Unsteady Operation of a Highly Supersonic Organic Rankine Cycle Turbine

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Enrico Rinaldi ◽  
Rene Pecnik ◽  
Piero Colonna
Keyword(s):  
Boundary Layer ◽  
Organic Rankine Cycle ◽  
Suction Side ◽  
Rankine Cycle ◽  
Boundary Layer Separation ◽  
Expansion Ratio ◽  
Ideal Gas ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Shock Interactions

Organic Rankine cycle (ORC) turbogenerators are the most viable option to convert sustainable energy sources in the low-to-medium power output range (from tens of kWe to several MWe). The design of efficient ORC turbines is particularly challenging due to their inherent unsteady nature (high expansion ratios and low speed of sound of organic compounds) and to the fact that the expansion encompasses thermodynamic states in the dense vapor region, where the ideal gas assumption does not hold. This work investigates the unsteady nonideal fluid dynamics and performance of a high expansion ratio ORC turbine by means of detailed Reynolds-averaged Navier–Stokes (RANS) simulations. The complex shock interactions resulting from the supersonic flow (M ≈ 2.8 at the vanes exit) are related to the blade loading, which can fluctuate up to 60% of the time-averaged value. A detailed loss analysis shows that shock-induced boundary layer separation on the suction side of the rotor blades is responsible for most of the losses in the rotor, and that further significant contributions are given by the boundary layer in the diverging part of the stator and by trailing edge losses. Efficiency loss due to unsteady interactions is quantified in 1.4% in absolute percentage points at design rotational speed. Thermophysical properties are found to feature large variations due to temperature even after the strong expansion in the nozzle vanes, thus supporting the use of accurate fluid models in the whole turbine stage.

Numerical Modelling of Fluid-Structure Interaction in a Turbine Stage for 3D Viscous Flow in Nominal and Off- Design Regimes

2010 ◽  
Author(s):  
Romuald Rzadkowski ◽  
Vitaly Gnesin ◽  
Lubov Kolodyazhnaya
Keyword(s):  
Boundary Layer ◽  
Viscous Flow ◽  
Gas Turbines ◽  
Element Model ◽  
Boundary Layer Separation ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Turbine Stage ◽  
Flow Distortion ◽  
Boundary Layer Interaction

In recent years there have been major developments in turbomachinery aeroelasticity methods. There are now greater possibilities to predict blade vibrations arising from self-excitation or inlet flow distortion. This is not only important with regard to aircraft compressor and fan blade rows, but also in the case of the last stages of steam and gas turbines working in highly loaded off-design conditions. In order to predict the unsteady pressure loads and aeroelastic behaviour of blades (including the computation of shock waves, shock/boundary layer interaction and boundary layer separation), complete Reynolds-averaged Navier-Stokes (RANS) equations are used in modelling complex and off-design cases of turbomachinery flows. In this paper the 3D RANS solver, including a modified Baldwin and Lomax algebraic eddy viscous turbulence model, is presented to calculate unsteady viscous flow through the turbine stage, while taking into account the blade oscillations but without the separating of outer excitation and unsteady effects caused by blade motion. The numerical method uses the second order by time and coordinates an explicit finite-volume Godunov’s type difference scheme and a moving H-O structured grid. The structure analysis uses the modal approach and a 3D finite element model of blade. To validate the numerical viscous code, the numerical calculation results were compared with the 11th Standard Configuration measurements. Presented here are the numerical analysis results for the aeroelastic behaviour of a steam turbine last stage with 760 mm rotor blades in a nominal and an off-design regime.

Investigation of Scramjet Inlet Flow Using a Flux-Splitting Solver

2010 ◽  
Author(s):  
D. M. Holian ◽  
R. R. Mankbadi
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Shock Interactions ◽  
Flux Splitting ◽  
Operation Conditions ◽  
Inlet Geometry ◽  
Proper Operation ◽  
Good Agreement

A detailed analysis is carried out on a rectangular scramjet inlet to analyze the flow field. The focus is on examining boundary layer separation and shock interactions to ensure proper operation of the inlet. We developed herein a flux-splitting Navier Stokes solver to be used for optimizing the inlet geometry and operation conditions. The results seem to be in good agreement with that of FLUENT CFD software and explain the experimental results of Haberle (2008).

