerosene jet in cross stream air flow. Unsteady computational analysis with VOF approach is carried out to simulate the two phase flow at three different cross flow Weber number conditions (150, 350 and 400) at constant momentum flux ratio of 17. Since the results of VOF technique is highly sensitive to the size and distribution of grid, grid optimization process is carried out, with both structured and unstructured forms of the grid. Since the structured grid with number of elements 17,96,181 displayed better matching with experimental results of upper trajectory of kerosene jet; this grid is used to investigate the effect of turbulence model and Weber number on the windward trajectory of kerosene jet in cross flow air stream. Initially to evaluate the results of computational analysis; simulations are carried out with larger computational domain (with number of elements 17,96,181). Windward trajectory of computational analysis is compared with experimental results of upper trajectory predicted using image processing technique and reasonable overall matching is observed.

To investigate the primary breakup process and deformation of liquid jet at three different increasing Weber number conditions, simulations are carried out with smaller computational domain with higher mesh density with number of elements 33,96,146. The computational technique used in the present analysis exactly captures the modes of breakup observed from experimental results at different Weber number operating conditions. To characterize the deformation of liquid jet at different Weber number conditions; near-field trajectory, cross stream dimension and wave length of liquid jet are quantified at different instants of time. With increase in Weber number, decrease in penetration of liquid jet along transverse direction and more bending of liquid jet along flow direction is observed. From the velocity profile along transverse direction of three different conditions, stronger shearing of liquid film is observed in higher Weber number conditions.

Commentary by Dr. Valentin Fuster
2017;():V001T04A008. doi:10.1115/GTINDIA2017-4710.

In this paper, we present a novel initial attempt on analysis of the mitigation mechanism of instability by rotating the otherwise static swirler in a lean premixed, swirl stabilized, labscale combustor. It has been reported in our previous work that increasing the swirler rotation rate mitigates the self-excited thermoacoustic instability in a model lab-scale combustor, over a range of conditions. Here, it is found that for a given period of observation, instead of a continuous and gradual decrease in the time localized pressure amplitude from the fully unstable state towards the fully mitigated state, the fraction of the time during which instability is present is reduced. With increasing swirler rotation rates, the instability becomes more intermittent with progressive reduction in frequency of their occurrence. High speed PIV results are also presented along with simultaneous pressure signals which support this claim. Such an intermittent route to instability mitigation could be attributed to the background turbulent flow field and is reminiscent of the intermittent opposite transition (implemented by changing the Reynolds number) from a fully chaotic state to a fully unstable state as recently discovered in Nair, Thampi and Sujith [1]. An attempt is made to model the behavior of pressure oscillations using the well established mean-field Kuramoto model. The variation of the order parameter r, which is the parameter for the measurement of synchronization between the oscillators provides critical insights on the transition from the unstable, intermittent to stable states.

Commentary by Dr. Valentin Fuster
2017;():V001T04A009. doi:10.1115/GTINDIA2017-4725.

Swirl cups (hybrid atomizers) are being widely employed in aero gas turbine engine combustors for their established merits in terms of achieving satisfactory atomization over the entire combustor operating regime. Even though several investigators have worked on development of these swirl cups, there is a scanty data reported in literature relevant to their design.

In the present study, flow behavior in a swirl cup assembled in a confined chamber similar to a gas turbine combustor has been analyzed. Flow analysis has been carried out using ANSYS Fluent and turbulence has been modeled using Realizable k-ϵ model. Six swirl cup configurations have been analyzed; mass flow ratio between primary and secondary swirler and venturi converging area ratio have been varied. The effect of these parameters on downstream flow field has been studied by analyzing the profiles of axial, tangential and radial velocities downstream of swirl cup. The size and shape of the recirculation zone has been analyzed and reported for all configurations. Also, the mass flow recirculated by swirl cup has been estimated and compared amongst the configurations analyzed. Data thus generated is very useful in designing such swirl cups of gas turbine combustors.

