ASME Conference Presenter Attendance Policy and Archival Proceedings

2015;():V001T00A001. doi:10.1115/GTINDIA2015-NS.

This online compilation of papers from the ASME 2015 Gas Turbine India Conference (GTINDIA2015) 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to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.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


2015;():V001T01A001. doi:10.1115/GTINDIA2015-1202.

Similarity theory principles are widely applied in gas dynamic design. But completely new solutions must be realized on a base of engineering approaches to predict performances. The heart of the Universal modeling engineering method is the physical model that is based on flow visualization and measurements inside rotating impellers. The math model is a sum of algebraic equations for calculation of head losses. Normalized velocity gradients along and normal to blade surfaces are arguments. Empirical coefficients values are established in a course of the identification — calculated performances are compared with the measured ones for several dozens of model stages tests with wide range of design parameters. The 4th version of the TU SPb modeling method (the set of the several computer programs) was applied in design practice in 1993–2010. Some Russian and foreign manufacturers realized several dozens of designed compressors with power up to 25 MW. The level of design point parameters prediction was so satisfactory that the manufacturers do not prove designs by model tests anymore. The whole performance prediction was not so good. The other difficulty — to predict design point efficiency with accuracy about 0.5% the very careful choice of empirical coefficients is necessary. The difficulties have been overcome in the new 5th and 6th versions. Most effective multistage compressors plant test performances were modeled successfully by 5th version program with the single set of empirical coefficients. Calculated performances and geometry of several dozens of stages of these compressors can be are applied in designs as usual model stages. The current designs are executed by the 5th and 6th version computer programs. The single stage 32 MW pipe line compressor was recently designed for the Ukrainian partner who offered high-RPM drive and favorable single-stage scheme. The test of the model at the 1:2 scale validated project parameters. Design pressure ratio and efficiency curves have matched completely. The predicted maximum efficiency 90% was proven.

Commentary by Dr. Valentin Fuster
2015;():V001T01A002. doi:10.1115/GTINDIA2015-1207.

While the effects of axisymmetric casing treatment on performance of an axial compressor stage have been extensively studied numerically as well as experimentally, the major geometrical parameters which govern these effects have been identified. Studies are now focused on understanding how each of these parameters individually impacts the performance of a casing treatment.

The present work aims to study the impact on performance of casing treatment geometry when aspect ratio of the grooves is varied in a circumferential groove casing treatment. The compressor geometry chosen for this study has design characteristics of a transonic compressor stage. Flow field solutions were derived for baseline model by solving steady state 3-D Reynolds-Averaged Navier-Stokes (RANS) equations for three grid densities and the grid independence was proved. The basic casing treatment geometry has 10 circumferential grooves of width 4mm and axial spacing of 2mm between each groove. The aspect ratio was varied by changing the depth of the grooves in each case. These casing treatment geometries were superimposed over the rotor domain with the grooves extending over the entire blade tip chord and flow field solutions were again obtained for various aspect ratios of grooves.

These results depict improvement in the range of operation in terms of mass flow rate. Results also show that the aspect ratio of the grooves significantly influences the overall effectiveness of casing treatment on the performance of compressor stage. Improvement in overall compressor efficiency is noted with lower aspect ratio casing treatments when compared to those with higher aspect ratios, however, the range improvement is higher with higher aspect ratios. It is also observed that, after a certain depth of grooves is reached, there is no significant improvement in performance on further increasing the depth and hence the aspect ratio.

Post processing results of the flow solutions are presented which confirm the trends and show that the flow behavior near rotor tip governs this effect.

Commentary by Dr. Valentin Fuster
2015;():V001T01A003. doi:10.1115/GTINDIA2015-1209.

Previous studies on circumferential groove casing treatments have shown that the effectiveness of casing Grooves highly depends on their axial location over blade tip. The present work aims to study the flow behavior and its impact on the performance of the compressor stage when the casing treatment grooves are placed to provide different axial coverage over rotor chord in each case. Geometry of a transonic compressor stage was modeled for this study. Flow field solutions for this model with smooth casing wall were obtained by solving steady state 3-D Reynolds-Averaged Navier-Stokes equations for three different grids to prove the grid independence of the solutions. Results obtained with the intermediate grid density were used as the baseline results to compare with results of casing treatment geometries. The basic casing treatment geometry has 10 circumferential groves of width 4mm, depth 16mm and axial spacing of 2mm between each groove. This casing treatment geometry was superimposed over the rotor domain with the grooves extending axially over the entire axial chord (58mm) of rotor blade tip and flow field solutions were again obtained. After that, for each case the grooves are removed from the rear side and axial coverage is shortened. Flow solutions for various axial coverage and hence for various number of grooves are thus obtained and compared. These results depict improvement in the operating range when compared to the Base-line results. Results also exhibit that as the grooves from the rear end are removed gradually, recovery in the overall efficiency is seen in compressor performance. Post processing of the flow solutions confirms the trend and shows that the grooves in the rear of the chord are almost idle not providing sufficient flow to pass over from pressure surface to suction surface of the blade and hence contributing very less towards performance enhancement.

Commentary by Dr. Valentin Fuster
2015;():V001T01A004. doi:10.1115/GTINDIA2015-1210.

This paper deals with the numerical studies on the combined effect of tip clearance and axisymmetric circumferential grooves casing treatment (CGCT) on the overall performance and stall margin of a single stage transonic axial flow compressor. Steady state numerical analysis was carried out by solving three dimensional Reynolds-averaged-Navier-Stokes (RANS) Equations using the Shear Stress Transport (SST) k-ω Turbulence Model. The numerical stall inception point was identified from the last converged point by the convergence criteria, and the stall margin was numerically predicted. Additionally, the stall margin and the isentropic peak stage efficiencies of the circumferential casing grooves with various tip clearances were compared and evaluated in order to explore the influence of the tip clearance. Results obtained were compared with those obtained on the baseline compressor with the smooth casing (SC). Further computational studies were conducted to study the role of the tip leakage flow in axial compressor in triggering the stall. The relationship between the tip clearance flow, flow field and surge margin extension from circumferential groove casing treatment with various rotor tip clearances were studied numerically. The application of the circumferential groove casing treatment with varying clearance leads to significant improvement in the operating stability of compressor with slight reduction in the isentropic peak stage efficiency for small tip clearances, whereas there was slight increment in the isentropic peak stage efficiency at higher tip clearance of 2.5 mm.

Commentary by Dr. Valentin Fuster
2015;():V001T01A005. doi:10.1115/GTINDIA2015-1211.

This paper describes the method to improve the stall margin of transonic axial flow compressor by controlling the boundary layer on the suction surface of the rotor blade tip through natural aspiration. Aspiration slots in the compressor blade are intended to energize the flow by increasing its momentum on the suction surface. This phenomenon of boundary layer control can delay the flow separation and hence results in enhancement of the stall margin of the compressor stage. Flow behavior with aspiration slots and its performance are evaluated using commercially available numerical software. Steady state RANS simulations with three dimensional implicit pressure-based coupled solver and turbulence model SST k-ω are used.

The effect of natural aspiration slot on the rotor blade performance is computed numerically. The main objective of the study was to identify the optimum location of the aspiration slot along the chord of the compressor on the rotor blade. The axial location chosen for the performance evaluations were 20%,40%,50%,60% and 70% of the rotor blade axial tip chord. By comparing the numerical simulation results with the steady state behavior in the absence of the aspirated slots, the optimized location of the aspiration slot that results in maximum stall improvement is identified. At the optimized location, natural aspiration slots on the rotor blade tip improved the stall margin with the minimum reduction in efficiency and stage pressure ratio when compared to base model. After critically understanding the performance with straight aspiration slots the compressor stage performance has enhanced further by orienting the aspiration slots. The numerical three dimensional results conclude an optimal improvement in the stall margin for the slots near the trailing edge of the rotor. The prediction shows that with the inclined aspiration slots at proper location it is possible to improve the stall margin of the compressor stage and also to restore the stage efficiency.

Topics: Axial flow
Commentary by Dr. Valentin Fuster
2015;():V001T01A006. doi:10.1115/GTINDIA2015-1213.

Laboratory “Gas dynamics of turbo machines” (LGDTM) has quite effective optimal design computer programs based on theoretic analysis and experimental data. The authors do not share an opinion that 3D impellers are superior in any case. A lot of designed compressors are provided with traditional 2D impellers with cylindrical blades disposed in a radial part of an impeller. The industrial partner tested recently 1:2 scale model of a single stage 32 MWt pipeline compressor. The flow path design is based on the medium specific speed 2D impeller. Good general scheme of the industrial partner, no constrains and profound design optimization have led to maximum efficiency 90% and to excellent performance in a whole. But if a design flow rate coefficient exceeds 0,070 … 0,08 application of 3D impeller is inevitable.

Meridian configuration and blade cascade shape of 3D impellers are much more complicated in comparison with 2D impellers. LGDTM has no at its disposal complete information on physical or numerical tests of 3D impeller candidates with different design solutions. Modern trend to apply CFD calculation for investigations to fill the gap seems to be most logical. But the authors’ own experience and published data show that CFD modeling of 3D impeller performance curves is not satisfactory. As a rule calculated performances are shifted to bigger flow rates and work coefficient is 6–9% higher. But the positive moment is that the efficiency at the design flow coefficient is predicted quite accurately. It opens a way to compare stage’s candidates at the design regimes efficiency at the design flow coefficient.

The initial design of the stage 3D impeller + vaneless diffuser + return channel with flow rate coefficient 0,105 and loading factor 0,56 is based on general principles of LGDTM: inlet velocity minimization, mean velocity deceleration control, Q-3-D non-viscid velocity diagrams with non-incidence inlet and minimal load at leading edges. CFD calculation has demonstrated necessity to apply a diffuser with tampered initial part, and better shape of the tampered part was defined. The better shape of the crossover was defined by CFD calculations too. The impeller candidates with gas dynamic and geometry principle of blade design, with different degree of flow deceleration, different axial dimension and different exit blade angles were compared.

The new 6th version of the optimal design computer programs (Universal modeling was widely presented at the conferences in Japan, Germany, Great Britain, etc.) is tuned on high flow rate stages with 3D impellers. Validation calculations demonstrated good level of performance curves modeling. The program was applied to study series of candidates with different dimensions in meridian plane. As these dimensions influence mean blade load each parameter was studied with different number of blades. Main results are: axial elongation of an impeller does not lead to efficiency grow, optimal leading edge position is at about 25% of meridian distance from an impeller inlet, optimal inlet diameter is 8,5% less that the diameter corresponding to minimal peripheral inlet velocity. The last conclusion is of particular interest and needs additional proof.

The comparison of 94 impellers candidates has led to the stage efficiency increase on about 1.5%. The results have verified general principles of design applied in the laboratory “Gas dynamics of turbo machines” and pointed out on some improvements of design principles.

Topics: Compressors , Design
Commentary by Dr. Valentin Fuster
2015;():V001T01A007. doi:10.1115/GTINDIA2015-1215.

Centrifugal compressors for gas industry consume huge amount of energy. As a rule, they are single-shaft, with two or more stages and with comparatively low pressure ratio. Compressors operate at low Mach numbers and high Reynolds numbers. Two design parameters influence mostly stage performances. Stage flow coefficient optimal values lie in range 0.060–0.11. Chosen number of stages establishes value of this coefficient if speed of a rotor rotation is fixed. Design loading factor optimal values are 0.42–0.52. It corresponds to high efficiency, shifts a surge limit far from a design point and makes power maximal in a design point. Some considerations about impeller and diffuser types are presented. Design procedure consists on application of the Universal modeling programs for main dimensions optimization and performance calculations. Q3D non-viscid velocity diagrams are analyzed for optimization of blade configuration. Samples of design are presented, 32 MW single-stage pipeline compressor stage with record efficiency included.

