Renewable energy generation in the rural environment has been receiving an increased attention over the recent years due to the proximity with the point of use. This paper presents the effect of design parameters on the performance of a Shutter Type Vertical Axis Wind Turbine (STVAWT). A STVAWT has been designed, manufactured and tested. The turbine performance has been investigated by varying the design parameters such as shutter angle and form of the shutter. The results were used for the comparison between the performance achieved while changing the design parameters. Significant numbers of experiments have been performed by changing the above mentioned parameters for different wind speeds. The effect of each parameter on the torque and power has been analyzed. It has been found that the shutter angle has a significant effect on the power of the turbine. The maximum power obtained in this investigation was 103 watts using a turbine with radius 150 cm, height 45 cm, shutter angle 30 degree and curved shutter form. The torque and power increases with increase in shutter angle up to 30 degree starting from 12 degree and then decreases with increase in shutter opening angle up to 48 degree. From this investigation, it is clear that the newly developed STVAWT is working efficiently at 30 degree shutter opening angle and the curved shutter form is found to be more efficient as compared to the straight shutter form.

The research project HEAVYcOPTer, a sub task of the European R&D program Clean-Sky GRC2 [1], is devoted to the efficient design and the shape optimization of the Agusta Westland AW101 helicopter turboshaft engine intake and exhaust system, to be carried out by means of advanced multi-objective optimization algorithms coupled with CFD Navier-Stokes solvers. The present paper describes the outcomes of HEAVYcOPTer in relation to the air intakes shape optimisation activities. This paper describes the technical details of such program. The optimisation method chosen for the redesign of the engine installation involves the application of the state of the art genetic algorithm GDEA, developed at the University of Padova and successfully applied in several fluid-dynamics applications, especially in the field of turbomachinery. For the present application, the set of geometrical designs constituting the genetic algorithm population are generated by means of morphing the original CFD model surface mesh: shapes are applied to baseline surface nodes with a displacement intensity driven by the GA chosen scaling factors. Then, CFD models of new designs are automatically generated and analyzed by the flow solver, returning to the GA the evaluation of the selected objective functions required in order to evolve the population in the next step of the evolutionary process. AW101 intakes have been optimised following a multi-objective/multi-point approach, minimizing inlet total pressure loss in both hovering and forward flight conditions simultaneously; optimised solutions were also constrained so as to not exceed the total pressure distortion level at the engine aerodynamic interface plane, so as to ensure inlet/engine compatibility with respect to the compressor surge limit. This approach ensured the improvement of the engine/airframe integration efficiency for the overall rotorcraft flight envelop, reducing fuel burn and increasing the helicopter propulsive efficiency.

A wave-rotor pressure-gain combustor (WRPGC) ideally provides constant-volume combustion and enables a gas turbine engine to operate on the Humphrey-Atkinson cycle. It exploits pressure (both compression and expansion) waves and confined propagating combustion to achieve pressure rise inside the combustor. This study first presents thermodynamic cycle analysis to illustrate the improvements of a gas turbine engine possible with a wave rotor combustor. Thereafter, non-steady reacting simulations are used to examine features and characteristics of a combustor rig that reproduces key features of a WRPGC. In the thermodynamic analysis, performance parameters such as thermal efficiency and specific power are estimated for different operating conditions (compressor pressure ratio and turbine inlet temperature). The performance of the WRPGC is compared with the conventional unrecuperated and recuperated engines that operates on the Brayton cycle. Fuel consumption may be reduced substantially with WRPGC introduction, while concomitantly boosting power. Simulations have been performed of the ignition of propane by a hot gas jet and subsequent turbulent flame propagation and shock-flame interaction.

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.

This paper presents experimental and numerical results on a single stage burner configuration with flameless/MILD combustion with liquid fuel. The proposed burner configuration is designed for 20 kW thermal input with heat intensity of ∼ 5 MW/m 3 using kerosene as fuel and air at ambient conditions as the oxidizer. Air is injected through four tangential injection holes near the bottom of the combustor results high swirl flow in the combustor helps to enhance the internal recirculation of the combustion products. Computational and experimental analysis is carried out simultaneously for optimization of combustor configuration. In swirl combustor configurations the reactants dilution ration (recirculation) is function of combustor geometry, exit diameter and inlet velocity of air. In the first step of study four different combustor configurations are considered, one cylindrical and three conical combustors with diverging angles of 30°, 45° and 60°. In the second step the effect of exit port diameter on the recirculation and quality of flameless combustion is studied. The exit port diameter varied from 80 mm to 25 mm. In the third step the inlet velocities of air varied by inserting different inlet diameters of 2 to 7 mm in a step of 1mm. Based on combustion completeness and emission analysis, the 60° diverging angle combustor with air inlet diameter of 4 or 5 mm and 25 mm exit diameter is considered as optimistic configuration to obtain flameless combustion mode with liquid fuels. The acoustic emissions and the emissions of CO and NO X are measured for different configurations.