Commentary by Dr. Valentin Fuster
2017;():V001T04A010. doi:10.1115/GTINDIA2017-4728.

Relight envelope of the combustor needs to be experimentally generated and established during the design and development of an aero gas turbine engine. Usually, during development stage of engine, compressor characteristics are not readily available at such low speeds and hence, it becomes difficult to specify the combustor inlet conditions such as pressure, temperature and Mach number during the engine light up studies. This paper compares the experimental test data generated on an annular combustor for windmill conditions during stand-alone mode and engine level tests under simulated flight conditions. The stand-alone combustor trials were conducted for the range of total pressure and temperature relevant to the flight altitude and Mach number range. During the engine level tests, combustor relight tests were conducted under simulated conditions (ISA+15) for altitudes ranging from 5.5 km to 10 km, flight Mach numbers in the range of 0.45 to 0.80. In this paper, effect of altitude and flight Mach number on the windmill spool speed, combustor pressure and temperature are studied.

Commentary by Dr. Valentin Fuster
2017;():V001T04A011. doi:10.1115/GTINDIA2017-4732.

Gas turbine engine combustors widely employ injection of liquid fuel in cross flow air as a fuel atomization process. Fuel atomization has a major impact on the efficiency, exit profile and emissions of the combustors. Since the quality of spray directly affects the performance of the combustors, it is important to understand a complex phenomenon of spray formation involving trajectory evolution, surface breakup, column fracture and dispersion of secondary droplets. There are several ways of analyzing spray formation which include experimental measurements, numerical analysis using Direct Numerical Simulation (DNS), a combination of output from Volume of Fluid (VOF) and empirical correlations or a hybrid Eulerian-Lagrangian multiscale method. In any of the numerical methods used for the analysis of the sprays, the mesh topology and resolution have a great impact on the accuracy of the predictions.

In the present work, evaluation of mesh topology particularly hexahedral and polyhedral meshes, is carried out to understand their effect on the accuracy of the spray prediction and explore the possibility of employing the polyhedral mesh for such a complex phenomenon. The mesh generation for polyhedral meshes is comparatively easier than hexahedral meshes and it has better control on the growth of the element size, resulting in lesser mesh count. A hybrid Eulerian-Lagrangian multiscale method developed earlier is used along with scale resolving turbulence model (Large Eddy Simulation) in this work. The test case considered for the analysis is taken from the experimental analysis done by Gopala et al. [2]. Momentum flux ratio of 10 and weber number of 1500 are selected for the present analysis. Simulation results from both the mesh topologies are compared with the experimental results for quantities like jet penetration length, Sauter mean diameter and droplet velocity profiles at different z/d locations. It is observed that both the mesh topologies considered, perform equally well and therefore, polyhedral mesh topology can be employed successfully for hybrid spray modeling.

Commentary by Dr. Valentin Fuster
2017;():V001T04A012. doi:10.1115/GTINDIA2017-4736.

Emission standard agencies are coming up with more stringent regulations on soot, given its adverse effect on human health. It is expected that Environmental Protection Agency (EPA) will soon place stricter regulations on allowed levels of the size of soot particles from aircraft jet engines. Since, aircraft engines operate at varying operating pressure, temperature and air-fuel ratios, soot fraction changes from condition to condition. Computation Fluid Dynamics (CFD) simulations are playing a key role in understanding the complex mechanism of soot formation and the factors affecting it.

In the present work, soot formation prediction from numerical analyses for turbulent kerosene-air diffusion jet flames at five different operating pressures in the range of 1 atm. to 7 atm. is presented. The geometrical and test conditions are obtained from Young’s thesis [1]. Coupled combustion-soot simulations are performed for all the flames using steady diffusion flamelet model for combustion and Mass-Brookes-Hall 2-equation model for soot with a 2D axisymmetric mesh. Combustion-Soot coupling is required to consider the effect of soot-radiation interaction. Simulation results in the form of axial and radial profiles of temperature, mixture fraction and soot volume fraction are compared with the corresponding experimental measured profiles. The results for temperature and mixture fraction compare well with the experimental profiles. Predicted order of magnitude and the profiles of the soot volume fraction also compare well with the experimental results. The correct trend of increasing the peak soot volume fraction with increasing the operating pressure is also captured.

