Vortex Transport and Blade Interactions in High Pressure Turbines

This paper presents a study of the three-dimensional flow field within the blade rows of a high-pressure axial flow steam turbine stage. Half-delta wings were fixed to a rotating hub to simulate an upstream rotor passage vortex. The flow field is investigated in a Low-Speed Research Turbine using pneumatic and hot-wire probes downstream of the blade row. The paper examines the impact of the delta wing vortex transport on the performance of the downstream blade row. Steady and unsteady numerical simulations were performed using structured 3D Navier-Stokes solver to further understand the flow field. The loss measurements at the exit of the stator blade showed an increase in stagnation pressure loss due to the delta wing vortex transport. The increase in loss was 21% of the datum stator loss, demonstrating the importance of this vortex interaction. The transport of the stator viscous flow through the rotor blade row is also described. The rotor exit flow was affected by the interaction between the enhanced stator passage vortex and the rotor blade row. Flow underturning near the hub and overturning towards the mid-span was observed, contrary to the classical model of overturning near the hub and underturning towards the mid-span. The unsteady numerical simulation results were further analysed to identify the entropy producing regions in the unsteady flow field.

The effects of a trip wire and unsteadiness on a high speed highly loaded low-pressure turbine blade

This paper presents the effect of a single spanwise 2D wire upon the downstream position of boundary layer transition under steady and unsteady inflow conditions. The study is carried out on a high turning, high-speed, low pressure turbine (LPT) profile designed to take account of the unsteady flow conditions. The experiments were carried out in a transonic cascade wind tunnel to which a rotating bar system had been added. The range of Reynolds and Mach numbers studied includes realistic LPT engine conditions and extends up to the transonic regime. Losses are measured to quantify the influence of the roughness with and without wake passing. Time resolved measurements such as hot wire boundary layer surveys and surface unsteady pressure are used to explain the state of the boundary layer. The results suggest that the effect of roughness on boundary layer transition is a stability governed phenomena, even at high Mach numbers. The combination of the effect of the roughness elements with the inviscid Kelvin-Helmholtz instability responsible for the rolling up of the separated shear layer (Stieger [1]) is also examined. Wake traverses using pneumatic probes downstream of the cascade reveal that the use of roughness elements reduces the profile losses up to exit Mach numbers of 0.8. This occurs with both steady and unsteady inflow conditions.

Development of blade profiles for low pressure turbine applications

This paper describes a program of work, largely experimental, which was undertaken with the objective of developing an improved blade profile for the low-pressure turbine in aero-engine applications. Preliminary experiments were conducted using a novel technique. An existing cascade of datum blades was modified to enable the pressure distribution on the suction surface of one of the blades to be altered. Various means, such as shaped inserts, an adjustable flap at the trailing edge, and changing stagger were employed to change the geometry of the passage. These experiments provided boundary layer and lift data for a wide range of suction surface pressure distributions. The data was then used as a guide for the development of new blade profiles. The new blade profiles were then investigated in a low-speed cascade that included a set of moving bars upstream of the cascade of blades to simulate the effect of the incoming wakes from the previous blade row in a multistage turbine environment.

A new alternative for reduction of secondary flows in low pressure turbines

This paper describes a new flow mechanism for the reduction of secondary flows in Low Pressure Turbines using the benefit of contoured endwalls. The extensive application of contoured endwalls in recent years has provided a deeper understanding of the physical phenomenon that governs the reduction of secondary flows. Based on this understanding, the endwall geometry of a linear cascade of solid-thin profiles typical of Low Pressure Turbines has been redesigned. Experimental data are presented for the validation of this new solution. Based on these data, a reduction of 72% in the SKEH and 20% in the mixed-out endwall losses can be obtained. CFD simulations are also presented to illustrate the effect of the new endwall on the secondary flows. Furthermore, an explanation of the flow mechanism that governs the reduction of the SKEH and the losses is given. Copyright © 2006 by ASME.

The prediction of hot streak migration in a high-pressure turbine

Accurate predictions of combustor hot streak migration enable the turbine designer to identify high-temperature regions that can limit component life. It is therefore important that these predictions are achieved within the short time scales of a design process. This article compares temperature measurements of a circular hot streak through a turning duct and a research turbine with predictions using a three-dimensional Reynolds-averaged Navier-Stokes solver. It was found that the mixing length turbulence model did not predict the hot streak dissipation accurately. However, implementation of a very simple model of the free stream turbulence (FST) significantly improved the exit temperature predictions on both the duct and research turbine. One advantage of the simple FST model described over more complex alternatives is that no additional equations are solved. This makes the method attractive for design purposes, as it is not associated with any increase in computational time.