Jets at higher Reynolds numbers have a high concentration of energy in the small scales in the nozzle vicinity. This is challenging for LES, potentially placing severe demands on grid density. To circumvent this, we propose a novel procedure based on well known Reynolds number (Re) independence of jets. We reduce the jet Re whilst rescaling the boundary layer properties to maintain incoming boundary layer thickness consistent with high Re jet. The simulations are carried out using hybrid largeeddy simulation type of approach which is incorporated by using near wall turbulence model with modified properties. No Subgrid Scale (SGS) model is used in these simulations. Hence, they effectively become Numerical Large Eddy Simulation (NLES) with Reynolds-averaged Navier-Stokes (RANS) covering the full boundary layer region. The noise post processing is carried out using Ffowcs-Williams-Hawking (FWH) approach. The simulations are made for Mach numbers (M) of 0.75 and 0.875. The results for Overall Sound Pressure Level (OASPL) are observed to be within 2–3% accuracy range and directivity of sound is also captured accurately for both the cases. The low Re simulations hence, can be more beneficial in saving time and cost of the simulation while providing reasonably accurate results.

The Bladeless Turbine is a remarkable machine in terms of simplicity, robustness, efficiency, and applicability but little as known, even among today’s engineers, about how it works and how well it performs alongside conventional turbines. This paper presents effect of disc spacing and disc surface roughness on the performance of bladeless turbine. In this investigation, instead of blades, closely packed parallel discs are used. Resistance to fluid flow between the plates results in energy transfer to the shaft. High velocity water enters the disc pack through inlet nozzle path tangent to the outer edge of the discs. Convergent nozzle imparts high velocity water jet tangentially on disc thickness. Lower-energy water spirals toward the central exit port, adhesion, drag and impulse forces continue to convert kinetic energy to shaft rotational power. However, The Bladeless Turbine and a flexible test rig have been designed and manufactured, and experimental results are presented. An analysis of the performance and efficiency of the disc turbine is carried out. The design philosophy of the flexible test rig has been explained. Various complementary methods of measurement have been implemented and compared, and several operational experiences have been noted Experimental results for a 152mm diameter and 2mm thick discs of turbine are presented, which shows the variation of torque, output power, and efficiency as a function of angular speed. Measurements of static pressure are also taken at the inlet, Many design considerations and operational experiences are discussed. The effect of each parameter on the torque and power has been analyzed. It has been found that the spacing and surface finish has a significant effect on the power of the turbine. The maximum power obtained in this investigation was 33watts for 6discs and 0.5 mm spacing between discs with rough surface ( spiral Groove). The torque and power increases with decrease in spacing upto 0.5mm and increase in surface roughness value (Ra) 500 microns. From this investigation, it is clear that the developed bladeless turbine is working efficiently at 0.5mm spacing and 500 microns roughness disc surface.

In this paper, employing an Eulerian-Eulerian method, based on a non-equilibrium point of view, the simulation of a two-phase vapor-liquid flow in the blade to blade passage has been numerically studied. In the following, the governing phenomena in blade to blade passage as well as important parameters in the blade design procedure have been investigated. In the first section, after introducing the physical governing phenomenon in the subsonic and supersonic outflows, the mechanisms will be discussed which play a way more important role in changing the outlet pressure compared to the other less important ones. In the next section, the blade loading distribution, which is one of the most important parameters in designing the steam-turbine blades, has been studied. At the last part noting the significance of gas mass phase fraction on the blade surface, this parameter is comprehensively investigated. In order to perform a more successful analysis, the blade to blade passage has been simulated by a convergent-divergent nozzle. Taking into account the governing phenomena in the convergent-divergent nozzle and using the developed simulation model, most of experimental observations could be reproduced and also successfully understood.

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.

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.

This paper discusses the results of a parametric study of a pair of contra-rotating axial fan rotors. The rotors were designed to deliver a mass flow of 6 kg/s at 2400 rpm. The blades were designed with a low hub-tip ratio of 0.35 and an aspect ratio of 3.0. Numerical and experimental studies were carried out on these contra-rotating rotors operating at a Reynolds number of 1.25 × 10 5 (based on blade chord). The axial spacing between the rotors was varied between 50 to 120 % of the chord of rotor 1. The performance of the rotors was evaluated at each of these spacing at design and off-design speeds. The results from the numerical study (using ANSYS CFX) were validated using experimental data. In spite of certain limitations of CFD under certain operating conditions, it was observed that the results agreed well with those from the experiments. The performance of the fan was evaluated based on the variations of total pressure, velocity components and flow angles at design and off-design operating conditions. The measurement of total pressure, flow angles etc. are taken upstream of the first rotor, between the two rotors and downstream of the second rotor. It was observed that the aerodynamics of the flow through a contra rotating stage is significantly influenced by the axial spacing between the rotors and the speed ratio of the rotors. With increasing speed ratios, the strong suction generated by the second rotor, improves the stage pressure rise and the stall margin. Lower axial spacing on the other hand, changes the flow incidence to the second rotor and thereby improves the overall performance of the stage. The performance is investigated at different speed ratios of the rotors at varying axial spacing.