Development and Design of a Two-Stage Contraction Zone and Test Section of an Organic Rankine Cycle Wind Tunnel

2016 ◽  
Author(s):  
Maximilian Passmann ◽  
Felix Reinker ◽  
Karsten Hasselmann ◽  
Stefan aus der Wiesche ◽  
Franz Joos
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Flow Separation ◽  
Test Section ◽  
Organic Rankine Cycle ◽  
Rankine Cycle ◽  
Boundary Layer Separation ◽  
Two Stage ◽  
Cross Sectional ◽  
Flow Deviation

This contribution presents the development and design of a two-stage contraction zone and modular test section for a closed loop Organic Rankine Cycle (ORC) wind tunnel. The first contraction consists of four truncated cones, whose length and angle of inclination are derived from a two-stage optimization procedure, with the objective to minimize flow deviation and to avoid boundary-layer separation. The geometrical optimization yields a profile with minor deviation to the ideal polynomial shape, whereas the flow optimized shape minimizes flow separation at the break-points between the single conical pieces. The second contraction has to perform two major tasks, namely the acceleration of the flow up to a Mach number of Ma = 0.8 for organic fluids and the transformation of the circular inlet to a rectangular outlet cross-sectional shape, required by the working section. The circular-to-rectangular transition is accomplished by variation of the generalized ellipse, also known as Lamé curve. Smooth polynomials are then used to define the reduction of cross-sectional area. A comprehensive number of contraction geometries with fixed contraction ratio, variable length, and different points of inflection are analyzed with regards to minimum flow deviation, the avoidance of flow separation, as well as a uniform velocity field at the contraction outlet. A semi-analytical approach based on a potential flow solution in combination with the Stratford criterion is the basis for evaluating boundary-layer separation. The design of a two-part modular diffuser, based on the concept of a dumped diffuser, as commonly encountered in gas turbine design, is presented. The numerical results are compared with analytical findings and special characteristics of the different designs are explained. Finally, the overall design concept of the test section is presented.

Effects of Tip Gap Size on the Flow Structure in the Tip Region of an Axial Turbomachine

2015 ◽  
Author(s):  
Yuanchao Li ◽  
David Tan ◽  
Huang Chen ◽  
Joseph Katz
Keyword(s):  
Boundary Layer ◽  
Flow Structure ◽  
Cross Sections ◽  
Substantial Reduction ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Rotor Blades ◽  
Gap Size

This experimental study examines the effect of tip gap size on the flow structure and turbulence in the tip region of an axial turbomachine. The experiments have been performed in the Johns Hopkins University (JHU) optically index-matched facility using an axial compressor settings designed based on the geometry of the inlet guide vanes (IGV) and the first stage of the Low Speed Axial Compressor (LSAC) facility at NASA Glenn. Two sets of rotor blades with similar cross sections, but with tip gap sizes of 0.49% and 2.3% of the blade chord (or 1.1% and 5.4% of the blade span) have been installed and tested. The measurements include performance tests, visualization of the tip leakage vortex (TLV) using cavitation, and stereo PIV (SPIV) measurements in several meridional planes. Increasing the tip gap size causes a substantial reduction in pressure rise across the machine for the same flow rate. The cavitation images, whose trends agree with the velocity and vorticity distributions obtained by the SPIV measurements, show that TLV rollup in the less loaded blade occurs at later chordwise location, and that the vortex remains located closer to the suction side (SS) corner of the originating blade. The delayed detachment from the blade with increasing gap is attributed to the increase of distance of the ‘image vortex’ (wall interaction) from the TLV. The wider gap also reduces the entrainment by the TLV of the endwall boundary layer after it separates at the point where the backward leakage flow meets the main passage flow. The previously observed TLV breakup, which is evident for the narrow gap in the aft part of the rotor passage, is delayed significantly for the wider gap. Consistent changes also appear in the distributions of turbulent kinetic energy, which peaks in the vicinity of the TLV core, the endwall boundary layer separation, and in the shear layer connecting the TLV center to the SS corner of the blade tip.

