Second Law Efficiency of an Intercooled-Reheated-Regenerative-Brayton Cycle With Variable Temperature Heat Reservoirs

The gas turbine engine works on the principle of the Brayton Cycle. One of the ways to improve the efficiency of the gas turbine is to make changes in the Brayton Cycle. In the present study, Brayton Cycle with intercooling, reheating and regeneration with variable temperature heat reservoirs is considered. Instead of the usual thermodynamic efficiency, the Second law efficiency, defined on the basis of lost work, has been taken as a parameter to study the deviation of the irreversible Brayton Cycle from the ideal cycle. The Second law efficiency of the Brayton Cycle has been found as a function of reheat and intercooling pressure ratios, total pressure ratio, intercooler, regenerator and reheater effectiveness, hot and cold side heat exchanger effectiveness, turbine and compressor efficiency and heating capacities of the heating fluid, the cooling fluid and the working fluid (air). The variation of the Second law efficiency with all these parameters has been presented. From the results, it can be seen that the Second law efficiency first increases and then decreases with increase in intercooling pressure ratio and increases with increase in reheating pressure ratio. The results show that the Second law efficiency is a very good indicator of the amount of irreversibility of the cycle.

A Strategy for Multi-Point Compressor Blading Using Multi-Objective Optimization

This paper introduces a framework to perform a multi-objective multipoint aerodynamic optimization for an axial compressor blade. This framework considers through-flow design requirements and mechanical and manufacturing constraints. Typically, components of a blade design system include geometry generation tools, optimization algorithms, flow solvers, and objective functions. In particular, optimization algorithms and objective functions are tuned to reduce blade design calculation cost and to match designed blade performance to the through flow design criteria and mechanical and manufacturing constrains. In the present study, geometry parameters of blade are classified to three categories. For each category, a distinct optimization loop is applied. In outer loop, Gradient-based optimization techniques are used to optimize parameters of the second category and a two-dimensional compressible viscous flow code is used to simulate the cascade fluid flow. Surface curvature optimization is carried out in inner loop, and its objective function is defined by integrating the normalized curvature and curvature slope. The genetic algorithm is used to optimize the parameters in the interior loop. To highlight the capabilities of the design method and to develop design know-how, an initial profile is optimized with three different design philosophies. The highest performance improvement in the first case is 15% reduction in loss at design incidence angle. In the second case, 16.5% increase in allowable incidence angle range, improves blade’s performance at off design conditions.

Simulations of 3D Separation in the Diffuser

Large eddy simulations of an asymmetric diffuser characterized by complex 3-D flow separation for which RANS models provide qualitatively wrong predictions have been performed.. An incompressible, turbulent, fully-developed flow in a rectangular duct (aspect ratio 1:3.33) expands into this diffuser. Two such diffusers were constructed by deflecting a pair of adjacent walls for the experiments in Cherry et al. [1, 2] (2006, 2008) and Buice, C. U. and Eaton, J. K. [3]. Most of our simulations consider Diffuser 1 with wall deflection angles 11.3° and 2.56°. In the experiments, flow begins to separate at the corner formed by the two deflected walls and then spreads so that flow is separated from the wall at the larger deflection angle. In simulations with RANS models, flow separates from the wall with the smaller deflection. It has been possible to obtain solutions with LES where flow separates correctly, off the wall at the larger deflection angle, as in the experiment. The LES finds a qualitatively correct separation, with characteristics in close quantitative agreement (within 5%) with the experimental values for Diffuser 1. The effects of variations in grid aspect ratio, grid refinement, inlet length, number of flow passes, and secondary flow structure upstream of the diffuser on solutions were determined. An LES was carried out for Diffuser 2 (deflection angles of 9° and 4° respectively), applying all lessons learnt in Diffuser 1 studies. It was found that the results for Diffuser 2 are not as quantitatively close to the experimental results as in case of the Diffuser 1, but the discrepancies appear to have a similar origin in some finer aspect of diffuser inflow conditions.

Aerodynamic Measurements on the Interaction of Secondary Jets and Separation Bubble

The dynamics of separation bubble under the influence of continuous jets ejected near the semi-circular leading edge of a flat plate is presented. Two different streamwise injection angles 30° and 60° and velocity ratios 0.5 and 1 for Re = 25000 and 55000 (based on the leading-edge diameter) are considered here. The flow visualizations illustrating jet and separated layer interactions have been carried out with PIV. The objective of this study is to understand the mutual interactions of separation bubble and the injected jets. It is observed that flow separates at the blending point of semi-circular arc and flat plate. The separated shear layer is laminar up to 20% of separation length after which perturbations are amplified and grows in the second-half of the bubble leading to breakdown and reattachment. Blowing has significantly affected the bubble length and thus, turbulence generation. Instantaneous flow visualizations supports the unsteadiness and development of three-dimensional motions leading to formation of Kelvin-Helmholtz rolls and shedding of large-scale vortices due to jet and bubble interactions. In turn, it has been seen that both the spanwise and streamwise dilution of injected air is highly influenced by the separation bubble.