Velocity distributions for low pressure turbines

A parametric set of velocity distributions has been investigated using a flat plate experiment. Three different diffusion factors and peak velocity locations were tested. These were designed to mimic the suction surfaces of Low Pressure (LP) turbine blades. Unsteady wakes, inherent in real turbomachinery flows, were generated using a moving bar mechanism. A turbulence grid generated a freestream turbulence level that is believed to be typical of LP turbines. Measurements were taken across a Reynolds number range of 50,000-220,000 at three reduced frequencies (0.314, 0.628, 0.942). Boundary layer traverses were performed at the nominal trailing edge using a Laser Doppler Anemometry system and hot-films were used to examine the boundary layer behaviour along the surface. For every velocity distribution tested, the boundary layer separated in the diffusing flow downstream of the peak velocity. The loss production is dominated by the mixing in the reattachment process, mixing in the turbulent boundary layer downstream of reattachment and the effects of the unsteady interaction between the wakes and the boundary layer. A sensitive balance governs the optimal location of peak velocity on the surface. Moving the velocity peak forwards on the blade was found to be increasingly beneficial when bubblegenerated losses are high, i.e. at low Reynolds number, at low reduced frequency and at high levels of diffusion. Copyright © 2008 by ASME.

Fluid mechanics of pulse pressure-gain combustors

The fluid dynamic operation of a valveless pulse combustor has been studied experimentally and numerically. Through phase-locked chemiluminescence and pressure measurements it is shown that mechanical energy is created periodically in the flame surface, with an efficiency of 1.6%. This mechanical energy leaves the pulse combustor through unsteady jets at the aerovalve inlet and the tailpipe exit stations. Two thermodynamically distinct flows are identified: a flow that is transported from inlet to exit and participates in combustion along the way, and a flow that is ingested and then ejected from the combustor without undergoing combustion. It is the latter of these two flows which has the greatest quantity of net work done on it. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

Performance of choked unsteady ejector-nozzles for use in pressure-gain combustors

If the conventional steady flow combustor of a gas turbine is replaced with a device which achieves a pressure gain during the combustion process then the thermal efficiency of the cycle is raised. All such 'Pressure Gain Combustors' (e.g. PDEs, pulse combustors or wave rotors) are inherently unsteady flow devices. For such a device to be practically installed in a gas turbine it is necessary to design a downstream row of turbine vanes which will both accept the combustors unsteady exit flow and deliver a flow which the turbine rotor can accept. The design requirements of such a vane are that its exit flow both retains the maximum time-mean stagnation pressure gain (the pressure gain produced by the combustor is not lost) and minimises the amplitude of unsteadiness (reduces unsteadiness entering the downstream rotor). In this paper the exit of the pressure gain combustor is simulated with a cold unsteady jet. The first stage vane is simulated by a one-dimensional choked ejector nozzle with no turning. The time-mean and rms stagnation pressure at nozzle exit is measured. A number of geometric configurations are investigated and it is shown that the optimal geometry both maximizes time mean stagnation pressure gain (75% of that in the exit of the unsteady jet) and minimizes the amplitude of unsteadiness (1/3 of that in the primary jet). The structure of the unsteady flow within the ejector nozzle is determined computationally. Copyright © 2009 by J Heffer and R Miller.

A transonic vane for use downstream of a pressure gain combustor

Replacing a conventional combustor in a gas turbine with one that produces a pressure gain could significantly raise cycle efficiency. For this efficiency gain to be achieved the exit flow from the combustor must be coupled to the downstream turbine such that the pressure gain produced by the combustor is retained and such that the turbine efficiency is maintained. The exit flow from a pressure gain combustor will often contain a high velocity unsteady jet. It has previously been proposed that ejectors should be used to harness the energy in the unsteady jet, this paper proposes combining an ejector with the first stage vane, producing a single compact component that preserves the combustion driven pressure gain and delivers a suitable flow to the turbine so that its efficiency is not compromised. This novel component has been experimentally tested for the first time. The performance of this first prototype design is found to be low due to high levels of loss generated by secondary flows. However possible mitigation strategies are discussed. It is shown that the unsteadiness at exit form the ejector-vane is reduced compared to the inlet flow. If a pulse combustor were incorporated into a gas turbine, it is unlikely that the level of unsteadiness experienced in a downstream rotor will be significantly larger that that due to the periodic passing of upstream wakes. Copyright © 2010 by Jonathan Heffer.

Mandrel peeling of a flexible laminate with a pressure-sensitive adhesive: Comparison of experiment with numerical analysis

Mandrel peel tests with mandrels or rollers of varying diameters have been carried out using Mylar backing of several thicknesses and a commercial synthetic acrylic adhesive. The results are critically compared with the numerical predictions of the peeling software package ICPeel. In addition, a finite element model of the mandrel peeling process has been completed which gives good agreement with experiment provided appropriate mechanical properties of adherend and adhesive are used which must include the effects of adherent constraint. The influence of the thickness of the backing is also considered and both experiment and analysis confirm that there is a backing thickness at which the peel force for a laminate of this sort will show a maximum. © 2010 Blackwell Publishing Ltd.