Commentary by Dr. Valentin Fuster
2015;():V001T01A008. doi:10.1115/GTINDIA2015-1216.

Stator part of a centrifugal compressor stage is a proper object of study by CFD calculations meaning better understanding of flow behavior, checking of field type design methods and possible improvements. Several stators with vane and vaneless diffusers for stages with different specific speed were designed by standard methodology and numerically analyzed. Results were verified. Calculation in a whole has demonstrated validity of existing recommendation. The specific velocity for stators is introduced which can be applied to match an impeller and a stator. Calculations demonstrated quick efficiency drop for stators with specific speed less than 0,215. Return channel vane cascades were studied in wide range of solidity with constant vane height and with constant radial component of velocity. Empirical formulae with non-dimensional circulation as an argument are proposed for loss coefficient, profile loss coefficient, optimal incidence angle and exit lag angle. Candidates of the low specific speed stator have demonstrated that an arbitrary channels’ wideness to diminish friction losses is not effective. Better flow organization is preferable. Modification of a crossover demonstrated positive results for high and low specific speed stators.

Commentary by Dr. Valentin Fuster
2015;():V001T01A009. doi:10.1115/GTINDIA2015-1231.

The steady and unsteady flow characteristics typically vary along and across the axial compressor stage. This coupled with asymmetric rotor tip clearance that occurs in practice makes flow even more complex. Understanding the complex flow behavior inside the transonic compressor stage will aid in developing flow control devices that are meant for purposes such as improving the rotating stall margin, flutter margin, etc. Here, a detailed time averaged numerical analysis is performed on the single stage transonic axial compressor with averaged rotor tip clearance (1.75% of rotor tip axial chord). An attempt is made to study the compressor stall phenomenon. Computational Fluid Dynamics (CFD) helps in resolving the complex flow features involved in a turbomachinery component and at transonic Mach numbers fairly well. Commercial tool ANSYS CFX is used for solving the 3D compressible Reynolds Averaged Navier-Stokes (RANS) equation with Shear Stress Transport (SST) turbulence model.

Grid independency is carried out for three different mesh size models. All mesh models chosen have fine mesh near wall boundary regions to capture the boundary layer effects. Overall performance maps of the compressor are generated for 50% to 100% rated design speeds in steps of 10% for the chosen optimum grid. Flow variations along the blade annulus are studied for three different operating conditions: choke/free flow, peak efficiency and near stall flow conditions and for different speeds.

Flow parameters such as Mach number, static and total pressure variations, etc. are studied at the inlet to rotor, exit to rotor and exit to stator for the various flow conditions and speeds. The boundary layer growth is clearly captured when the flow is throttled from choke/free flow conditions to near stall condition for all the speeds investigated. Mach number variation along blade height clearly shows decrease in Mach number as stall is approached. Blade loading distribution of the rotor at hub, mean and tip sections are clearly captured. Shock motion from around mid-chord region at free flow condition to towards the leading edge at near stall condition is clearly highlighted. Velocity streamlines near the tip section show the complex interaction of the tip leakage and clearance flows. Velocity vectors near the blade tip shows, the backflow near the trailing edge and tendency for leading edge spillage as the back pressure is increased. The flow blockage region is captured in the meridional plot and the motion of vortex core region as stall is approached is demarcated in the r-θ plots.

Tangential velocity variation across the annulus for the two flow conditions investigated shows stall initiating from the tip section of the blade as compressor is throttled. Flow compensation at near stall conditions is explained.

Commentary by Dr. Valentin Fuster
2015;():V001T01A010. doi:10.1115/GTINDIA2015-1237.

An unsteady one-dimensional dynamic model has been developed at Georgia Tech to investigate the impact of stage characteristics as well as load distribution on the compression and expansion waves that develop prior to a surge event in a multistage axial compressor. In the developed model, each of the blade rows is replaced by a duct of varying cross-sectional area with force and work source terms. The source terms model the force and energy imparted by a blade row to the working fluid. The modeling assumes the flow to be inviscid, unsteady, compressible and axisymmetric. While rotating stall cannot be explicitly modeled in a 1D mean-line method, the effect of rotating stall can be captured by a judicious choice of source terms that reflects the loss of pumping capability of a stage. Conservation of mass, momentum and energy are applied to an elemental control volume resulting in one-dimensional quasi-linear Euler system of equations. A non-uniform grid and the second-order central difference Kurganov-Tadmor (KT) scheme are used to discretize the one-dimensional computational domain. The resulting ODEs are solved with an explicit second order Runge-Kutta solver. A throttle schedule is used to introduce perturbations at a selected operating condition in order to study flow oscillations that can lead to a stall event. The current study is aimed at validation of the developed flow solver using an industrial compressor database. Further, the current study is aimed at understanding the interaction between the stages with regards to pressure oscillations leading to stall.

Commentary by Dr. Valentin Fuster
2015;():V001T01A011. doi:10.1115/GTINDIA2015-1276.

Circumferential nonuniformity of gas flow is one of the main problems that can occur in the gas turbine engine. Usually, the flow circumferential nonuniformity appears near the support, located in the flow passage of the engine. The presence of circumferential nonuniformity leads to the increased dynamic stresses in the blade rows and the blade damage. The goal of this research was to find the ways of the flow non-uniformity reduction, which would not require a fundamental changing of engine design. A new method for reducing the circumferential nonuniformity of gas flow was proposed. It has been suggested to increase the gap of the leading edges of support racks from the trailing edge of the upstream guide vane blades which will result in achieving the desired results. An important advantage of this method is that the internal cavities of racks remain unchanged for the placement of engine systems. Moreover, the proposed method allows the prediction of the pressure peak values after the rotor blades without .

Commentary by Dr. Valentin Fuster
2015;():V001T01A012. doi:10.1115/GTINDIA2015-1277.

Accurate laminar-turbulent prediction is very much important to understand the complete performance characteristics of any airfoil which operates at low and medium Reynolds number. In this article, a numerical study has been performed over two different thick airfoils operating at low Reynolds number using k-ω SST, k-kl-ω and Spalart-Allmaras (SA) RANS models. The unsteady two dimensional (2D) simulations are performed over NACA 0021 and NACA 65-021 at Re 120,000 for a range of angle of attacks. The performances of these models are assessed through aerodynamic lift, drag and pressure coefficients. To obtain better comparison, the simulated results are compared with the experimental measurements and XFOIL results as well. In this present study, it is found that the k-kl-ω transition model is capable of predicting correct lift, drag coefficient and separation bubble as reported in experiments. At high angles of attack, this model fails to predict performance variables accurately. The SA and SST models are fail to predict laminar separation bubble. However, At high angle of attack, SA model shows better predictions compared to k-kl-ω and k-ω SST models.

Commentary by Dr. Valentin Fuster
2015;():V001T01A013. doi:10.1115/GTINDIA2015-1311.

Tandem blade arrangement of axial compressors has been proposed to obtain high loading and turning compared to a single blade. The objectives of the current study is to investigate the effect of percent pitch, axial overlap and incidence angle for a low speed axial compressor stator cascade and to supplement the results with the flow structures observed. 2-D numerical study was conducted using a finite volume scheme which solves the RANS equations along with the Spalart-Allmaras turbulence model. Comparison offlow structures corresponding to different percent pitch, axial overlap and incidence angle has been made to highlight all prominent flow mechanisms. It is observed that the flow through the gap nozzle between the two blades has significant effects on losses. The incidence range of the tandem cascades is found to be superior to the corresponding single blade cases.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2015;():V001T01A014. doi:10.1115/GTINDIA2015-1313.

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. Computational Fluid Dynamics (CFD) has been extensively used to analyze the flow through rotating machineries, in general and through axial compressors, in particular. The present work is focused on the studying the effects of Vortex Generator (VG) on test compressor at CSIR National Aerospace Laboratories, Bangalore, India using CFD. The compressor consists of NACA transonic rotor with 21 blades and subsonic stator with 18 vanes. The design pressure ratio is 1.35 at 12930 RPM with a mass flow rate of 22 kg/s. Three configurations of counter rotating VGs were selected for the analysis with 0.25δ, 0.5δ and δ height, where δ was equal to the physical thickness of boundary layer (8mm) at inlet to the compressor rotor [11]. The vortex generators were placed inside the casing at 18 percent of the chord ahead to the leading edge of the rotor. A total of 63 pairs of VGs were incorporated, with three pairs in one blade passage. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 3.5% from baseline at design speed. The reasons for this numerical stall margin improvement are discussed in detail.

Commentary by Dr. Valentin Fuster
2015;():V001T01A015. doi:10.1115/GTINDIA2015-1322.

The primary function of an aero engine fuel system is to supply metered fuel to the combustion chamber at all operating conditions based on the flow demand set by the engine controller. Centrifugal pump is one of the components in fuel system that feeds fuel to the HP pump. Impeller design is a critical one which dictates the overall performance of centrifugal pump. This paper discusses about design and performance evaluation of impeller with twisted blade configuration. Impeller design is performed based on Euler’s one dimensional theory.

Steady state performance of the impeller at design and off-design operating conditions is analyzed by using commercial CFD code ANSYS-CFX with a standard SST turbulence model. The governing mathematical model for flow analysis is a three dimensional incompressible Reynolds Averaged Navier-Stokes (RANS) equation. Cavitation phenomenon is simulated in the CFD multiphase analysis to assess the pump impeller performance under cavitation at different NPSH values. Rayleigh-Plesset cavitation model is used along with RANS equation to perform cavitation analysis.

From this study, the simulation method and technique adapted is appropriate for predicting the performance of impeller of centrifugal pump with or without cavitation. Performance of the impeller reduces drastically as there is decrease in NPSH. The prediction of critical NPSH is vital for safer operation of the pump, specifically at high altitude the pump inlet pressure falls which may result in cavitation during operation.

Commentary by Dr. Valentin Fuster
2015;():V001T01A016. doi:10.1115/GTINDIA2015-1333.

The benefits of minimizing the weight of an aircraft are substantial, due to which all aircraft components are designed so as to perform acceptably with minimal weight. As a result, modern compressor, fan and turbine blades are increasingly being designed with thinner airfoil profiles, while also being subjected to high loading, so as to improve the thrust-to-weight ratio of the engine. These conditions make the blades extremely likely to result in large-amplitude vibrations. A class of critical vibrations are caused due to aeroelastic instabilities within the turbomachinery blade rows. Forced response or the self-excited flutter can lead to high-cycle fatigue which can cause a catastrophic failure of the blades. The absence of a reliable prediction methodology for the occurrence of these instabilities signify the importance of unsteady aerodynamic studies in turbomachine cascades. The present experiments are conducted on a newly commissioned annular cascade tunnel. The test section consisting of 14 compressor blades is studied under subsonic conditions. The profiles of the blades have a constant span design, and the cascade parameters are chosen as EPFL’s Second Standard Test Configuration at the mid-span location. A set of guide vanes upstream of this blade row set the required incidence to the cascade by imparting a circumferential component to the velocity. For the present unsteady studies, selected blades are subjected to controlled vibrations while the unsteady response on a reference stationery blade is measured. Two blades are connected to individual servo motors through a mechanism so as to execute controlled, low-amplitude, torsional (pitching) oscillations about its mid-chord axis. A reference blade is mounted on a load cell to enable measurement of both the axial and transverse forces and moments. In order to simulate the effect of the inter-blade phase angle prevalent in a rotating turbomachine blade row, the phase angle between the vibrating blades is varied to all admissible values. The fundamental parameters for evaluating the stability are the phase difference between blade position and the forces responses. This is estimated from the Fourier transform of the displacement and load signals. The parameters are evaluated for a range of reduced frequencies and inlet velocities to evaluate the stability of the cascade at all specified flow conditions.