Investigations Into Creep Behavior of Gas Turbine Component Assemblies

Modern day gas turbines are very complex in construction and consist of a very large number of smaller parts and subassemblies. Hence the most vital parts of the entire assembly are the mechanical devices which are deployed to connect and keep them together. In gas turbines two approaches are normally used in the assembly process. They are the threaded fasteners such as bolt and nut and shrunk-fit or interference-fit assemblies. In the high temperature regions of the gas turbines the effect of creep on the integrity of such fastening arrangements needs to be assessed at the design stage. A problem commonly faced pertains to lack of creep data which would facilitate detailed nonlinear analysis. The available data invariably exhibit scatter. In this paper parametric studies are undertaken. Creep curves are chosen so that both primary and secondary stages are accounted for. The coefficients are chosen to meet the design needs. The performance of bolted joints and shrunk-fit assemblies get affected over time due to stress relaxation leading to loss of bolt pretension or the effective interference. The bolt preload as well as the interference is to be optimally chosen. Higher the preload or the interference the more effective is the joint. At the same time the stress levels are higher and hence the stresses will relax to a greater extent. For a design stage assessment of the behavior of assemblies there is need for correlation among the various operating parameters such as stress, temperature and time. For individual components one normally uses empirical correlations such as Larson-Miller to predict rupture life and also creep growth. For assemblies in which relaxation is the main design issue, such parameters are usually not available. There is need to carry out detailed nonlinear analysis. Typical bolted flange and shrink-fit assemblies are chosen for study. Parametric studies are carried out. Using creep properties as described earlier, nonlinear structural responses are studied. The purpose is to correlate the creep properties, in terms of creep strain with respect to time, stress and temperature, with the joint behavior. The key joint behavior indices are the bolt tensile stress in the case of threaded fastening and the compressive force of “effective interference” in the case of shrunk-fit assemblies. The studies have established the need for rigorous creep analysis of components having interference fits or threaded fasteners. Once the operational requirements are known, the approach presented helps in material selection.

Effective Radiation Modelling Technique for Transient Temperature Prediction of Gas Turbine Components

Gas Turbine engine components like Combustor, Diffuser, and Turbines are subjected to very high temperatures. Predicting accurate temperatures of such components demand accurate Radiation modeling along with Conduction and Convection. Radiation heat transfer modeling is very complex due to non linear dependence on temperature and additional parameters driving the heat transfer like shape factor, emissivity, surface area and absorbtivity of material. The commercial software ANSYS developed various Radiation techniques like ‘Radiation Matrix’, ‘Radiosity’ and ‘Radiation modeling between a surface and a point’. A detailed study has been carried out to compare different Radiation models. The ease of building the model, computational time, accuracy, and limitations are thoroughly examined. It is found that all existing methods have some limitations in accuracy, computational time or system requirements. To overcome some of these limitations, a new technique called ‘Surface Effect Element Method’ is proposed in this paper. This method uses ‘Radiosity’ for the shape factor computation and ‘Radiation modeling between a surface and a point’ for modeling Radiation between two surfaces. The average of one surface temperature is transferred to a single point which in turn is used to model the Radiation to the second surface and the same procedure is repeated for the second surface too. A detailed study is carried out and the proposed technique is compared against the available methods. The new technique enables accurate computation of transient temperatures for gas turbine components leading to accurate life prediction for these components. It is shown that ‘Surface Effect Element Method’ has comparable accuracy but significantly lower cycle time and efforts compared to existing methods.

The Effect of Reynolds Number on Transonic Compressor Blade Rotor Section

To achieve at a more precise designing procedure in axial-compressors as well as a higher pressure ratio value, a comprehensive understanding on the flow aerodynamics and the governing phenomena is required. Existence of these complicated phenomena e.g., simultaneous production of supersonic and subsonic flows, shock-boundary layer interaction, unique incidence phenomenon, etc, makes it difficult to analyze the flow in the transonic compressors. One of the methods which is useful in the modeling of the phenomena occur in the compressors is investigating the flow in the blade to blade passage. In this paper, employing the simultaneous solution of the full Navier-Stokes equations (using the Roe-FDS numerical method) and turbulence equations (using the K–w (SST) model) the flow has been simulated in the blade to blade passage of a transonic compressor. In the following, in order to comparison the predicted results with experimental data, required adjustments and conditions have been taken into account. After passing through the first transonic compressor stages, the flow becomes remarkably compressed. In such conditions, the Reynolds number considerably changes compared to the inflow Reynolds number. In the present work, it is intended to numerically investigate the effects of the inflow Reynolds number on the unique incidence, flow losses, deviation angle, and also shock position changes, in three different important states of “Minimum loss” and “Choked flow” in started conditions and “Stall operation” in unstarted conditions.

