Rotor Blade Sweep Effect on the Performance of a Small Axial Supersonic Impulse Turbine (original) (raw)
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12th European Conference on Turbomachinery Fluid Dynamics and hermodynamics
The paper focuses on methods for an efficiency increasing of supersonic axial impulse turbines. There was chosen the axial turbine stage with axisymmetric nozzles and mean diameter 103.5 mm as the base model for a numerical and experimental investigation. There were chosen two modifications of the base design to provide an efficiency increasing: modification of the rotor leading edge design and modification of the rotor hub endwall design. The effect of these two factors on the turbine efficiency was analyzed separately. A preliminary numerical investigation showed that the rotor hub endwall modification provided an efficiency increasing up to 2% in comparison with the base model. Meanwhile, modification of the rotor leading edge shape did not provided significant efficiency increasing in comparison with the base model. * blade wheel outlet angle in relative frame (deg.
Comparison of Various Supersonic Turbine Tip Designs to Minimize Aerodynamic Loss and Tip Heating
Volume 7: Turbomachinery, Parts A, B, and C, 2011
The rotor tips of axial turbines experience high heat flux and are the cause of aerodynamic losses due to tip clearance flows, and in the case of supersonic tips, shocks. As stage loadings increase, the flow in the tip gap approaches and exceeds sonic conditions. This introduces effects such as shock-boundary layer interactions and choked flow that are not observed for subsonic tip flows that have been studied extensively in literature. This work simulates the tip clearance flow for a flat tip, a diverging tip gap and several contoured tips to assess the possibility of minimizing tip heat flux while maintaining a constant massflow from the pressure side to the suction side of the rotor, through the tip clearance. The CFD code GlennHT was used for the simulations. Due to the strong favorable pressure gradients the simulations assumed laminar conditions in the tip gap. The nominal tip gap width to height ratio for this study is 6.0. The Reynolds number of the flow is 2.4×105 based on ...
12th European Conference on Turbomachinery Fluid Dynamics and hermodynamics, 2017
The paper describes the results of a numerical and experimental research program addressing the aerodynamic investigation on the performance of blade profiles specifically developed for application in highly loaded impulse type turbine stages. The industrial requirements driving toward the adoption of highly loaded stage solutions are presented, along with an estimation of the profiles operating parameters. Two stator vanes and one rotor blade profile have been developed and extensively tested by means of flow field measurements and schlieren visualization in a transonic blow-down wind tunnel for linear cascades. Experimental results for the relevant operating conditions are presented, providing validation data for the CFD model used for blade design and evidencing that the main goals of the design optimization procedure have been achieved.
Tip leakage phenomenon in axial compressors is sensitive to the flow incidence, flow coefficient, tip gap height and the pressure gradients. All these geometric/flow features are considerably altered by blade stagger angle. Literature on the stagger angle effects in compressors is scarce; and indeed, such studies for various tip gap heights have not been reported yet. The present paper reports the effect of rotor stagger angle on the performance of subsonic axial compressor rotor with different forward sweep configurations and for various rotor tip clearances. The computational model for the study utilizes finest hexahedral grids. A commercial CFD package ANSYS® CFX 11.0 was used with standard k-ω turbulence model for the simulations. CFD results were well validated with experiments. The following observations were made: At higher stagger angles, flow separates from upstream suction surface locations. Little tip clearance had a positive effect for certain stagger angle increments owing to beneficial interaction of leakage flows with the local flow field. However, severe performance loss was observed at higher stagger settings with large clearances. As the stagger angle was increased, vena contracta effect was highly reduced. At high stagger angles, the flow was observed to leak in a more "axially-reversed" fashion through the tip gap. The deep lowest pressure zones near the pressure surface of the tip are due to the effect of 'vena contracta.' Such zones near the suction surface edge of the tip are due to flow acceleration. This particular feature is directly correlated with the tip aerofoil loading and thickness-to-tip gap ratio.
Application of Sweep to Low Pressure Turbine Blade for Tip Flow Containment
48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2012
A numerical investigation on a low-speed linear cascade has been done to study the impact of sweep applied at the tip of a turbine rotor blade on tip leakage flow. Two forward and two backward swept blade tip modifications have been applied to the T106 profile to create new blade cascade configurations. The aim of applying sweep to the blade's tip is an attempt to reduce tip losses. This can be achieved by reducing the leakage mass flow rate or by altering the leakage flow to reduce the leakage vortex strength, which is the main contributor to tip losses. A detailed comparison of the tip region including blade loading, mass flow, turbulent kinetic energy, pressure gradients, and velocity vectors, has been conducted to gain insight into the flow structures within the tip gap. A similar detailed comparison has been conducted for the main passage to determine leakage vortex formation, location, size, turbulent kinetic energy and interaction with secondary flow. Forward sweep was observed to reduce mass flow rate, however, an increase to the tip gap vortex size and strength was also observed; which raised the turbulent kinetic energy introduced into the leakage flow, thereby increasing the size and strength of the leakage vortex and increasing pressure loss. Backward sweep, contrarily, increased mass flow but reduced the tip gap vortex, thus decreasing the turbulent kinetic energy introduced. Therefore the leakage vortex size and strength was reduced, ensuring a reduction in pressure loss. Nomenclature
Applied and Computational Mechanics, 2021
Numerical simulations of 2D compressible flow through the tip-section turbine blade cascade with a flat profile and the supersonic inlet were carried out by the OpenFOAM code using the Favre-averaged Navier-Stokes equations completed by the γ-Re_θt bypass transition model with the SST turbulence model. Predictions completed for nominal regimes were concentrated particularly on the effect of the shock-wave/boundary layer interaction on the transition to turbulence. Further, the link between the inlet Mach number and the inlet flow angle i.e. the so called unique incidence rule was studied. Obtained numerical results were compared with experimental data covering optical and pressure measurements.