Commentary by Dr. Valentin Fuster
2017;():V001T04A013. doi:10.1115/GTINDIA2017-4739.

Reverse flow can annular combustor configuration becomes the inevitable option for industrial and marine gas turbine engine, due to its advantages over other configurations. The complexity associated with can annular configuration is optimum design of annular diffuser, as its flow field is dominated by downstream blockage created by transition duct geometry.

In the present study, flow behavior in the annular diffuser has been analyzed by simulating realistic downstream combustor liner and transition duct geometry. Flow analysis has been carried out using ANSYS Fluent and turbulence has been modeled using Realizable k-ε model. The diffuser is designed based on G* method, for optimum pressure recovery. Six diffuser configurations have been analyzed by varying the inner wall profile. The effect of parameters on flow field within diffuser and dump region has been studied. Also, the static pressure recovery and total pressure loss coefficient of diffuser is calculated and compared. The results show that the profile of the inner wall and the dump region needs to be tailored to get optimum performance from diffuser.

Commentary by Dr. Valentin Fuster
2017;():V001T04A014. doi:10.1115/GTINDIA2017-4816.

Oil sealing in a turbocharger is a key design challenge. Under certain engine operating conditions oil in the lubrication system is likely to enter the compressor or turbine wheel crossing the piston rings which are used to arrest the undesirable oil flow. Compressor side oil leakage can cause white smoke and particulate emissions. Limited experimental and analytical methods are available to aid the designers in developing the oil flow path. The oil flow path has dimensions of the order of a few microns in certain areas and in mm in other areas. In addition, the flow is comprised of oil and exhaust gas mixture in certain regions. The combined effects of disparate geometric length scales and two-phase flow adds to the complexity of the flow. Understanding the oil flow allows the designer to correctly size the components, flow path and also specify the appropriate clearances between for instance shaft and bearing journals. In this study a Computational Fluid Dynamics (CFD) Model has been built and validated through several experiments conducted particularly to check the oil leak through the piston rings. The study shows that CFD based models can predict within engineering accuracy the flow through leakages in a turbocharger. The importance of manufacturing tolerances on the leakages is also highlighted.

Commentary by Dr. Valentin Fuster
2017;():V001T04A015. doi:10.1115/GTINDIA2017-4838.

Fuel injector coking involves deposit formation on the external or the internal surfaces of an injector or nozzle. This deposition of carbonaceous particles can result in uneven fuel-spray characteristics or localised burning (hot spots), which may eventually lead to mechanical failure or simply have a detrimental effect on the combustion system. This study focuses on the use of numerical methods to investigate the effect of coke formation on both the atomiser internal flow passages and its spray characteristics. Three different cases are examined; one investigating the clean injector; the second investigating the effect of internal coking; and the third investigating the effect of nozzle tip coking. A pressure swirl atomiser was considered for the purpose of the study. Validation of the numerical results for the clean injector condition is carried out against published experimental data. Two arbitrary geometries of coke deposits were created. The Volume of Fluid (VOF) multiphase model has been used in conjugation with a Geometrical Reconstruction Scheme (GRS) to simulate the interface representing the two phases. Spray cone angle and the liquid film thickness for the clean injector condition predicted by numerical simulation agreed well with the experimental data. Instabilities in the air core and the spray angle were also observed because of the presence of coke layers. Fouling present on the injector tip resulted in an earlier breakup of the film which can thereby affect the flame lift-off length. These stated observations can have significant implications both on the performance as well as the life of the combustion systems, thereby establishing the relevance of this study.

Commentary by Dr. Valentin Fuster
2017;():V001T04A016. doi:10.1115/GTINDIA2017-4872.