Commentary by Dr. Valentin Fuster
2015;():V001T01A017. doi:10.1115/GTINDIA2015-1334.

The leading edge serrations are a type of passive flow control techniques in a compressor cascade. They are particularly attractive as they have been observed to increase the stall angle. This stall postponing character of the serrations is helpful in preventing compressor surge and widens the operational window of the compressor. Due to the simpler geometry of the serration type used in this study, it can be easily implemented onto the existing compressor blades. An experimental study on the flow modifications and losses due to these serrations are conducted in a linear cascade tunnel. The experiments are conducted on blades of NACA 65209 airfoil with and without leading edge serrations at Re of 120,000. Four serration profiles of various width and amplitude are compared. End plane measurements taken with 5-hole probe are studied for the better serration profile and surface flow visualizations are conducted to study the variation in the surface flow pattern on the suction side. The surface flow visualization reveals the presence of local recirculation zones and stream wise vortices created from each wave of the serration leading to flow attachment. These serrated blades have higher losses at 0 deg incidence; the reason for the same is found to be the flat leading edge surfaces formed from serration.

Commentary by Dr. Valentin Fuster
2015;():V001T01A018. doi:10.1115/GTINDIA2015-1337.

The operating range of any compressor is controlled by Surge and Choke. Surge occurs at lower mass flow rates with large pressure fluctuations and flow reversals, while choke occurs at higher mass flow rates when the flow rate reaches the limit which compressor can discharge. Ported shroud is a cost effective casing treatment that can greatly improve operating range of centrifugal compressors. By removing the stagnant and reverse flow from shroud wall boundary-layer region and recirculating it to impeller inlet, it has been demonstrated that larger range of operability can be achieved without much loss on compressor efficiency. This paper demonstrates the improvement of a centrifugal compressor operational range with ported shroud configuration.

A series of CFD simulations were carried out with open source centrifugal compressor geometry (NASA HPCC 4:1) to create performance characteristics/speed-lines. The CFD methodology and practices were validated by comparing the results with the experimental data. Performance evaluation of ported shroud configuration is done with respect to solid shroud.

Ported shroud compressor is proven to give higher choke mass flow and also a better surge margin compared to the Solid shroud model. The phenomena of in-flowing and out-flowing port have also been demonstrated. Emphasis was given to understand how ported shroud helps to achieve a better performance. A design optimization study has also been carried out in order to establish the optimum ported shroud configuration. Design parameter such as port location has been selected and the effect of this parameter on the performance of the compressor is studied using CFD. Optimum port geometry was proposed.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2015;():V001T01A019. doi:10.1115/GTINDIA2015-1341.

An industrial axial compressor has to meet a wide range of operation requirements. These machines have to run continuously for four to five years before going for overhaul. Hence, overall high level of efficiency may be slightly relaxed to meet this requirement. This requires axial flow compressor design to be more conservative and flexible to accommodate changes required for process industry through modern design & development approaches.

This paper deals with finding of optimum flow path configuration that will allow a successful detailed design to follow. The effect of various parameters such as hub to tip ratio, proper selection of design rpm, reactions, work coefficient & flow coefficient has been investigated and selected for optimal performance of the machine. Last stage of the compressor is selected as radial stage with the advantage of reduction in axial length and to provide radial outlet, which is more suitable outlet configuration. Meanline design and streamline analysis for each configuration is determined to find out good operating range (stall-free operation) before starting the detailed design.

Topics: Design , Axial flow
Commentary by Dr. Valentin Fuster
2015;():V001T01A020. doi:10.1115/GTINDIA2015-1345.

Centrifugal compressors used in applications like enhanced oil recovery using gas re-injection and Carbon capture and sequestration operate at very high pressures and often have to to deal with supercritical CO2 which is considerably viscous. Increased viscosity leads to energy dissipation and introduces damping in the acousto-elastic interaction between supercritical CO2 and the impeller of a centrifugal compressor, thereby altering the frequency response of the system especially near resonant frequencies. In this paper, the damping introduced by visco-thermal effects in such acousto-elastic systems is accounted for in a numerically efficient manner.

The acoustics in the fluid are modelled using the Boundary Layer Impedance (BLI) model and the centrifugal impeller as a linear elastic structure. The coupled acousto-elastic system is then solved using the finite element method. The finite element solution becomes computationally expensive especially when working with large three-dimensional models. In order to reduce the computational cost, a model order reduction technique based on a multi-point second-order Arnoldi (SOAR) procedure is developed. It is demonstrated that the reduced order model significantly brings down the computational time while being sufficiently accurate.

Commentary by Dr. Valentin Fuster
2015;():V001T01A021. doi:10.1115/GTINDIA2015-1350.

Return channel de-swirl vanes form an integral part of a centrifugal compressor stage for multi-stage configuration. In this paper, a few configurations of return channel vanes (RCV) are arrived at by modifying the blade angle and thickness distribution from leading edge to the trailing edge. Influence of these two parameters on the overall performance of return channel in terms of total pressure loss co-efficient and static pressure recovery co-efficient along with stage exit flow angle are evaluated through CFD analysis.

CFD results show that, proper thickness distribution after maximum thickness point to the trailing edge improves the stage exit flow angle but not the total pressure loss co-efficient and static pressure recovery co-efficient. Whereas, by suitably modifying the blade angle distribution, all the three performance parameters can be improved considerably.

Topics: Compressors , Design , Blades
Commentary by Dr. Valentin Fuster
2015;():V001T01A022. doi:10.1115/GTINDIA2015-1351.

The modern engine has the requirement of high pressure ratio compressors. High diffusion blades are used to cater to this requirement. The high diffusion blades suffer from the low incidence range. A variable geometry inlet guide vane is used to improve the incidence range and to have an increased stable operating range.

In this paper a variable camber inlet guide is proposed in place of an existing inlet guide vane (IGV) to operate the compressor at increased stable operating range or to operate at improved efficiency at off design point. Numerical analysis is carried out in ANSYS CFX©. The existing compressor consists of IGV (20 blades) , rotor (43 blades) and stator (52 blades). The rotor rotates at 2400 rpm in clockwise direction.

The IGV blade is split two part forward blade and aft blade. Numerical studies are conducted to study the effect of varying the stagger angle on the performance of the compressor. The aft blade is given rotation in clockwise direction for +5° and +10°. The numerical results obtained are compared to the same stagger angle with full blades. It is observed that marginal improvement in the pressure ratio and efficiency. 7% stall margin improvement is achieved with slotted blade in place a fixed IGV at 0° setting angle. A new compressor characteristics is estimated which shows that the compressor can be operated to the left of the fixed-IGV-stage peak pressure with high efficiency.

Commentary by Dr. Valentin Fuster
2015;():V001T01A023. doi:10.1115/GTINDIA2015-1356.

Since the production of oil and gas is declining, the development and implementation of new, cost-efficient technology is an important focus area in order to enhance production and recovery from existing fields. The wet gas compressor represents a key element because it has been identified as one of the most relevant technologies to be developed.

In recent years, research on subsea installations has been carried out. A traditional centrifugal compressor may not be applicable to handle liquid during a standard process. A complete comprehension of the machine behaviour, in particular investigation with the presence of gas-liquid mixture is essential. Because of liquid impact, compressor performances and consequently the margin of stability might be modified. Delayed instability inception should also be pointed out.

The objective of this paper is to provide a consistent basis for the surge and rotating stall analysis of wet gas compressors and describe the mechanism leading to instabilities and their evolution. Visualisation of flow in the compressor stage is accomplished by means of fluorescent injection in correspondence with the impeller labyrinth seal and diffuser section. Simultaneously, a fast Fourier transform examination is realised in order to identify characteristic frequencies of unsteady events.

Commentary by Dr. Valentin Fuster
2015;():V001T01A024. doi:10.1115/GTINDIA2015-1369.

This paper describes the computational results on the performance of a centrifugal compressor stage with Vaneless diffuser (VLD) and low solidity vaned diffuser (LSVD) by varying blade shape and its setting angle. The centrifugal compressor stage configuration consists of a 2-D impeller with a diffuser. Analysis was conducted for VLD and four different blade shapes of LSVD namely Un-cambered constant thickness flat plate (FP), Cambered curved arc constant thickness plate (CCAP), Un-cambered aerofoil profile NACA 0010(NACA 0010) and Cambered aerofoil profile NACA 2410 (NACA 2410) at five different setting angles ranging from 16° to 32° in steps of 4° for each blade. The study is conducted at five different flow coefficients, at 0.8, 0.9, 1.0, 1.1 and 1.2 of design mass flow rate representing the design and off design cases for VLD and LSVD. CFD results are validated with experimental results for stages with VLD and LSVD for certain chosen performance parameters such as head coefficient, stage input power and exit flow angle. The computational results indicate that variations in diffuser vane geometry and its setting angle causes changes in all significant performance parameters like the total head coefficient, total-to-static stage efficiency, power coefficient of the stage and static pressure recovery coefficient of the diffuser. Contour plots were generated from CFD results and analyzed for better understanding of effect of diffuser vane shape and its setting angle on the performance of the centrifugal compressor. As a result of this study, it can be concluded that the centrifugal compressor shows improved performance characteristics for chosen blade shape of low solidity vaned diffuser than VLD.

Commentary by Dr. Valentin Fuster
2015;():V001T01A025. doi:10.1115/GTINDIA2015-1383.

A Counter-Rotating System (CRS) is composed of a front rotor and a rear rotor which rotates in the opposite direction. Compared with traditional rotor-stator system, the rear rotor is used not only to recover the static head but also to supply energy to the fluid. Therefore, to achieve the same performance, the use of a CRS may lead to a reduction of the rotational speed and may generate better homogeneous flow downstream of the stage. On the other hand, the mixing area in between the two rotors induces complicated interacting flow structures.

Blade sweep has attracted the turbomachinery blade designers owing to a variety of performance benefits it offers. However, the effect of blade sweep on the performance, stall margin improvements whether it is advantageous/disadvantageous to sweep one or both rotors has not been studied till now. In the current investigation blade sweep on the performance characteristics of contra rotating axial flow fans are studied. Two sweep schemes (axial sweeping and tip chord line sweeping) are studied for two sweep angles (20° and 30°). Effect of blade sweep on front rotor and rear rotor are dealt separately by sweeping one at a time. Both rotors are swept together and effect of such sweep scheme on the aerodynamic performance of the stage is also reported here. The performance of contra rotating fan is significantly affected by all these parameters. Blade sweep improved the pressure rise and stall margin of front rotors. Axially swept rotors are found to have higher pressure rise with reduced incidence losses near the tip for front rotors. Sweeping the rear rotor is not effective since the pressure rise is less than that of unswept rotor and also has less stall margin.

