Aeroelastic Flutter Investigation and Stability Enhancement of a Transonic Axial Compressor Rotor Using Casing Treatment

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

Mean-Line Modelling of a Variable Geometry Turbocharger (VGT) and Prediction of the Engine-Turbocharger Coupled Performance

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

Design and Analysis of a Marine Current Turbine

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

Experimental and Numerical Studies on a Curved Back Transonic Airfoil

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

CFD Studies on the Performance of a Centrifugal Compressor With Single Wall Rotating Vaneless Diffusers at the Wall Extension Ratios of 1.1 and 1.15

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

Experimental Investigation of Low Emission Liquid Fuelled Reverse Cross Flow Combustor

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

Flow Studies on a Single Stage Transonic Axial Flow Compressor Retrofitted With Circumferential Grooves and Varied Rotor-Stator Axial Gap

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

Impact of Inlet Fogging on the Performance of Steam Injected Cooled Gas Turbine Based Combined Cycle Power Plant

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

Development of LP Blade Module for High Back Pressure-Aerodynamic Design

In steam turbine power plants, the appropriate design of the last stage blades is critical in determining the plant efficiency and reliability. The development of LP module for desert applications is finding applications for a number of industrial steam turbine operating with air cooled condensers. The conventional LP Module for water cooled condenser operates at low back pressure (Pexit = 0.09 bar) and are generally not suitable for high back pressure application. This paper focuses on the aerodynamic design & optimization of last stages of LP blade module for high back pressure application and validation through 3D CFD. The guide and moving blade are designed with seven equally-spaced profiles section from hub to shroud through Axstream S/w. The profile and incidence losses are minimized for the design and off-design conditions. Aeromechanical design of LP blade module consisting of 2 stages for 0.2 bar back pressure, 1.1 bar inlet static pressure and a mass flow of 61.2 kg/s is carried out. An optimization process through a streamline curvature code and design optimization software using Optimus is established and flow path contours is optimized thoroughly, a total to total efficiency of 81.4% is achieved for the rated condition. The off-design performance is investigated for a wide range of operating conditions, especially at low volume flow rate of steam condition.

Impact of Geometric Parameters of Fuel-Air Distribution System on the Temperature Variation and Emission of a Sideway Faced Porous Radiant Burner (SFPRB)

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