Self-Excited Oscillation of Transonic Flow Around an Airfoil in Two-Dimensional Channel

1989 ◽  
Author(s):  
Kazuomi Yamamoto ◽  
Yoshimichi Tanida
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Transonic Flow ◽  
Stokes Equations ◽  
Flow Region ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Layer Separation ◽  
Excited Oscillation

A self-excited oscillation of transonic flow in a simplified cascade model was investigated experimentally, theoretically and numerically. The measurements of the shock wave and wake motions, and unsteady static pressure field predict a closed loop mechanism, in which the pressure disturbance, that is generated by the oscillation of boundary layer separation, propagates upstream in the main flow and forces the shock wave to oscillate, and then the shock oscillation disturbs the boundary layer separation again. A one-dimensional analysis confirms that the self-excited oscillation occurs in the proposed mechanism. Finally, a numerical simulation of the Navier-Stokes equations reveals the unsteady flow structure of the reversed flow region around the trailing edge, which induces the large flow separation to bring about the anti-phase oscillation.

1-D Model Analysis of Tesla Turbine for Small Scale Organic Rankine Cycle (ORC) System

2017 ◽  
Author(s):  
Jian Song ◽  
Chun-wei Gu
Keyword(s):  
Large Scale ◽  
Low Cost ◽  
Organic Rankine Cycle ◽  
Axial Flow ◽  
Rankine Cycle ◽  
Expansion Ratio ◽  
Working Fluid ◽  
Small Scale ◽  
Low Grade ◽  
D Model

Energy shortage and environmental deterioration are two crucial issues that the developing world has to face. In order to solve these problems, conversion of low grade energy is attracting broad attention. Among all of the existing technologies, Organic Rankine Cycle (ORC) has been proven to be one of the most effective methods for the utilization of low grade heat sources. Turbine is a key component in ORC system and it plays an important role in system performance. Traditional turbine expanders, the axial flow turbine and the radial inflow turbine are typically selected in large scale ORC systems. However, in small and micro scale systems, traditional turbine expanders are not suitable due to large flow loss and high rotation speed. In this case, Tesla turbine allows a low-cost and reliable design for the organic expander that could be an attractive option for small scale ORC systems. A 1-D model of Tesla turbine is presented in this paper, which mainly focuses on the flow characteristics and the momentum transfer. This study improves the 1-D model, taking the nozzle limit expansion ratio into consideration, which is related to the installation angle of the nozzle and the specific heat ratio of the working fluid. The improved model is used to analyze Tesla turbine performance and predict turbine efficiency. Thermodynamic analysis is conducted for a small scale ORC system. The simulation results reveal that the ORC system can generate a considerable net power output. Therefore, Tesla turbine can be regarded as a potential choice to be applied in small scale ORC systems.

Numerical Study of Wall Influence on Boundary-Layer Transition for Two-Dimensional NACA 16-012 and 4412 Hydrofoil Sections

1990 ◽  
Vol 34 (01) ◽  
pp. 38-47
Author(s):  
R. Latorre ◽  
R. Baubeau
Keyword(s):  
Boundary Layer ◽  
Test Section ◽  
Numerical Study ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Turbulent Separation ◽  
Side Pressure ◽  
Effective Angle

One of the difficulties in hydrofoil model tests is the relatively low Reynolds number of the test piece and the presence of the test section walls. This paper presents the results of systematic calculations of the potential flow field of NA 4412 and NACA 16-012 hydrofoil in a test section with wall-to-chord ratios h/c -1.0. The corresponding boundary-layer calculations using the CERT calculation scheme are presented to show the influence of the nearby walls on shifting the location of the boundary-layer laminar-turbulent separation as well as turbulent separation. By introducing an effective angle of attack, it is possible to obtain close agreement in the calculated and measured suction side pressure distortion as well as the locations of the boundary-layer separation and transition.

Modelling of By-Pass Transition With Conditioned Navier-Stokes Equations and a K-E Model Adapted for Intermittency

1994 ◽  
Author(s):  
J. Steelant ◽  
E. Dick
Keyword(s):  
Boundary Layer ◽  
Stokes Equations ◽  
Suction Side ◽  
Navier Stokes ◽  
Intermittent Flow ◽  
Time Averaging ◽  
Bypass Transition ◽  
Navier Stokes Equations ◽  
Direct Excitation ◽  
Energy Equations