Study of Flow Patterns in Radial and Back Swept Turbine Rotor under Design and Off-Design Conditions
Journal of Applied Fluid Mechanics
Paper details the numerical investigation of flow patterns in a conventional radial turbine compared with a back swept design for same application. The blade geometry of a designed turbine from a 25kW micro gas turbine was used as a baseline. A back swept blade was subsequently designed for the rotor, which departed from the conventional radial inlet blade angle to incorporate up to 25° inlet blade angle. A comparative numerical analysis between the two geometries is presented. While operating at lower than optimum velocity ratios (U/C), the 25° back swept blade offers significant increases in efficiency. In turbocharger since the turbine typically experiences lower than optimum velocity ratios, this improvement in the efficiency at offdesign condition could significantly improve turbocharger performance. The numerical predictions show offdesign performance gains of the order of 4.61% can be achieved, while maintaining design point efficiency.
International Journal of Mechanical Engineering Technologies and Applications, 2023
An impulse turbine uses drag force on its blades to produce torque on its rotor. As fluid flows over the blades, pressure changes occur at the nozzle, which increases the fluid's velocity and reduces the static pressure at the nozzle outlet. The high-momentum fluid then impinges on the rotor blades, generating frictional force and resulting in torque production. To study the impact of blade shape and number on the turbine's performance, simulations were conducted. The results indicate that blades with an angle of 0° and 180° are optimal for creating high-pressure vortices on the concave surface of the blade. Additionally , more blades always result in higher torque and power output by increasing the active area of the blades. However, in the case of blades with an angle of 0° and 180°, 8 blades produced more torque than 12 blades with an angle of 0° and 90°. Therefore , blades with an angle of 0° and 180° are highly effective at generating drag force and producing torque.
Computational Estimation of Flow through the C-D Supersonic Nozzle and Impulse Turbine Using CFD
In this paper, CFD analysis of flow within, Convergent – Divergent rectangular super sonic nozzle and super sonic impulse turbine with partial admission have been performed. The analysis has been performed according to shape of a super sonic nozzle and length of axial clearance and the objective is to investigate the effect of nozzle-rotor interaction on turbine's performance. It is found that nozzle-rotor interaction losses are largely dependent on axial clearance, which affects the flow within nozzle and the extent of flow expansion. Therefore selecting appropriate length of axial clearance can decrease nozzle-rotor interaction losses. The work is carried in two stages:1) Modeling and analysis of flow for rectangular convergent divergent super sonic nozzle. 2) Prediction of optimal axial gap between the nozzle and rotor blades by allowing the above nozzle flow. In the present work, using a finite volume commercial code, ANSYS FLUENT 14.5, carries out flow through the convergent divergent nozzle study. The nozzle geometry is modeled and grid is generated using ANSYS14.5 Software. Computational results are in good agreement with the experimental ones.
Computational Analysis of Unsteady Flow in a Partial Admission Supersonic Turbine Stage
Volume 2D: Turbomachinery, 2014
Turbines used in upper stage engine for a rocket are sometimes designed as a supersonic turbine with partial admission. This study deals with numerical investigation of supersonic partial admission turbine in order to understand influences on the unsteady flow pattern, turbine losses and aerodynamic forces on rotor blades due to partial admission configuration. Two-dimensional CFD analysis is conducted using "Numerical Turbine" code. Its governing equation is URANS (Unsteady Reynolds Averaged Navier-Stokes Simulation) and fourth-order MUSCL TVD scheme is used for advection scheme. The unsteady simulation indicates that strongly nonuniform circumferential flow field is created due to the partial admission configuration and it especially becomes complex at 1 st stage because of shock waves. Some very high or low flow velocity regions are created around the blockage sector. Nozzle exit flow is rapidly accelerated at the inlet of blockage sector and strong rotor LE shock waves are created. In contrast, at rotor blade passages and Stator2 blade passages existing behind the blockage sector, working gas almost stagnates. Large flow separations and flow mixings occur because of the partial admission configuration. As a result, additional strong dissipations are caused and the magnitude of entropy at the turbine exit is approximately 1.5 times higher than that of the full admission. Rotor1 blades experience strong unsteady aerodynamic force variations. The aerodynamic forces greatly vary when the Rotor1 blade passes through the blockage inlet region. The unsteady force in frequency domain indicates that many unsteady force components exist in wide frequency region and the blockage passing frequency component becomes pronounced in the circumferential direction force. Unsteady forces on Rotor2 blades are characterized by a low frequency fluctuation due to the blockage passing.