Topics: Aerodynamics , Fans , Blades
Commentary by Dr. Valentin Fuster


2015;():V001T02A001. doi:10.1115/GTINDIA2015-1221.

Compound lean implemented on stator of an industrial steam turbine stage in order to reduce secondary losses are discussed. Baseline stator is a prismatic vane with aspect ratio of unity. Compound lean stator blade is designed by shearing the airfoil sections in tangential direction. Modifications are analyzed numerically using commercial code CFX. Three blade rows i.e. one complete stage with a downstream stator are analyzed. Steady state Reynolds averaged Navier Stokes equations are solved. Total pressure loss (TPL) is used as objective function to monitor reduction in secondary losses. Rotor is retained the same for baseline as well as compound leaned stator. Results show reduction in total pressure loss of stator in excess of 5 %. Also, computations of co-efficient of secondary kinetic energy shows significant reduction in secondary losses in excess of 30 % in stator. Efficiency gained by implementation of compound lean are discussed.

Topics: Steam turbines
Commentary by Dr. Valentin Fuster
2015;():V001T02A002. doi:10.1115/GTINDIA2015-1225.

Unsteady transonic flows in diffuser have become increasingly important, because of its application in new propulsion systems. In the development of supersonic inlet, air breathing propulsion systems of aircraft and missiles, detail investigations of these types of flow behavior are very much essential. In these propulsion systems, naturally present self-sustaining oscillations, believed to be equivalent to dynamically distorted flow fields in operational inlets, were found under all operating conditions. The investigations are also relevant to pressure oscillations known to occur in ramjet inlets in response to combustor instabilities. The unsteady aspects of these flows are important because the appearance of undesirable fluctuations generally impose limitation on the inlet performance. Test results of ramjet propulsion systems have shown undesirable high amplitude pressure fluctuations caused by the combustion instability. The pressure fluctuations originated from the combustor extend forward into the inlet and interact with the diffuser flow-field. Depending on different parameters such as the diffuser geometry, the inlet/exit pressure ratio, the flow Mach number, different complicated phenomena may occur. The most important characteristics are the occurrence of shock induced separation, the length of separation region downstream of the shock location, and the oscillation of shock location as well as the oscillation of the whole downstream flow. Sajben experimentally investigated in detail the time mean and unsteady flow characteristics of supercritical transonic diffuser as a function of flow Mach number upstream the shock location and diffuser length. The flows exhibited features similar to those in supersonic inlets of air-breathing propulsion systems of aircraft. A High-order LES turbulence model developed by the author is assessed with experimental data of Sajben on the self-excited shock oscillation phenomena. The whole diffuser model configuration including the suction slot located at certain axial location around the bottom and side walls to remove boundary layer, are included in the present computation model. The time-mean and unsteady flow characteristics in this transonic diffuser as a function of flow Mach number and diffuser length are investigated in detail. The results of study showed that in the case of shock-induced separation flow, the length and thickness of the reverse flow region of the separation-bubble change, as the shock passed through its cycle. The instabilities in the separated layer, the shock /boundary layer interaction, the dynamics of entrainment in the separation bubble, and the interaction of the travelling pressure wave with the pressure fluctuation region caused by the step-like structure of the suction slot play very important role in the shock-oscillation frequency.

Commentary by Dr. Valentin Fuster
2015;():V001T02A003. doi:10.1115/GTINDIA2015-1245.

An investigation of five models used to assess the profile losses in axial turbine cascades appears in this article: Soderberg model, Ainley&Mathieson model, Dunhem&Came model, Kaker&Ocapu model and Central Institute of Aviation Motors (CIAM, Russia) model. Using them, the calculation results were compared with experimental data for 170 airfoil cascades of axial turbines. These cascades include a diversity of blade profiles of axial turbines used in aircraft gas turbine engines. Direct comparison of the calculated and experimental results did not make it possible to uniquely choose the best model. For this reason, the analysis method of loss models based on the statistical analysis of calculation and experimental data deviation was developed. It is shown that the deviations are subject to the normal distribution law. Based on the analysis of mathematical expectations μΔζ and standard deviation σΔζ, it was found that CIAM model gives the results closest to the experimental data. It shows the deviation from the real values of the loss 2±82% with a probability of 95%.

Commentary by Dr. Valentin Fuster
2015;():V001T02A004. doi:10.1115/GTINDIA2015-1372.

A specific design of mixed flow variable geometry turbine for an automotive sub 1.5 litre diesel engine turbocharger is proposed in this paper. An experimental set up is developed for measuring the steady state and transient response behaviour of the turbine at different nozzle vane opening positions. The rotor speed, pressure and temperature before and after the turbine are measured and recorded using high frequency data logging system. The steady state performance for mass flow, efficiency, velocity ratio, specific speed and the transient response behaviour of the mixed flow variable geometry turbine (MFVGT) are compared against the same parameters of a radial flow variable geometry turbine (RFVGT) of similar dimensions. Typical result indicates that the transient response of the MFVGT is faster by about 350 milliseconds than the radial at turbine inlet pressure of 0.2 bar (g).

Commentary by Dr. Valentin Fuster

Combustion, Fuels and Emissions

2015;():V001T03A001. doi:10.1115/GTINDIA2015-1212.

Experiments were performed on the central pilot body (RPL-rich-pilot-lean) of Siemens prototype 4th generation DLE burner to investigate the flame behavior at atmospheric pressure condition when varying equivalence ratio, residence time and co-flow temperature. The flame at the RPL burner exit was investigated applying OH planar laser-induced fluorescence (PLIF) and high-speed chemiluminescence imaging. The results from chemiluminescence imaging and OH PLIF show that the size and shape of the flame are clearly affected by the variation in operating conditions. For both preheated and non-preheated co-flow cases, at lean equivalence ratios combustion starts early inside the burner and primary combustion comes to near completion inside the burner if residence time permits. For rich conditions, the unburnt fuel escapes out through the burner exit along with primary combustion products and combustion subsequently restarts downstream the burner at leaner condition and in a diffuse-like manner. For preheated co-flow, most of the operating conditions yield similar OH PLIF distributions and the flame is stabilizing at approximately the same spatial positions. It reveals the importance of the preheating co-flow for flame stabilization. Flame instabilities were observed and Proper Orthogonal Decomposition (POD) is applied to time resolved chemiluminescence data to demonstrate how the flame is oscillating. Preheating has strong influence on the oscillation frequency. Additionally, combustion emissions were analyzed to observe the effect on NOX level for variation in operating conditions.

Commentary by Dr. Valentin Fuster
2015;():V001T03A002. doi:10.1115/GTINDIA2015-1249.

In the reaction zone of flame, electronically excited species are formed such as CH*, OH* etc. During de-excitation these radicals emit electromagnetic radiation of certain wavelength. This process is called chemiluminescence. The intensity of chemiluminescence, is in general captured using a photo multiplier tube (PMT), which is used to measure unsteady heat release rate from premixed flames. This technique is well established and is now a standard for unsteady heat release rate measurements in the parlance of combustion instability, however has certain limitations. In fuel rich mixtures, unreacted heated carbon emits broad band black body radiation, which in some cases large enough to mask the chemiluminescence signal. Hence, this technique is not valid for fuel rich conditions. Moreover, it cannot be applied, when the heat source is diffusion/partially premixed flames or electrically heated wires. We propose an alternative in this regard: two microphone technique. In this technique, we relate the acoustic velocity jump across the heat source to measure the unsteady heat release rate. The up and downstream acoustic velocity, in turn is obtained by two microphone technique. Experiments are performed in a premixed multiple flame burner at fuel lean conditions. This burner is enclosed in a duct, which acts as an acoustic resonator. Results indicate that the magnitude of the unsteady heat release rate obtained from both the techniques is found to agree within 18 %. Experiments are conducted for various lengths of the duct, thereby changing the oscillating frequency. This method is valid as long as the heat source is compact in comparison to the duct, which is true in most of the combustors during combustion instability and is irrespective of its type.

Commentary by Dr. Valentin Fuster
2015;():V001T03A003. doi:10.1115/GTINDIA2015-1278.

In many combustion systems, fuel atomization and the spray breakup process play an important role in determining combustion characteristics and emission formation. Due to the ever-rising need for better fuel efficiency and lower emissions, the development of a fundamental understanding of its process is essential and remains a challenging task. The Spray-A case of the Engine Combustion Network (ECN) is considered in the study, in which liquid n-Dodecane (Spray-A) is injected at 1500 bar through a nozzle diameter of 90 μm into a constant volume vessel with an ambient density of 22.8 kg / m3 and an ambient temperature of 900 K. The unsteady Reynolds averaged Navier-Stokes (URANS) in conjunction with k-ε turbulence model is used to investigate the flow physics in a two-dimensional axisymmetric computational domain. A reduced chemical mechanism from Wang et al. [1] with 100 species and 432 reactions is invoked to represent the kinetics. The gas and liquid phases are modeled using Eulerian-Lagrangian coupled approach. The present model is validated with the experimental data as well as computational data of Pei et al. [2]. Initially, the effects of various turbulence models with modified constants are examined without introducing the breakup phenomena in the computational physics. Later on, primary and secondary breakup processes of the liquid fuel are taken into account. In the present study, we examine the effects of secondary breakup modeling on the spray under high-pressure conditions using different breakup models, including Wave, Kelvin-Helmholtz and Rayleigh-Taylor (KH-RT) and Stochastic Secondary Droplet (SSD) models. It has been observed that KH-RT model is more dominant in such high-pressure sprays and predict physics more accurately as compared to other models. The dominance of convection as well as diffusion controlled vaporization model is also realized over the diffusion controlled vaporization model. The investigations at different fuel injection pressures are also modeled and validated with the experimental data [3]. The results strongly suggest that applying high-pressure, leads to high injection velocity and momentum which enhances the air entrainment near the injector region and the mixing process.

Commentary by Dr. Valentin Fuster
2015;():V001T03A004. doi:10.1115/GTINDIA2015-1339.

This study deals with the investigations on the sources and the control of combustion noise, in an atmospheric gas turbine combustor. Combustion noise encountered here is also termed as hooting, as it occurs within a limited bandwidth of frequencies ranging from 300–450 Hz. Combustion noise is usually classified as direct and in-direct combustion noise. The present study emphases on the direct combustion noise which occurs when a volume of gas expands at constant pressure, as soon as it is heated by combustion; this results in a sound wave which propagates outside the boundary of the flame. At certain conditions, if the unsteady heat release rate drives the acoustic oscillations, satisfying Rayleigh criterion, pressure oscillations grow leading to discrete tonal sound and this phenomena is termed as combustion instability.

Experiments are conducted in a liquid fuelled swirl stabilized atmospheric gas turbine combustor, whose aspect ratio is 2.5, combustion intensity varies from 25MW/m3 atm to 50MW/m3 atm. Air is passed through various stages: primary, secondary, quenching and atomizing air. Aviation turbine fuel is injected through an air-blast atomizer. An unsteady pressure transducer is located at the primary zone to measure the acoustic oscillations. The frequency of sound generated during the combustion process is compared with a microphone located at 1.25 m away from the combustor at an angle of 45° from the axis of the combustor.