Numerical Simulation of Effects of Tip Geometry on the Performance of an Axial Compressor

The purpose of this paper is to investigate numerically the effects of the tip geometry on the performance of an axial compressor rotor. There are three case studies which are compared with the base line tip geometry. 1) baseline (flat tip); 2) Cavity (tip with a cavity); 3) SSQA (suction side squealer tip) and 4) SSQB (modified suction side squealer tip). The case of SSQB is a combination of suction side squealer tip and the cavity tip. From leading edge to 10% chord, the tip has a cavity. From 10% chord to trailing edge, the tip has a suction side squealer. The numerical results of 2) show that the cavity tip leads to lower leakage mass flow and greater loss in tip gap and the rotor passage. The loading near the blade tip is lower than the baseline, thus the tangential force of the blade is lower. It leads to lower pressure rise than the baseline. The performance of the compressor for the tip with cavity is worse than the baseline. The results of 3) show that the higher curvature of the suction side squealer increases the loading of the blade and the tangential blade force. With the suction side squealer tip, the leakage flow experiences two vena contractor thus the mass of the leakage flow is reduced which is benefit for the performance of the compressor. The loss in the tip gap is lower than baseline. The performance is better than the baseline with greater pressure rise of the rotor, smaller leakage mass flow and lower averaged loss. For the case the SSQB, the leakage mass flow is lower than the SSQA and the loss in the tip gap and the rotor passage is greater than SSQA. The performance of the case of the SSQB is worse than the case of SSQA.

Non-Axisymmetric Flow Structure in Vaneless Diffuser of Centrifugal Compressor for Turbochargers

Turbocharger technology is widely used in internal combustion engines. With the downsizing of internal combustion engines and the introduction of strict emission regulations, there is urgent demand for turbochargers featuring centrifugal compressors with a wide flow range. The flow in a centrifugal compressor of a turbocharger is non-axisymmetric due to the inherent asymmetry of the discharge volute. The asymmetric flow field inside the diffuser has great influence on the performance of centrifugal compressor. In order to develop a flow control method that facilitates a wider flow range of turbocharger compressors, further understanding of the asymmetric flow structure is very important. The main subject of this study is to reveal the asymmetrical characteristics of the flow field in the vaneless diffuser of a centrifugal compressor followed by a volute. Oil flow visualizations and numerical simulations were used. The results of the numerical simulations are consistent with that of the oil flow visualizations near choke and at designed flow rate. The results show that a “dual-zone mode” asymmetric flow structure exists near the shroud of the vaneless diffuser at near choke condition. A bifurcation point at the volute tongue that divides the flow and creates two distinct flow patterns was found. The asymmetry of the flow structure near the hub was much less significant than that near the shroud. At the design flow rate, asymmetric flow patterns are found neither near shroud nor near hub. At near surge condition, the pattern of the oil flow traces near the shroud is very different from those near choke.

Effect of Erodent Particle Initial Velocity on the Erosion of Propeller Blades for Turboprop Engines

Propeller driven-engines operate efficiently at low speeds, and ground maneuvers, but its performance is affected by operating in unsuitable environment. Actually, it is susceptible to encounter many physical problems such as erosion, corrosion, foreign object damage, and icing. These problems not only cause changes in air path boundaries but also yield changes in the aerodynamic performance of the engine components due to the change of the propeller profile shape and increase in the overall surface roughness. This work aims to study the effect of the particle initial velocity on the propeller erosion phenomena and the subsequent deterioration for the blades profile. Particle trajectory, erosion rate, frequency and the critical erosion area on the blade are the main issues under investigation. The domain selected for computational study is a periodic sector through the propeller bounding and the boundary conditions are set corresponding to that exist in the propeller manuals. A three dimensional unstructured grid was generated and adopted using commercial turbomachinery grid generator GAMBIT software. The governing equations are solved using FLUENT6.3.26 a commercial CFD code, which uses a control volume approach on a grid over the computational domain. A Lagrangian-formulated particle equation of motion is added to predict particle velocity and trajectories once the air flow field is obtained.