Wind-milling occurs when air flowing through the flamed out engine results in increasing spool rotation. Aircraft forward speed and dive angles play an important role in achieving sustainable spool rotation for relight. Aircraft fuel pump connected to power take-off shaft of Engine shuts when engine rpm falls below design speed and cannot deliver pressurized fuel to Engine during wind-milling. Under this condition, engine has to suck the fuel from aircraft using its own fuel pump. The atmospheric pressure available on the aircraft fuel tank assists the engine to operate in suction mode. The datum height between engine inlet and fuel tank outlet changes with the dive angle of the aircraft.

A test set up was established in the engine test bed to vary the datum between aircraft fuel tank and Engine inlet. The datum was varied to simulate various dive angles of the aircraft. The negative gravity head (Engine fuel inlet above fuel tank outlet) between engine and fuel tank was varied in steps. Total four test cases were carried out in an engine test bed located at an altitude of 920 m above sea level. The engine was successfully started in suction mode without external assistance of fuel pump. This paper presents the test setup, comparison of engine start cycle under various dive angles of the aircraft to evaluate optimum flight conditions to attempt windmill relight.

These tests show that as the negative gravity head is increased, the time taken to start the engine increases. The slope of N2 build up becomes shallower with increase in gravity head.

Topics: Suction , Engines , Turbofans
Commentary by Dr. Valentin Fuster
2017;():V001T04A017. doi:10.1115/GTINDIA2017-4882.

Search for potential alternative jet fuels is intensified in recent years to meet stringent environmental regulations imposed to tackle degraded air quality caused by fossil fuel combustion. The present study describes atomization characteristics of blends of jatropha-derived biofuel with conventional aviation kerosene (Jet A-1) discharging into ambient atmospheric air from a dual-orifice atomizer used in aircraft engines. The biofuel blends are characterized in detail and meet current ASTM D7566 specifications. The experiments are conducted by discharging fuel spray into quiescent atmospheric air in a fuel spray booth to measure spray characteristics such as fuel discharge behavior, spray cone angle, drop size distribution and spray patternation at six different flow conditions. The characteristics of spray cone angle are obtained by capturing images of spray and the measurements of spray drop size distribution are obtained using laser diffraction particle analyzer (LDPA). A mechanical patternator system comprising 144 measurement cells is used to deduce spray patternation at different location from the injector exit. A systematic comparison on the atomization characteristics between the sprays of biofuel blends and the 100% Jet A-1 is presented. The measured spray characteristics of jatropha-derived alternative jet fuels follow the trends obtained for Jet A-1 sprays satisfactorily both in qualitative and quantitative terms.

Commentary by Dr. Valentin Fuster
2017;():V001T04A018. doi:10.1115/GTINDIA2017-4905.

Control of emissions is a big challenge plaguing the gas turbine industry for years. This necessitates new combustor designs addressing the problem. This paper discusses the characterization of a novel burner* employing Lean Direct Injection (LDI) technology for reduced pollutant emissions and improved combustion. The burner is an array of multiple swirlers arranged closely, facilitating distributed mixing of fuel and air at each swirler throughout the length of the burner. This results in a uniform and rapid mixing, thus eliminating hot spots and enabling efficient combustion. The burner thus developed is capable of operating at very lean conditions of fuel, leading to overall temperatures being low. The burner is characterized in terms of lean blow out equivalence ratio, pressure drop, average exit temperature of the burnt mixture, pattern factor and emissions — CO, CO2, unburned hydrocarbon (UHC), NOx and soot. Results show very low NOx emissions. Enhanced combustion also results in reduction in overall emissions. It overcomes the drawback of flame flashback encountered in lean premixed pre-vaporized concept. LDI is also less susceptible to combustion instability. Pressure drop across the burner is observed to be very less compared to the conventional gas turbine combustors. Thus, this concept of multi-swirl LDI burner can be a potential contender to be employed in the combustors of gas turbine engines.

Commentary by Dr. Valentin Fuster

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