The main objective of this paper is to investigate the cause of hooting and the passive control techniques in order to reduce it. This is achieved by two ways, systematically by two ways, i) reducing the quenching air methodologically reducing the quenching air and ii) varying the air to liquid (fuel) ratio (ALR). By imparting these methods the sound pressure level inside the combustor is reduced from 143dB to 128dB. Since, combustion noise occurs in a broad-band of frequencies, the unsteady data obtained with the pressure transducer are analyzed using octave bands, where it shows linear decrement of energy present in-between the two frequencies. In order to perform parametric study, two swirlers of swirl numbers.0.77 and 0.86 are used. Established by the physics of Helmholtz oscillator, the combustor hooting is dictated, by which the sound produced by the combustor is explained.

Commentary by Dr. Valentin Fuster
2015;():V001T03A005. doi:10.1115/GTINDIA2015-1340.

The presented study is on a laboratory scaled industrial gas turbine combustor of intensity 25MW/m3 atm, where an open loop active control technique is investigated. Combustion noise is classified as direct and in-direct combustion noise. The present study is focused on the investigation of direct combustion noise. It occurs when the volume of the gas fluctuates due to the fluctuations in heat release rate, caused perhaps due to flow turbulence. This results in sound waves, which propagate outside the boundary of the flame. The radiated acoustic waves are reflected from the boundaries of the combustion chamber, perturbing the fuel flow rate and hence the spray characteristics. This eventually leads to perturbation in the heat release rate and thus a feedback loop is established. At certain conditions, if the unsteady heat release rate drives the acoustic oscillations, satisfying Rayleigh criterion, pressure oscillations grow leading to discrete tonal sound and this phenomena is termed as combustion instability.

Experiments are performed in a scaled down swirl stabilized liquid fueled gas turbine combustor, where a new scheme for open-loop control of combustion noise using periodic fuel injection is employed without drastically altering the combustor design or forfeiting its performance. Fuel is modulated in the frequency range of 0.6 to 5 Hz with various duty cycles [25–75%] using square wave. Fuel modulation is achieved by passing fuel through a direct current (DC) powered solenoid valve, which is being controlled using a custom-made circuit. The modulated fuel enters the combustor through an air-blast atomizer and is metered through a turbine flow meter.

The main objective of this paper is to investigate the potential of active control to reduce combustion noise in laboratory scaled gas turbine combustor. Pressure transducer is used to capture the sound pressure level inside the combustor. A reduction in overall sound pressure level of 14dB is achieved by modulating fuel with 50% duty cycle at 1.5Hz.

Commentary by Dr. Valentin Fuster
2015;():V001T03A006. doi:10.1115/GTINDIA2015-1342.

This paper presents an experimental study of primary breakup of liquid jet in an annular passage in a cross flow of air at a fixed Mach number of 0.12, at atmospheric pressures. The experiments were conducted for various velocities of liquid jet from 1.417 m/s to 7.084 m/s (based on orifice diameter = 1 mm) and the corresponding liquid-air momentum flux ratios varied from 1 to 25. The droplet sizes and velocities were measured using a Phase Doppler Particle Analyzer (PDPA) downstream of the liquid inlet port along the axial direction at the centerline of the annular passage along the plane of injection. Observed droplet sizes and velocity variations at different momentum flux ratios, in the axial direction, show three distinct zones. The first zone is the ligament formation zone represented by large variation in droplet Reynolds number with momentum ratio. The second zone is the primary droplet formation zone in which a fairly monotonic decrease in droplet size and droplet acceleration due to the breakup is observed. However, the Reynolds number of the droplets is almost invariant with momentum ratio. The third zone is where the spray attains the critical state where the size and velocity does not vary in the axial direction and the variation in size in this zone with the momentum ratio is primarily due to the initial conditions established in the ligament formation zone.

Commentary by Dr. Valentin Fuster
2015;():V001T03A007. doi:10.1115/GTINDIA2015-1347.

An experimental study has been conducted to investigate the interaction between the conical spray produced by simplex atomizer and the swirling flow from an axial swirler. This work has been carried out in an unconfined ambience at isothermal conditions, using water. Malvern spray analyzer with a three dimensional traverse is used to characterize the swirling flow and spray interactions at various axial and radial locations. Images of spray at different conditions of air and water mass flow rates have been analyzed. Increasing the air mass flow through swirler at constant water flow rate, changes the spray structure significantly. These structural changes are sudden and highly dependent on the initial conditions of the spray. At smaller air flow rates, single-mode droplet size distribution at mid-plane changes into a bi-modal distribution at an air flow rate of about 35 kg/hr, with higher contribution of larger droplets. With further increase in air flow rate (90, 110 and 130 kg/hr), the bi-modal size distribution is maintained but with a larger volumetric fraction of small droplets. At different axial distances, the droplet size distributions are similar (single mode and bimodal distributions depending on air flow rate). But volume percentage of larger droplets is less compared to those of smaller droplets, at larger axial distance. At outer radial locations of the spray, volume percentage of larger droplets reduces and that of smaller droplets increases significantly, due to secondary droplet breakup. The interaction between the swirl and spray causes droplets to move radially outwards, resulting in droplet break-up by impact on the dome. Cases with higher air to water flow ratios exhibit significant changes in drop size distribution due to such swirl-spray interactions.

Commentary by Dr. Valentin Fuster
2015;():V001T03A008. doi:10.1115/GTINDIA2015-1406.

Dimension reduction is a popular and attractive approach for modeling turbulent reacting flow incorporating finite rate chemistry effects. One of the earliest and most popular approaches in this category is the Laminar Flamelet Model (LFM), which represents the turbulent flame brush using statistical averaging of laminar flamelets whose structure is not affected by turbulence. The other common reduction approach is the intrinsic low dimensional manifold (ILDM). While, the LFM has limitations in predicting the non-equilibrium effects, the ILDM model suffers in the prediction of the low temperature kinetics. A combination of the two approaches where flamelet based manifold are generated called, Flamelet Generated Manifold (FGM) model considers that the scalar evolution in a turbulent flame can be approximated by the scalar evolution similar to that in a laminar flame. This model does not involve any assumption on flame structure. Therefore, it can be successfully used to model ignition, slow chemistry and quenching effects, which are far away from equilibrium. In the FGM, the manifold can be created using different flame configurations. For premixed flames, 1D unstrained flamelets are solved in reaction-progress space. In the case of diffusion flames, a counter flow configuration is used to generate a series of steady flamelets with increasing scalar dissipation and also an unsteady laminar flamelet is generated to create the diffusion FGM manifold. In the present work, a diffusion flamelet based FGM model is compared with the FGM model using premixed unstrained flamelet configurations. The performance and predictive capabilities of the two approaches are compared for a turbulent lifted methane flame in a diluted hot co-flow environment, where the reacting flow associated with the central jet exhibits similar chemical kinetics, heat transfer and molecular transport as recirculation burners without the complex recirculating fluid structures. It is observed that though the diffusion flamelet based FGM predicts a lifted flame, but the lift off height is lower compared to the premixed configuration. A parametric study with different normalization for the progress variable is done to study its impact on the flame characteristics and the manifold created. Finally, the computations are performed for different definitions of the progress variable from previously published works. It is seen that the results are sensitive to the various progress variable definitions, particularly when the number of species are higher and involve different time scales.

Commentary by Dr. Valentin Fuster

Heat Transfer

2015;():V001T04A001. doi:10.1115/GTINDIA2015-1201.

This paper deals with a computational study to predict important dimensions of a rectangular fin used in gas turbine blade cooling for satisfying a prescribed internal heat generation. The heat transfer is assumed to occur by simultaneous conduction, convection and radiation. The effect of temperature-dependent thermal conductivity has been also taken into consideration. Rectangular fin geometry has been considered due to its simplicity and easiness of fabrication. Corresponding to known values of various thermo-physical parameters, at first using the fourth order implicit Runge-Kutta-based forward method, the relevant steady-state temperature distribution is evaluated. Forward method has been well-validated with three numerical schemes and experimental data. Thereafter, an inverse problem is solved using the genetic algorithm (GA) for predicting fin dimensions satisfying a prescribed temperature distribution corresponding to a fixed internal heat generation. The relevant objective function has been formulated using a three-point error minimization technique represented by square of residuals between guessed and available temperature distributions. The analysis has been done for three different fin materials such as Inconel, Hastelloy and Titanium. These materials are generally used in gas turbine blade applications due to their high melting point along with good fatigue, corrosion and creep properties. Effects of random measurement errors following a Gaussian profile are analyzed. The variations of relevant parameters are studied at different generations of GA. It is observed that for a given fin material, many feasible dimensions can sustain a given amount of internal heat generation which offer sufficient scopes to the fin designer. For the required amount of heat generation, the suitability of estimated parameters has been verified by the comparison between actual and reconstructed temperature distributions alongwith minimization of total fin volume. The present work is proposed to be useful in selecting appropriate dimensional fin configurations corresponding to a given material which can satisfy a fixed amount of internal heat generation.

Topics: Heat , Dimensions
Commentary by Dr. Valentin Fuster
2015;():V001T04A002. doi:10.1115/GTINDIA2015-1206.

Established numerical approaches for performing detailed flow analysis happens to be an effective tool for industry based applied research. In the present study, computations are performed on multiple gas turbine combustor geometries for turbulent, non-reactive and reactive swirling flow conditions for an industrial swirler. The purpose of this study is to identify the location of peak convective heat transfer along the combustor liner under swirling inlet flow conditions and to investigate the influence of combustor geometry on the flow field. Instead of modeling the actual swirler along with the combustor, an inlet swirl flow profile is applied at the inlet boundary based on previous literature. Initially, the computed results are validated against available experimental data for an inlet Reynolds number flow of 50000 using a 2D axi-symmetric flow domain for non-reacting conditions. A constant heat flux on the liner is applied for the study. Two turbulence models (RNG k-ε and k-ω SST) are utilized for the analysis based on its capability to simulate swirling flows. It is found that both models predict the peak liner heat transfer location similar to experiments. However, k-ε RNG model predicts heat transfer magnitude much closer to the experimental values except displaying an additional peak whereas k-ω model predicts only one peak but tends to over-predict in magnitude. Since the overall characteristic liner heat transfer trend is captured well by the latter one, it is chosen for future computations. A 3D sector (30°) model results also show similar trends as 2D studies. Simulations are then extended to 3 different combustors (Case 1: full cylinder and Case 2 and 3: cylinders with downstream contractions having reduced exit areas) by adopting the same methodology for same inlet flow conditions. Non-reacting simulations predict that the peak heat transfer location is marginally reduced by the downstream contraction of the combustor. However the peak location shifts towards downstream due to the presence of accelerated flow.

Reacting flow simulations are performed with Flamelet Generation Manifold (FGM) model for simulating premixed combustion for the same inlet flow conditions as above. It is observed that Case 3 predicts a threefold increase in the exit flow velocity in comparison to non-reacting flow simulations. The liner heat transfer predictions show that both geometries predict similar peak temperatures. However, only one fourth of the initial liner length experiences peak temperature for Case 1 whereas the latter continues to feel the peak till the end. This behavior of Case 3 can be attributed to rapid convection of high temperature products downstream due to the prevailing accelerated flow.

Commentary by Dr. Valentin Fuster
2015;():V001T04A003. doi:10.1115/GTINDIA2015-1296.

Advanced gas turbines are designed to operate at increasingly higher inlet temperature that poses a greater challenge to the designer for more effective blade cooling strategies. In this paper, a generic high-pressure turbine (HPT) blade of a gas turbine, which is cooled by film cooling in conjunction with internal convective cooling, has been analysed by solving Navier-Stokes and energy equations. The intricate internal cooling passages and a series of holes on the suction surface are considered for the simulations. Large numbers of cell in different zones are used to truly replace the blade with cooling holes and the internal cooling passage. The CFD analysis with conjugate heat transfer condition is accomplished by Fluent, version 6.3. A detailed discussion has been made regarding the aerodynamics and heat transfer. In brief, the suction surface is well protected by film cooling, whereas, the pressure surface demands some additional protection for a longer life. The leading edge is under the metallurgical limit because of internal cooling for the present configuration.

