Measurement of Unsteady Flow and Heat Transfer in a Linear Turbine Cascade

A detailed experimental study has been conducted on the wake-induced unsteady flow and heat transfer in a linear turbine cascade. The unsteady wakes with passing frequencies in the range zero to 240 Hz were generated by moving cylinders on a squirrel cage device. The velocity fields in the blade-to-blade flow and in the boundary layers were measured with hot-wire anemometers, the surface pressures with a pressure transducer and the heat transfer coefficients with a glue-on hot film. The results were obtained in ensemble-averaged form so that periodic unsteady processes can be studied. Of particular interest was the transition of the boundary layer. The boundary layer remained laminar on the pressure side in all cases and in the case without wakes also on the suction side. On the latter, the wakes generated by the moving cylinders caused transition, and the beginning of transition moves forward as the cylinder-passing frequency increases. Unlike in the flat-plate study of Liu and Rodi (1991a) the instantaneous boundary layer state does not respond to the passing wakes and therefore does not vary with time. The heat transfer increases under increasing cylinder-passing frequency even in the regions with laminar boundary layers due to the increased background turbulence.

An Investigation of Turbulence Modelling in Transonic Fans Including a Novel Implementation of an Implicit κ–ϵ Turbulence Model

An explicit Navier-Stokes solver has been written with the option of using one of two types of turbulence models. One is the Baldwin-Lomax algebraic model and the other is an implicit k-ϵ model which has been coupled with the explicit Navier-Stokes solver in a novel way. This type of coupling, which uses two different solution methods, is unique and combines the overall robustness of the implicit k-ϵ solver with the simplicity of the explicit solver. The resulting code has been applied to the solution of the flow in a transonic fan rotor which has been experimentally investigated by Wennerstrom. Five separate solutions, each identical except for the turbulence modelling details, have been obtained and compared with the experimental results. The five different turbulence models run were: the standard Baldwin-Lomax model both with and without wall functions, the Baldwin-Lomax model with modified constants and wall functions, a standard k-ϵ model and an extended k-ϵ model which accounts for multiple time scales by adding an extra term to the dissipation equation. In general, as the model includes more of the physics, the computed shock position becomes closer to the experimental results.

Recess Vane Passive Stall Control

An experimental investigation has been carried out on the influence of a vaned recessed casing treatment on the stall margin improvement of axial flow fans with different hub to tip ratio, with and without inlet distortion. The inlet distortion tests were conducted on a 0.5 hub to tip ratio fan and significant increases in the flow range with only small drops in operating efficiency were observed. The clean flow tests were conducted on higher hub to tip ratio fans (0.7 and 0.9). In each case the stage characteristic was compared with the results obtained with a solid casing. Significant increases in the flow range, with only modest or no loss in operating efficiency, were observed for optimum configurations at both diameter ratios.

Compressible Flow Pressure Losses in Wye-Junctions

This paper considers the compressible flow pressure losses in sharp cornered wye-junctions with symmetrical branches under dividing and combining flow conditions. Determination of the additional total pressure losses occurring in flow through several three-leg junctions, using dry air as the working fluid, has been made experimentally. Results covering a wide speed range up to choking are presented for three different wye-junction geometries. Separate flow visualisation Schlieren tests detected the presence of normal shock waves, located at up to one duct diameter downstream of the junction, and therefore confirmed the choking of the flow at the vena contracta. The highest attainable Mach number (M3) of the averaged whole flow was 0.9 for one of the dividing flow geometries and 0.65 for several of the combining flow cases. These values of M3 were the maximum possible and hence represent a limiting condition dictated by choking. In general, the compressible flow loss coefficients, caused by the presence of the wye-junctions, can be expected to be higher for dividing flows and lower for combining flows than would be the case for incompressible flows because of the influence of Mach number (M3) on the magnitude of the denominator.

Transition Prediction in Attached and Separated Shear Layers Using an Integral Method

Calculation of the aerodynamic parameters of axial turbomachinery blades, and an accurate assessment of the flow over the blade surfaces under today’s increasingly demanding requirements for higher efficiencies and optimized blade shapes, at both design and off-design conditions, impose a need for accurate prediction methods able to compute through two sensitive but highly critical phenomena: separation and transition. The present study describes work done on the modelling and prediction of transitional regions, such as those appearing on turbomachinery blading, covering both attached and separated flows. The concept of an engineering method, cheap to run and avoiding complex CFD and turbulence model formulations was always kept in mind. Results include comparisons of integral quantities and velocity profiles in zero, favourable or adverse pressure gradient attached flows, and velocity distributions including points of separation, transition and reattachment in separated airfoil flows, obtained either from a straightforward shear layer calculation or from a viscous-inviscid interaction procedure.