Performance of intake duct with fixed inlet trajectory and different area distributions have been analyzed using a commercial CFD (Computational Fluid Dynamics) software. The performance have been evaluated for fixed boundary conditions. The area distributions studied are defined by varying cross sectional area at different locations of intake duct by keeping the inlet and exit area same. The performance of the intake ducts are studied in terms of the pressure recovery coefficient, total pressure loss, pressure recovery factor and distortion coefficient in the present work. The motion caused by the change in centerline curvature is analyzed. The objective of the work is to derive a shape of the duct with minimum distortion of the flow and maximum pressure recovery.

In this paper, a numerical method is presented to solve the two-dimensional two-phase steam flow over a series of geometries (such as nozzles, expansion corners and steam turbine blade-to-blade passages) by means of equilibrium thermodynamics model. The flow is assumed to be compressible and inviscid and obeys the ideal gas equation of state. The resulted equations are then numerically solved by the Roe’s FDS time marching scheme that has recently been modified to allow for two-phase effects. Validations of condensing steam flow through vapor nozzles have been performed, where good agreement has been achieved. Detailed parametric studies monitoring the influence of (I) the geometry expansion rate, (II) the inlet total temperature and pressure, and (III) the expansion fan or shock waves on the location of condensation onset and the rate of condensation are given. Finally as a case study, expansion of steam flow through a steam turbine blade-to-blade passage is considered, and condensation or evaporation of the steam flow through the passage and fate of the wet flow through the fan or shocks were observed.

The main objective of this work is the validation of Computational Fluid Dynamics (CFD) code used for analysis of transonic axial compressors. NASA Rotor 35 is used here as test case for validation. In this work, computations are performed using parallelized RANS code, to predict the transonic axial compressor rotor flow characteristics. Advection Upstream Splitting Method (AUSM) scheme has been used. A Multiple Frame of Reference approach has been used to model the rotor passage. Spalart-Allmaras turbulence model is used to model turbulence. Multiblock Structured mesh is used. Performance characteristics for the entire range of operation, from maximum mass flow rate till maximum pressure ratio, have been simulated. The results obtained are comparable with experimental data within 5–10% error. Investigations have been carried out to study the effect of varying tip clearance in NASA Rotor 35. The present work is intended to study the clearance flow trajectory as a function of varying tip clearance. The effects of shock/vortex interaction in tip clearance region are also studied. The effects of tip clearance size on the generation and evolution of the end-wall vortical structures are discussed by investigating their evolutionary trajectories. By this study, it is observed that as tip clearance reduces, clearance flow trajectory moves downstream. From this it can be concluded that if tip clearance increases, tip clearance vortices expand. This may help in casing-treatment or tip-treatment to mitigate the loss in the performance, if the tip clearance increases.

Studies on the effects of stator reduced frequency in low pressure turbines have shown that periodic wake-induced unsteadiness can increase steady flow circulation by as much as 15% and reduce losses compared to a steady flow datum. A large separation bubble downstream of peak suction that formed under steady flow conditions was periodically suppressed by wake passing events, resulting in significantly reduced losses within the boundary layer. This research extends this concept to a controlled diffusion compressor stator blade with a circular arc leading edge. The blade was placed inside a large scale, two-dimensional, cascade with a rotating bar mechanism used to simulate an upstream rotor blade row. The blade profile has been shown to experience leading edge separations and subsequent transition on both the pressure and suction surfaces due to a velocity overspeed caused by discontinuities in surface curvature. Testing was carried out at reduced frequencies of 0.47, 0.94 and 1.88 at the design inlet flow angle 45.5° and Reynolds number based on chord of 230,000. The freestream turbulence intensity was 4.0%. A range of experimental measurements were used to look at the blade’s performance: high resolution time-averaged blade surface static pressure measurements, inlet and exit 3-hole probe traverses and instantaneous, ensemble averaged and time average surface mounted hot-film measurements for the calculation of turbulent intermittency and quasi wall-shear stress. Results showed that increasing the stator reduced frequency from, 0–1.88, increased the overall blade pressure loss. The losses generated by the pressure surface and suction surface differed significantly and are affected very differently. The pressure surface demonstrated a clear reduction in loss with an increase in reduced frequency whereas the opposite trend was seen on the suction surface. Wake-induced turbulent strips suppressed the formation of leading edge separation bubbles that formed under steady flow conditions and in between wake passing events. Wake-induced turbulent strips reduced in width and level of turbulent intermittency through the favorable pressure gradients leading to peak suction and grew in the adverse pressure gradient of the velocity overspeed. The flow between wake-induced turbulent strips partially relaminarised through the favorable pressure gradient leading to peak suction.