Commentary by Dr. Valentin Fuster
2015;():V001T04A004. doi:10.1115/GTINDIA2015-1319.

The impingement/effusion cooling is a method of using cooling air to protect the hot combustor liner surfaces from high temperature effectively. This paper investigates the impingement/effusion cooling over two perforated flat plates and proposes a better cooling scheme for high temperature combustion liners in aircrafts and electrical power generation application. The adiabatic cooling effectiveness distribution over the liner surface is numerically studied by control volume technique in CFD. In this hybrid scheme the hydraulic diameter (d) of the hole is 1mm and impingement plate is provided with holes normal to the plate over its whole length of 250d. While effusion plate has only 20 rows of holes inclined at 30° to its surface. The effect of blowing ratio (BR) over this hybrid scheme of cooling is studied for different BR of 0.5, 1.0, 1.5 and 2.0. It has been found that the area averaged effectiveness increases steeply for BR 0.5 to 1.0 but further increase in BR results only in a small increase. The results also show that increasing the hole diameter increases averaged effectiveness while increasing the center-to-center spacing decreases averaged effectiveness.

Commentary by Dr. Valentin Fuster
2015;():V001T04A005. doi:10.1115/GTINDIA2015-1362.

Pulsating heat pipes (PHP) receives heat from the working fluid distributes itself naturally in the form of liquid–vapor system, i.e., receiving heat from one end and transferring it to other end by a pulsating action of the liquid–vapor system. Pulsating heat pipes have more advantages than other heat pipes. The problem identified is, to calculate the performance of the pulsating heat pipes with respect to different inclinations using various parameters. In this paper, experiment on performance of closed single loop pulsating heat pipe (CLPHP) using water as a working fluid is considered. The parameters such as thermal resistance (Rth), heat transfer coefficient (h), and variation of temperature with respect to time for the given input at different inclinations such as 0°, 45°, and 90° are taken for the present work. Water is used as the working fluid and is subjected to 50% filling ratio and vacuumed at a pressure of 2300Pa. The performance is calculated at different inclinations of the CLPHP with single turn/loop. The factors such as heat transfer coefficient, thermal resistance, time taken for heating the pulsating heat pipe with the given input are calculated. Finally, it has been concluded that is preferable orientation for PHP and it was found be at vertical orientation i.e., at 90° inclination, because more pulsating action is taken place at this inclination and henceforth, heat transfer rate is faster at this inclination.

Topics: Heat pipes
Commentary by Dr. Valentin Fuster
2015;():V001T04A006. doi:10.1115/GTINDIA2015-1377.

The gas turbine combustor liner which is subjected to high temperature requires efficient cooling. In earlier days concept of slot film cooling is utilized in the combustion liners and in modern combustors multiple row film cooling (effusion cooling) is mainly used. This study aims at the experimental investigation of overall film cooling effectiveness of an effusion plate with and without impingement holes at the backside. The experiments are done at different blowing ratios and the surface temperature measurements are taken using infrared thermography. The effusion and impingement holes are arranged in staggered manner on two parallel plates and each effusion hole is surrounded by four impingement holes. Effusion holes are drilled at an angle of 27° and the impingement plate is kept at a distance of 6D away from the effusion plate. The experiments are done on the effusion plate with and without impingement plate at the backside. The results show, increase in cooling effectiveness as the blowing ratio increases. The comparative results shows that at a particular blowing ratio the overall cooling effectiveness is higher for effusion plate with impingement holes at the backside due to the higher convective heat transfer coefficients produced by the impinging jets at the cold side of the effusion plate.

Commentary by Dr. Valentin Fuster
2015;():V001T04A007. doi:10.1115/GTINDIA2015-1392.

Film cooling is often adopted, where coolant jets are ejected to form a protective layer on the surface against the hot combustor gases. The bending of jets in crossflow results in Counter Rotating Vortex Pair (CRVP), which is a cause for high jet lift-off and poor film cooling effectiveness in the near field. There are efforts to mitigate this detrimental effect of CRVP and thus to improve the film cooling performance. In the present study, the effects of both downwash and upwash type of vortex generator on film cooling are numerically analysed. A series of discrete holes on a flat plate with 35° streamwise orientation and connected to a common delivery plenum is used here, where the vortex generators are placed upstream of the holes. The blowing ratio and the density ratio are considered as 0.5 and 1.2 respectively with a Reynolds number based on free-stream velocity and diameter of hole being 15885. The computations are performed by ANSYS Fluent 13.0 using k-ε realizable turbulence model. The results show that vortices generated by downwash vortex generator (DWVG) counteracts the effect of CRVP preventing the jet lift-off, which results in increased effectiveness in streamwise as well as in spanwise directions. However, upwash vortex generator (UWVG) augments the effect of CRVP, resulting in poor performance of film cooling.

Commentary by Dr. Valentin Fuster
2015;():V001T04A008. doi:10.1115/GTINDIA2015-1394.

The experimental investigation of adiabatic film cooling effectiveness is carried out on a flat plates with 4:1 scaled up hole geometries, similar to that of typical turbine nozzle guide vane film cooling holes. Under this study, three flat plate models are considered with the two rows of holes having circular, fan and laidback fan shapes arranged in a staggered manner. These flat plate models are generated using solid works design software and fabricated using low thermal conductivity nylon based material using RPT technique. The mass flow results indicated the average nominal coefficient of discharge for the cooling holes as 0.71, for all these three models based on the inlet hole area and length of the holes. The laterally averaged adiabatic film cooling effectiveness is found along the stream wise direction at a density ratio of 1.62 by varying the blowing ratio in the range of 0.5 to 2.5. The surface temperatures of the test models are captured using the infrared camera, to evaluate the film cooling effectiveness. The experimentally evaluated results shows that, there is no increase in cooling effectiveness for the blowing ratio of 2.0 to 2.5 in the stream wise direction up to the X/d of 25 and there is a marginal increase above the X/d of 25 in the cases of these type of two row circular and Fan shaped hole models. Where as in the Laidback fan shaped hole model, the increase in cooling effectiveness is found significant up to the blowing ratio of 2.5 in the considered range. From the comparative results of adiabatic film cooling effectiveness of these three models, the laidback fan shaped hole model shows the higher film cooling effectiveness than the circular and fan shaped holes model at all the considered blowing ratios.

Commentary by Dr. Valentin Fuster

Structure and Dynamics

2015;():V001T05A001. doi:10.1115/GTINDIA2015-1247.

Flexible supports are used in many aero and automobile industrial applications. They transmit loads, accommodate misalignment, allow axial displacement, ensure no loss of lubricants, absorb shock and dampen vibration, withstand high temperatures, allow easy installation and disassembling. Flexible supports react on connected equipment components when subjected to misalignment and torque. The reaction forces and moments on components due to flexible supports should be within the allowable limits or otherwise it can cause failure of gears, shafts, bearings, and other equipment components.

These flexible supports used in aero engine applications expected to meet design and manufacturing criteria. Flexible supports should have required stiffness values in different directions to meet rotor dynamic stability criteria. Flexible supports also required to meet strength and durability criteria for the given material at the required maximum operating temperature. The designed component should be producible and meet manufacturing limitations.

The main objective of this paper is to optimize single and multiple convolutes types of flexible supports with in the manufacturing limits and in the given design space. A methodology is developed to optimize the components to meet required stiffness, strength and durability criteria. Parametric models of flexible support are developed in UNIGRAPHICS NX9. Design parameters such as overall length, convolute height, convolute radius and angle are considered for the optimization study. ANSYS Workbench is used for the analysis and optimization of flexible support.

Commentary by Dr. Valentin Fuster
2015;():V001T05A002. doi:10.1115/GTINDIA2015-1250.

The aero engine rotating parts are always fracture critical components and their failure in service affects the aircraft safety. Rotors / disks will burst at a certain speed if they operate at ever-increasing speed. Rotor burst is one of important failure mode in aero engine and resulting in disk disintegration into multiple fragments with high speed resulting in containment breach. Disks are subjected to fatigue loading and it limits the service life. Fatigue loading on disk includes high temperature environment, tremendous centrifugal and aerodynamic forces caused by blades.

The main aspect of turbine disk design is to safe guard against LCF failure. Design of disk should ensure that stresses due to thermal, centrifugal and aerodynamics loads during operating conditions should be within the limits. Turbine disks are also designed to operate at speed above 20% of maximum operating speed for maximum power and referred as over speed capability or burst margin. This over speed capability may require for the aircraft during emergency conditions.

The objective of this study is to design a turbine disk for minimum weight. A numerical investigation is performed to predict stresses and burst margins of turbine disk. A parametric disk model is developed with bore width, bore height, web width and web height parameters. Optimization of turbine disk design is carried out to achieve minimum weight. Sensitivity studies are carried out to understand the geometry parameters influence on the stress and burst margins.

Commentary by Dr. Valentin Fuster
2015;():V001T05A003. doi:10.1115/GTINDIA2015-1275.

The influence of deterministic surface texture on the sub-synchronous whirl stability of a rigid rotor has been studied. Non-linear transient stability analysis has been performed to study the stability of a rigid rotor supported on two symmetric journal bearings with a rectangular dimple of large aspect ratio. The surface texture parameters considered are dimple depth to minimum film thickness ratio and the location of the dimple on the bearing surface. Journal bearings of different Length to diameter ratios have been studied. The governing Reynolds equation for finite journal bearings with incompressible fluid has been solved using the Finite Element Method under isothermal conditions. The trajectories of the journal center have been obtained by solving the equations of motion of the journal center by the fourth-order Runge-Kutta method. When the dimple is located in the raising part of the pressure curve the positive rectangular dimple is seen to decrease the stability whereas the negative rectangular dimple is seen to improve the stability of the rigid rotor.

Commentary by Dr. Valentin Fuster
2015;():V001T05A004. doi:10.1115/GTINDIA2015-1291.

This experimental study presents the Lamb wave based baseline-free diagnostics of damage in a carbon-fiber composite fan blade using the Modified Time Reversal Method (MTRM). A fan blade with delamination was selected as the test specimen and two piezoelectric transducers (PZT) were surface bonded using epoxy adhesive. The system (fan blade and transducers) was then waveform-tuned from which 5.5 cycle tone burst was selected as the optimum excitation waveform. The system was then mode-tuned in order to determine the optimum excitation center frequency. The 55 kHz, 5.5 cycle tone burst was used to generate fundamental asymmetric Lamb waves with high signal-to-noise ratio and low dispersion. Damage Index (DI) values for the fan blade was calculated at equidistant points on a line using both PZTs. Results show that magnitudes of the DI values have strong correlation with the severity of damage present in the fan blade. The procedure followed in this experimental study is directly applicable to the Non Destructive Evaluation (NDE) conducted on composite fan blades. Additionally, this method can be extended to develop a real-time structural health monitoring system for composite fan blades which employs embedded PZTs for actuating and sensing waves.

Commentary by Dr. Valentin Fuster
2015;():V001T05A005. doi:10.1115/GTINDIA2015-1297.