Turbomachinery flows are characterized by a very high intensity turbulent mean part. As a consequence, laminar flow in boundary layer regions undergoes transition through direct excitation of turbulence. This is the so-called bypass transition. Regions form that are intermittently laminar and turbulent. In particular in accelerating flows, as on the suction side of a turbine blade, this intermittent flow can extend over a very large part of the boundary layer. Classical turbulence modelling based on global time averaging is not valid in intermittent flows. To take correctly account of the intermittency, conditioned averages are necessary. These are averages taken during the fraction of time the flow is turbulent or laminar respectively. Starting from the Navier-Stokes equations, conditioned continuity, momentum and energy equations are derived for the laminar and turbulent parts of an intermittent flow. The turbulence is described by the classical k-ε model. The supplementary parameter introduced by the conditioned averaging is the intermittency factor. In the calculations, this factor is prescribed in an algebraic way.

scholarly journals Study and Control of a Radial Vaned Diffuser Stall

2012 ◽  
Vol 2012 ◽  
pp. 1-12 ◽  
Author(s):  
Aurélien Marsan ◽  
Isabelle Trébinjac ◽  
Sylvain Coste ◽  
Gilles Leroy
Keyword(s):  
Boundary Layer ◽  
Skin Friction ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Point Of View ◽  
Vaned Diffuser ◽  
Operating Range ◽  
Stable Operating ◽  
And Control ◽  
Friction Line

The aim of the present study is to evaluate the efficiency of a boundary layer suction technique in case of a centrifugal compressor stage in order to extend its stable operating range. First, an analysis of the flow pattern within the radial vaned diffuser is presented. It highlights the stall of the diffuser vanes when reaching a low massflow. A boundary layer separation in the hub-suction side corner grows when decreasing the massflow from the nominal operating point to the surge and finally leads to a massive stall. An aspiration strategy is investigated in order to control the stall. The suction slot is put in the vicinity of the saddle that originates the main separating skin-friction line, identified thanks to the analysis of the skin-friction pattern. Several aspiration massflow rates are tested, and two different modelings of the aspiration are evaluated. Finally, an efficient control is reached with a removal of only 0,1% of the global massflow and leads—from a steady-state calculations point of view—to an increase by 40% of the compressor operating range extent.

Scaling of Gas Turbine From Air to Refrigerants for Organic Rankine Cycle Using Similarity Concept

2015 ◽  
Vol 138 (6) ◽  
Author(s):  
Choon Seng Wong ◽  
Susan Krumdieck
Keyword(s):  
Gas Turbine ◽  
Test Bench ◽  
Pressure Ratio ◽  
Organic Rankine Cycle ◽  
Rankine Cycle ◽  
Expansion Ratio ◽  
Performance Estimation ◽  
Working Fluids ◽  
Turbine Performance ◽  
Performance Map

Similitude, or similarity concept, is an essential concept in turbomachinery to allow the designer to scale a turbine design to different sizes or different working fluids without repeating the whole design and development process. Similarity concept allows the testing of a turbomachine in a simple air test bench instead of a full-scale organic Rankine cycle (ORC) test bench. The concept can be further applied to adapt an existing gas turbine as an ORC turbine using different working fluids. This paper aims to scale an industrial gas turbine to different working fluids, other than the fluid the turbine was originally designed for. The turbine performance map for air was generated using the 3D computational fluid dynamics (CFD) analysis tools. Three different approaches using the similarity concept were applied to scale the turbine performance map using air and generate the performance map for two refrigerants: R134a and R245fa. The scaled performance curves derived from the air performance data were compared to the performance map generated using CFD analysis tools for R134a and R245fa. The three approaches were compared in terms of the accuracy of the performance estimation, and the most feasible approach was selected. The result shows that complete similarity cannot be achieved for the same turbomachine with two different working fluids, even at the best efficiency point for particular expansion ratio. If the constant pressure ratio is imposed, the location of the optimal velocity ratio and optimal specific speed would be underestimated with calculation error over 20%. Constant Δh0s/a012 was found to provide the highest accuracy in the performance estimation, but the expansion ratio (or pressure ratio) is varying using different working fluids due to the variation of sound speed. The differences in the fluid properties and the expansion ratio lead to the deviation in turbine performance parameters, velocity diagram, turbine's exit swirl angle, and entropy generation. The use of Δh0s/a012 further limits the application of the gas turbine for refrigerants with heavier molecular weight to a pressure ratio less than the designed pressure ratio using air. The specific speed at the best efficiency point was shifted to a higher value if higher expansion ratio was imposed. A correction chart for R245fa was attempted to estimate the turbine's performance at higher expansion ratio as a function of volumetric flow ratio.

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