Radial Transport and Momentum Exchange in an Axial Compressor

The objective of this work was to examine radial transport in axial compressors from two perspectives. The first was to compare the mixing coefficient based on a secondary flow model (using measured radial velocities) with that based on a turbulent diffusion model. The second was to use measured airfoil pressure forces and momentum changes to assess the validity of the assumption of diffusive radial transport which is common to both models. These examinations were carried out at both design and off-design conditions as well as for two rotor tip clearances. In general it was seen that radial mixing was strongest near the hub and that it increased dramatically at near-stall conditions. It was also seen that radial transport could cause large differences (≈ 100%) between the force on an airfoil and the change in momentum across the airfoil at the same spanwise location.

The Design and Evaluation of a High Pressure Ratio Radial Turbine

A single stage, high speed, high pressure ratio radial inflow turbine was designed for a single shaft gas turbine engine in the 200 kW power range. A model turbine has been tested in a cold rig facility with correct simulation of the important non-dimensional parameters. Performance measurements over a wide range of operation were made, together with extensive volute and exhaust traverses, so that gas velocities and incidence and deviation angles could be deduced. The turbine efficiency was lower than expected at all but the lowest speed. The rotor incidence and exit swirl angles, as obtained from the rig test data, were very similar to the design assumptions. However, evidence was found of a region of separation in the nozzle vane passages, presumably caused by a very high curvature in the endwall just upstream of the vane leading edges. The effects of such a separation are shown to be consistent with the observed performance.

Incidence Angle and Pitch-Chord Effects on Secondary Flows Downstream of a Turbine Cascade

This paper describes the results of an experimental investigation of the three-dimensional flow downstream of a linear turbine cascade at off-design conditions. The tests have been carried out for five incidence angles from −60 to +35 degrees, and for three pitch-chord ratios: s/c = 0.58,0.73,0.87. Data include blade pressure distributions, oil flow visualizations, and pressure probe measurements. The secondary flow field has been obtained by traversing a miniature five hole probe in a plane located at 50% of an axial chord downstream of the trailing edge. The distributions of local energy loss coefficients, together with vorticity and secondary velocity plots show in detail how much the secondary flow field is modified both by incidence and cascade solidity variations. The level of secondary vorticity and the intensity of the crossflow at the endwall have been found to be strictly related to the blade loading occurring in the blade entrance region. Heavy changes occur in the spanwise distributions of the pitch averaged loss and of the deviation angle, when incidence or pitch-chord ratio is varied.

A Novel Approach to High Resolution Compressible Cascade Flow Analysis Using the Navier-Stokes Equations

This paper deals with the development of a technique aimed at improving the accuracy of 2-D flow solutions of turbomachinery problems. The basic concept is to take a quasi-3D Navier-Stokes or Euler solver on a coarse mesh (the “outer code”) and couple it to a 2-D space marching parabolised Navier-Stokes solver on a finer sub-mesh (the “inner code”). The “inner-code” includes the FLARE approximation to permit reverse flow. The inner and outer codes are coupled by adopting an approach analogous to classical multigrid methods. The combination forms a cheap and fast solver to provide fine resolution solutions using only mini-computer resources. Predictions of the flow through a compressor and a turbine cascade are described and show good agreement with the experimental results.

Recent Advances in Simulating Unsteady Flow Phenomena Brought About by Passage of Shock Waves in a Linear Turbine Cascade

High-pressure ratio turbines have flows dominated by shock structures that pass downstream into the next blade row in an unsteady fashion. Recent numerical results have indicated that these unsteady shocks may significantly affect the aerodynamic and mechanical performance of turbine blading. High cost and limited accessibility of turbine rotating equipment severely restrict the quantitative evaluation of the unsteady flowfield in that environment. Recently published results of the Virginia Tech transonic cascade facility indicate high integrity in simulation of the steady state flow phenomena. The facility has recently been modified to study the unsteady effects of passing shock waves. Shock waves are genarated by a shotgun blast upstream of the blade row. Shadowgraph photos and high-response pressure data are compared to previously published experimental and numerically predicted results. Plots are included which indicate large fluctuations in estimated blade lift and cascade loss.