Majority of the failures in Gas turbine Blades are caused by High Cycle Fatigue induced by the vibratory stresses in the rotor blades. The first step in blade design is the prevention of coincidence of natural frequencies of the blades with the frequencies of the fluctuating Gas loads.

The forcing frequency is a function of number of upstream and downstream stator blades, and rotational speed. In gas turbines with multiple stages, modal analysis of bladed-disks is individually performed i.e. stage by stage. As the structure is rotationally periodic, cyclic symmetric boundary conditions can be utilized, over 360 degree modeling. The advantage of cyclic symmetry over full model lies in reduced degrees of freedom and hence reduced computational time. In most of the available tools, cyclic symmetry for modal analysis is limited to single stage. As such there is no provision to model and analyze multiple stages at the same time. This leads to inaccurate values of natural frequencies as the flexibility introduced by the adjacent stages is not being taken into consideration. An alternative to this is full 3D modeling and analysis of all the combined stages.

Bladh et al. (2003) [1] have shown that interstage coupling can significantly affect the dynamics of the multi-stage assembly and in some cases lead to an underestimation of vibratory levels. Sokolowski et al [2] studied the influence of inclusion of shaft in the model on the natural frequencies and mode shapes of the shrouded bladed discs up to four nodal diameters for first two frequency series (mode shapes). Rzadkowski and Drewczynski (2006) [3] have used full 360 degrees models to study the free and forced dynamics of multi-stage systems. However this method is avoided as the computational cost is prohibitive.

Multi stage cyclic symmetry overcomes this obstacle in which each stage is cyclically modeled and an inter-stage coupling is introduced between adjacent stages. The advantage of multi stage cyclic symmetry lies in the significant reduction in the number of elements and therefore computational time. Laxalde et al. (2007) [4] were the first to come up with the method of dynamic analysis of turbo machinery rotors with multi stage cyclic symmetry using interstage coupling. They considered an example of two-stage High Pressure compressor. The results were validated against a complete 360 degrees reference model. Forced response analysis of rotor stages to fluctuating gas loads with and without interstage coupling definition was also presented and compared. In the present work a complete Gas Turbine rotor system with multiple stages of Compressor, Shaft and Turbine were analyzed together.

Commentary by Dr. Valentin Fuster
2015;():V001T05A006. doi:10.1115/GTINDIA2015-1298.

Complex stress strain response of a turbine rotor used in a gas turbine engine was studied. Simple and comprehensive approximation techniques developed by Muralidharan–Manson, Bäumel-Seeger (from data obtained from tension tests) and Roessle–Fatemi (from data obtained from hardness tests) were used to predict the fatigue constants of the rotor material. Multiaxial Fatigue damage models like von Mises equivalent strain model, Smith Watson Topper model, Fatemi–Socie Model, Kandil Brown and Miller model were used to predict the fatigue life of the rotor. Predictions were then compared with the life obtained from the same damage models using the experimental fatigue constants and the life obtained from Low Cycle Fatigue (LCF) testing of the turbine rotor. Acceptable life predictions were obtained with SWT model and FS model using the fatigue constants obtained from the experiment as well as from the approximation techniques. von-Mises equivalent strain model failed to give reasonable life predictions with fatigue constants obtained from the experiment and approximation techniques. The life predicted by KBM model using fatigue constants obtained from approximation techniques (Bäumel-Seeger and Roessle-Fatemi) was found unsatisfactory. The approximation technique proposed by Muralidharan-Manson in combination with all the damage models fitted the failure data within a factor of 5. Finite Element tools were used to determine the stress/strain response of the component under the mutiaxial loading condition.

Commentary by Dr. Valentin Fuster
2015;():V001T05A007. doi:10.1115/GTINDIA2015-1309.

The fatigue life of the titanium alloy centrifugal impeller is estimated based on strain life method. a) Equivalent Strain Method and b) Smith Watson Topper Method are used. The probabilistic analysis of fatigue life is carried out by Weibull distribution method. The fatigue life of the impeller is assessed through cyclic spin test. The results are compared. The Equivalent Strain Method is found non conservative. The average fatigue life of the impeller is estimated as 12000 cycles by SWT method, which is found more closer to test results. The reliability improvement from 95% to 99% reduces the fatigue life around 18%.

Commentary by Dr. Valentin Fuster
2015;():V001T05A008. doi:10.1115/GTINDIA2015-1312.

This paper is concerned with the coupled thermo-mechanical stress analysis of functionally graded (FG) gas turbine rotor shaft system. Gas turbine shaft may expose in high temperature environments which demands to use functionally graded materials (FGMs). The aim of the present work is to study the stresses developed in the FG turbine shaft due to temperature variations and mechanical loading due to unbalance masses. For the present analysis aluminum oxide (Al2O3) and stainless steel (SUS304) are taken as shaft materials, power law gradation is used for the determination of FG material properties of the turbine shaft. Three nodded Timoshenko beam element with six degree of freedom (DOF) per node is considered for the finite element modelling of FG shaft. First order shear deformation theory (FSDT) is used with rotary inertia, strain and kinetic energy. Solution for governing equation of motion is obtained by the Hamilton principle. Complete MATLAB code has been developed for thermosmechanical stress analysis. Comparative study between steel shaft and FG shaft have been carried out. Normal stress (σxx) on plane perpendicular to axial direction, shear stress (τxr) on plane perpendicular to axial direction in radial direction and shear stress (τ) on plane perpendicular to axial direction in circumferential direction are obtained against time and along radius of shaft. Also these stresses are obtained for different parameters like power law indexes and speed of rotation of shaft.

Commentary by Dr. Valentin Fuster
2015;():V001T05A009. doi:10.1115/GTINDIA2015-1325.

In great majority of situations, bolted flanges are designed to work well within elastic limits. The main design considerations concern preload which should be high enough to limit fatigue loading of the bolt and also prevent flange opening in the operating range. Elastic design of bolted joints as well their finite element simulation have been well understood.

The present paper deals with post elastic behavior of bolted joints. Equipment and structures are often exposed to loads much higher than the normal operating loads. Civil engineering structures experiencing earthquakes is one such example. Military ships and submarines are subject to torpedo loads and other types of blast loads. Aircraft engines are subject to ultimate load conditions such as fan blade-off (FBO) and foreign object damage (FOD) like bird-hit. To meet such eventualities it is uneconomical to attempt elastic designs. The approach is to go for plastic designs. The criterion is that the structure or equipment can yield and distort but should not rupture. The philosophy is that the distorted parts are replaced once the event is over. Bolted joints are invariably present in these categories of equipment.

The present paper deals with simulation and structural behavior of components fastened together by way of threaded fasteners. Sector of bolted flange is considered for study and elastic-plastic analyses are carried out. This is an extension of the work carried out earlier by the authors for simple axisymmetric joints. The earlier study was for conceptual understanding. In the present study focus will be on design aspects. Three different simulation models are compared. In addition, parametric studies are conducted to get deeper insight into structural behavior.

Commentary by Dr. Valentin Fuster
2015;():V001T05A010. doi:10.1115/GTINDIA2015-1330.

Nowadays, heavy and bulky rotors are replaced by the light yet strong rotor, where the composite material is only supplementary. The composite may be constructed either by reinforcing long unidirectional fiber into matrix material or stacking of lamina, where each lamina has different orientation of fiber. But mathematical modelling of such type of rotor is little difficult when considering different orientation of fiber. This invokes us to construct multilayer rotors of different isotropic material and associated formulation to show its better dynamic performance.

Generally internal damping has an enormous effect on the dynamics of rotor shaft system. For the sake of modelling, all layers are assumed to made of viscoelastic material and perfectly bonded. The constitutive relationship of each layer is represented by two element voigt model and equation of motion is obtained in time domain.

This paper involves the development of both classical and finite element mathematical model of multilayer viscoelastic rotors, which contents system characteristics. Under these conditions, the complex modal behaviour of the rotor-shaft is studied to get an insight of the dynamic characteristics of the system, in terms of Decay rate, Stability Limit of Spin-speed, First Natural Frequency and also Unbalance frequency response.

Commentary by Dr. Valentin Fuster
2015;():V001T05A011. doi:10.1115/GTINDIA2015-1366.

The present work emphasizes the dynamic response of double cracked cantilever beam subjected to a traversing mass. The cracks are located at different positions of the beam with random crack depths. The response of the damaged structure has been evaluated employing a numerical procedure of Runge-Kuuta method. The effects of crack depth, traversing mass, traversing speed and crack location on the response of the structure are studied. Finite element analysis (FEA) using the commercial ANSYS 15 has been presented to validate the adopted numerical method.

Commentary by Dr. Valentin Fuster
2015;():V001T05A012. doi:10.1115/GTINDIA2015-1390.

Forced vibration analysis has been carried out on functionally graded plates where the material properties vary along axial direction. The geometric nonlinearity is incorporated in the system using nonlinear strain displacement relations. An indirect methodology is adopted in which the dynamic system is assumed to satisfy the force equilibrium condition at peak excitation amplitude, thus reducing the problem to an equivalent static case. The computational points are selected and start functions are generated at those points by satisfying the flexural and membrane boundary conditions of the plate. The start functions are later used for generating higher order functions using Gram-Schmidt orthogonalisation procedure. The mathematical formulation is based on the variational form of energy principles and the governing equations are derived using Hamilton’s principle. The set of nonlinear governing equations is solved using an iterative direct substitution method employing an appropriate relaxation technique. The results are generated for combinations of clamped and simply supported boundary conditions and presented in amplitude-frequency plane. Three dimensional operational deflection shape plots along with contour plots are also provided for some cases. Results are validated with the works available in the literature.

Commentary by Dr. Valentin Fuster

Controls, Diagnostics and Instrumentation

2015;():V001T06A001. doi:10.1115/GTINDIA2015-1248.

This paper deals with the development of multi-parametric Model Predictive Control (mp-MPC) strategies for a laboratory SR-30 gas turbine setup. The objective is to control the engine speed and turbine exhaust pressure of the gas turbine. Firstly, an empirical transfer function model is obtained experimentally, between the fuel flow-shaft speed and nozzle diameter-turbine exhaust pressure (TEP) of the gas turbine. Then, Model Predictive Control (MPC) is designed based on the empirical models. The output responses under MPC are found to be satisfactory, however, with large computational time. Next, mp-MPC controllers are designed based on the empirical models. Relevant operating constraints are also added as design specifications. It is found that the multi-parametric MPC delivers superior results in comparison with conventional MPC, in terms of computational time, while delivering the same transient performance.

Commentary by Dr. Valentin Fuster
2015;():V001T06A002. doi:10.1115/GTINDIA2015-1264.

Fractional-order modeling and controller design by a simplified way is the demanding research area and is gearing more and more momentum. This paper is the attempt of application of fractional-order modeling and controller design for the power plant gas turbine. The Gas Turbine is most important equipment in power, aviation and automotive industry. It converts the thermal energy of fuel into the mechanical power. Therefore, important requirement of gas turbine system is to control the flow of input fuel. The existing identified model of the gas turbine between the input fuel flow and the output speed, is of high-order and integer type, which is reduced to the simple and compact integer-order (IO) and fractional-order (FO) models using local optimization technique. The fractional-order internal model controller (FO-IMC) is designed and to show the performance efficacy it is compared with integer-order internal model controller (IO-IMC), which is also designed using the same methodology and specification. Simulation results show that FO-IMC based controller gives better performance for the set point tracking, plant uncertainty and disturbance rejection than the IO-IMC. FO-IMC controller also satisfy the robust stability condition.