Reviewed the historical development of the supersonic axial flow compressor, and gave an outlook for its future developments and research orientations. According to the internal flow characteristics of the conventional supersonic axial flow compressors, put forward a high load of supersonic axial compressor aerodynamic design principle. A preliminary design verification of the principle has been carried. The 3D viscous numerical simulation results show that, under the tip tangential speed 360m / s, has achieved a stage pressure ratio 2.3 with efficiency 86.5%. In addition, considering the rotor under impulse condition can get the maximum rotor total pressure ratio with high efficiency, a design principle has also been put forward to solve the high entrance Mach number problem of the downstream stator. But the numerical simulation results show that the multi-shock structure does not have any advantages to reduce the stator losses.

Present study is aimed to investigate the relationship between blade shape of Savonius rotor and its optimum overlap, an important parameter that can improve the rotor’s performance superbly. To set a relation between rotor’s shape and its overlap, a new parameter called “trailing angle” was introduced in this study and by a novel design method, several models were designed to represent a wide range of trailing angles. The investigation has been carried out by using Reynolds Stress Method (RSM) through ANSYS Fluent commercial software package. The application of the method was validated by comparing the results of a conventional model with experimental data acquired from one of the references. Results of this study have proved that optimum overlap ratio is associated with blades’ trailing angle and by increasing the trailing angle, optimum overlap would decrease. The results also demonstrate that by application of a proper overlap ratio, power coefficient of the rotor would increase by nearly 22 percent.

This paper presents the results of numerical studies on the effect of endwall reshaping on the performance and stall margin of a centrifugal compressor. The endwall (shroud) of a baseline compressor was reshaped by introducing a convex hump at a location downstream of the inducer where the flow turns from axial to radial direction. The depth of the hump was varied as 3%, 6%, and 9% of the local impeller blade height. The impeller was also locally reshaped to match the shroud, maintaining the baseline tip clearance. Other geometric parameters of the reshaped impeller were kept same as those of the baseline impeller. An implicit, coupled, density based solver (ANSYS FLUENT) was used with first order upwind discretization scheme. The local reshaping of the endwall and impeller blades showed a positive tendency of increase in compressor stall margin without any appreciable reduction in efficiency. Endwall reshaping of 9% of the local blade height resulted in reducing the stall mass flow rate by 6.4 % of the baseline stall flow rate. Further, the peak total pressure ratio showed a marginal increase of 0.8%. Reshaping of the endwall was observed to cause local acceleration of the flow, thus weakening separation bubble and delaying inception of stall.

Various rib turbulator geometries have been used earlier to investigate the heat transfer and fluid flow characteristics owing to its vast industrial applications. Permeable ribs are reported to provide better performance in terms of heat transfer enhancement and pressure penalty. The trapezoidal rib with variable downstream chamfering and a centrally placed longitudinal slit has been used for the detailed flow field and heat transfer investigations using particle image velocimetry and liquid crystal thermography. Objective of the present work is to study the combined effect of rib chamfering along with a continuous slit and establish an understanding of the underlying flow mechanism and the corresponding heat transfer distribution. Experiments were carried out for hydraulic diameter based Reynolds numbers of 61480 in a rectangular duct for flow over rib with 12.5% blockage ratio and 25% open area ratio. The chamfer angles were varied from 0 to 20 degree. It has been observed that the slitted rib cause shorter reattachment length and reduced pressure penalty as compared to its solid counterpart.

In a gas turbine combustor, it is necessary to use a diffuser to decelerate the high velocity air stream delivered by the compressor and thus avoid high total pressure loss. The interaction between the diffuser and combustor external flows plays a key role in controlling the pressure loss, air flow distribution around the combustor liner. Flow through casing-liner annulus is crucial as it feeds air to the primary, secondary and dilution holes. It is important that the annulus flow has sufficient static pressure to achieve adequate penetration of the jets. Moreover, the correct proportion of air enters the combustor liner through the dome and the various ports to maintain stable operation and good quality outlet condition. Length of combustor can be reduced if a provision is made for sufficient diffusion in the dump region. In the present numerical study, three can-combustor models of different geometry with a constant dump-gap have been analyzed with emphasis on the flow through annulus. A comparison has been made amongst the three models in terms of flow uniformity, static pressure recovery and total pressure loss. It is observed that flow uniformity in the annulus region is improved if a small divergence in length and a curved shape step height casing is made.