Commentary by Dr. Valentin Fuster
2015;():V001T06A003. doi:10.1115/GTINDIA2015-1344.

In house development of lubrication oil pump of gerotor type used in gas turbine engine is described in this paper. A host of geometrical parameters determine the inner and outer rotor profiles which in turn define the flow characteristics of the pump. In this paper, a simulation model has been developed based on AMESim and MatLab which used to predict the flow characteristics of a pump. This model incorporates both the methodologies of design and parametric analysis of the pump which aids to design the pump in accordance with designer needs by varying the parameters. A gerotor pump is designed having fixed geometrical parameters using this model. The prototype pump has been tested for its flow characteristics and compared with estimated result. The comparison indicates that simulation results agree well with the measured data. Thus, this simulation model will be useful in designing and analyzing the lubrication pump of a gas turbine engine.

Commentary by Dr. Valentin Fuster

Manufacturing, Materials and Metallurgy

2015;():V001T07A001. doi:10.1115/GTINDIA2015-1352.

LD slag (LDS) is a major solid waste generated in huge quantities during steel making. It comes from slag formers such as burned lime/dolomite and from oxidising of silica, iron etc., while refining the iron into steel in the LD furnace. This work aims at utilization of waste LDS to develop surface coatings by plasma spraying technique. This technology has the advantage of being able to process various low-grade ore minerals to obtain value-added products and also to deposit materials, generating near homogeneous coatings with the desired microstructure. Coatings prepared for this investigation are characterized in terms of their thickness, hardness, adhesion strength and porosity. Coating deposition efficiency is calculated in order to assess the coatability of LDS and XRD is carried out in order to ascertain the various phases present in the coating. Premixing of TiO2 powder with LDS is found to substantially improve the interfacial adhesion. It is also found that the operating power levels of the plasma torch affect the adhesion strength, coating deposition efficiency and mean thickness of the coatings. This work reveals that LD slag is eminently coatable and can be gainfully used as a potential cost-effective material for deposition of plasma spray coatings on metallic substrates.

Commentary by Dr. Valentin Fuster

GT Operation and Maintenance

2015;():V001T08A001. doi:10.1115/GTINDIA2015-1240.

Gas turbine based combined cycle power plants are found to play vital role in electric power generation in recent years. Modern gas turbines are integral elements in these plants operating at very high temperature with high efficiency. Improvements in plant reliability, availability and maintainability (RAM) have been major areas of concern for power producers to ensure competitive positions. In gas turbine based power generation systems, the performance of the turbine drops due to several reasons like compressor fouling and inlet filter clogging. For improving plant RAM, advanced methods of health monitoring are vital for gas turbine plant components such as inlet air filter, compressor, combustion chamber and turbine. This paper focuses on health monitoring of gas turbine compressor considering major fault condition of compressor fouling. The health monitoring is achieved by developing a compressor model, to predict the performance of the compressor at design and off-design operational conditions. A thermodynamic model of the gas turbine system has limited applicability for health monitoring applications. The modelling framework has to incorporate the complex assembly of various components that make up the overall system and the real time off-design operations of the system. With the recent developments in computational methods and availability of vast computing power, process history data models are found to be convenient options for the system modelling. Among the process history data based methods, Artificial Neural Networks (ANN) have proved to be effective for modelling non-linear and complex processes. Hence, ANN is used as the modelling platform for this study and the model is developed from process data of a GE frame 9E machine. Residuals generated from the model are used for analysing the health of the system. The prediction of future events achieved through the model is found to provide vital information for the decision making and planning of maintenance actions. Principal Component Analysis (PCA) is a suitable method that is efficient in accounting the variability of the data. It derives the loading vectors and is suitable for improving the effectiveness of the ANN model. Different fault conditions relevant to the gas turbine compressor are demonstrated with actual plant data using the ANN based health monitoring system. Effect of compressor fouling and recouping of fouling effect with off line compressor water wash are also analysed in this paper.

Commentary by Dr. Valentin Fuster
2015;():V001T08A002. doi:10.1115/GTINDIA2015-1273.

The purpose of this research was the gas-dynamic design of the pneumatic brake system on the basis of the existing geometry of the three stage axial flow low-pressure compressor for application in the GTE test benches.

The pneumatic brake system has been obtained by adding of the add stage to the low-pressure compressor, the geometry of which repeats the geometry of the stages of the base compressor. In this work, two options were considered. First, the second stage of the base compressor was used as the add stage. Second, the third stage of the base compressor was used as the add stage.

Commentary by Dr. Valentin Fuster

CCPP, Heat Recovery Steam Generators and Steam Turbines

2015;():V001T09A001. doi:10.1115/GTINDIA2015-1261.

Market demands such as generating power at lower cost, increasing reliability, providing fuel flexibility, increasing efficiency and reducing emissions have renewed the interest in Integrated Gasification Combined Cycle (IGCC) plants in the Indian refinery segment. This technology typically uses coal or petroleum coke (petcoke) gasification and gas turbine based combined cycle systems as it offers potential advantages in reducing emissions and producing low cost electricity. Gasification of coal typically produces syngas which is a mixture of Hydrogen (H) and Carbon Monoxide (CO). Present state of gas turbine technology facilitates burning of low calorific fuels such as syngas and gas turbine is the heart of power block in IGCC. Selecting a suitable gas turbine for syngas fired power plant application and optimization in integration can offer the purchaser savings in initial cost by avoiding oversizing as well as reduction in operating cost through better efficiency.

This paper discusses the following aspects of syngas turbine IGCC power plant:

• Considerations in design and engineering approach

• Review of technologies in syngas fired gas turbines

• Design differences of syngas turbines with respect to natural gas fired turbines

• Gas turbine integration with gasifier, associated syngas system design and materials

• Syngas safety, HAZOP and Hazardous area classification

• Retrofitting of existing gas turbines suitable for syngas firing

• Project execution and coordination at various phases of a project

This paper is based on the experience gained in the recently executed syngas fired gas turbine based captive power plant and IGCC plant. This experience would be useful for gas turbine technology selection, integration of gas turbine in to IGCC, estimating engineering efforts, cost savings, cycle time reduction, retrofits and lowering future syngas based power plant project risks.

Commentary by Dr. Valentin Fuster

GT Cycle Innovations, Renewable Applications

2015;():V001T11A001. doi:10.1115/GTINDIA2015-1230.

The need of an efficient, low cost and environment friendly hydraulic energy converter provides motivation to many researchers to contribute in the field of renewable energy. In the present investigation, one such requirement is addressed for the possibility of electrical power generation from free stream of water. The implementation of such a low head or zero head turbine does not require a dam for energy conversion, thereby making it a low cost and environmental friendly source of power generation. The present study deals with the development of a zero head vertical-axis helical water turbine, and its subsequent testing in an open channel. The main parameters that influence the performance of a helical water turbine are its blade profile, aspect ratio, helix angle, number of blades and solidity ratio. Considering these parameters and the numerical work reported in literature, a three-bladed helical turbine has been developed and tested under different loading conditions. The variation of power coefficient at various tip-speed ratios of the turbine has been investigated and analyzed.

Commentary by Dr. Valentin Fuster
2015;():V001T11A002. doi:10.1115/GTINDIA2015-1263.

In the recent times, there has been a much concerned about the efficient utilization and conservation of energy across the globe. The fossil fuels supply most of the energy requirements, however, it is a non-renewable energy source having a limited reserve across the globe. In view of this, there has been a continuous drive to explore the renewable, easily available and environment friendly energy sources to partially or fully replace the fossil fuel. With this background, this paper investigates the viability of utilizing producer gas, derived from the biomass briquettes, as the primary fuel to run a diesel engine in dual fuel mode. Biomass used was sun-dried pine leaves and cow dung. Briquettes were prepared at the proportion of 75 % cow dung and 25 % pine leaves, by mass, using water as the binder. The producer gas, generated from a downdraft gasifier, was fed to single cylinder, four-stroke, water-cooled, 5.2 kW compression ignition engine to run on dual fuel mode. A minor modification of engine was carried out to suit the dual fuel mode operation. The performance and emission characteristics of the engine were studied at various loads. The system sustains well in a dual fuel mode although there is a drop in brake thermal efficiency in the range of 7 to 22%. There is a significant reduction, up to 93%, in NOX emission in the exhaust but CO and CO2 emission increases, more than 40% and 8% respectively, in the dual fuel mode.

Commentary by Dr. Valentin Fuster
2015;():V001T11A003. doi:10.1115/GTINDIA2015-1266.

With the rising fossil-fuel prices, energy scarcity and climate-change, renewable energy plays an important role in producing local, clean and inexhaustible energy source to supply world rising demand for electricity. The selection of suitable wind turbine plays a vital role for urban power generation where wind is characterised by unsteadiness and turbulence. Thus, blade aerodynamics of wind turbine has a significant effect on turbine efficiency.

In this study, the aerodynamic aspect of a straight bladed Darrieus turbine is numerically analyzed. Two dimensional numerical modelling and simulation of unsteady flow through the rotor blades (NACA 0018) of the turbine is performed using ANSYS FLUENT 14.5. The unsteady Reynolds averaged Navier-Stokes (RANS) equation is used to demonstrate the effects on the performance of two dimensional Darrieus turbine blade. The Shear Stress Transport (SST) k-ω model has been adopted for the turbulence closure. For the proposed analysis, the flow field characteristics are investigated at different azimuthal angle and tip speed ratio. Further, the parametric quantities such as solidity, number of blades and blade thickness have being investigated for a uniform free stream velocity of 6 m/s. The effect of laminar boundary layer separation on performance of the Darrieus turbine has also been taken into account during the study of flow physics around the blade. The results obtained are compared with the reported experimental and computational data.

Commentary by Dr. Valentin Fuster
2015;():V001T11A004. doi:10.1115/GTINDIA2015-1329.

The rotor is the first element in the chain of functional elements of a wind turbine. Therefore, its aerodynamic and dynamic properties have a decisive influence on the entire system in many respects. The capability of the rotor to convert a maximum proportion to the wind energy, flowing through its swept area, into mechanical energy is obviously the direct result of its aerodynamic properties. These will determine the overall efficiency of the energy conversion in the wind turbine. To enhance the performance of wind turbine, one has to analyze the effects of variations in nomenclature of a rotor blade on the performance of wind turbine.

The present paper incorporates the analysis of primary selected blade design by the Solar Energy Research Institute (SERI -8) with BEMT based simulations; five different suggested models of the rotor blade are also analyzed with same BEM based simulations, including Tip losses, 3-D corrections and Reynolds number drag correction. These six models incorporate twist optimization, chord distribution optimization, and the combination of airfoils, new designed airfoils and base model SERI-8 rotor blade. Without modifying the size of the rotor blade, different analysis is done and improved the rotor aerodynamics efficiency. Highly aerodynamics efficient new designed airfoils produce more lift and minimize the drag by avoiding the stall; it develops high-value of coefficient of lift to drag (Cl/Cd) ratio. The rated power for all models considered at 15m/s for analysis. The result suggests that the rotor power efficiency is 25.772% with SERI-8, 29.420% with the twist optimized model, 55.51% with the model with a combination of SG series and SERI series airfoils and 61.62% with new designed airfoils (SV Series) model. Respectively, efficiency is increased by 3.64%, 29.745%, and 35.84% compared to SERI-8 blade.

Commentary by Dr. Valentin Fuster

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