Thomas Povey

The effect of hot-streaks on HP vane surface and endwall heat transfer : An experimental and numerical study

Pronounced nonuniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of endwall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the effects of these temperature nonuniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimize the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and endwall heat transfer rate for a high-pressure (HP) nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two nonuniform temperature profiles. The temperature profiles were nondimensionally similar to profiles measured in an engine. A difference of one-half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and endwall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualization study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short-duration engine-sized turbine facility. Mach number, Reynolds number, and gas-to-wall temperature ratios were correctly modeled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

Developments in Hot-Streak Simulators for Turbine Testing

The importance of understanding the impact of hot-streaks, and temperature distortion in general, on the high pressure turbine is widely appreciated, although it is still generally the case that turbines are designed for uniform inlet temperature—often the predicted peak gas temperature. This is because there is an insufficiency of reliable experimental data both from operating combustors and from rotating turbine experiments in which a combustor representative inlet temperature profile has accurately been simulated. There is increasing interest, therefore, in experiments that attempt to address this deficiency. Combustor (hot-streak) simulators have been implemented in six rotating turbine test facilities for the study of the effects on turbine life, heat transfer, aerodynamics, blade forcing, and efficiency. Three methods have been used to simulate the temperature profile: (a) the use of foreign gas to simulate the density gradients that arise due to temperature differences, (b) heat exchanger temperature distortion generators, and (c) cold gas injection temperature distortion generators. Since 2004 three significant new temperature distortion generators have been commissioned, and this points to the current interest in the field. The three new distortion generators are very different in design. The generator designs are reviewed, and the temperature profiles that were measured are compared in the context of the available data from combustors, which are also collected. A universally accepted terminology for referring to and quantifying temperature distortion in turbines has so far not developed, and this has led to a certain amount of confusion regarding definitions and terminology, both of which have proliferated. A simple means of comparing profiles is adopted in the paper and is a possible candidate for future use. New whole-field combustor measurements are presented, and the design of an advanced simulator, which has recently been commissioned to simulate both radial and circumferential temperature nonuniformity profiles in the QinetiQ/Oxford Isentropic Light Piston Turbine Test Facility, is presented.

Effect of Simulated Combustor Temperature Nonuniformity on HP Vane and End Wall Heat Transfer: An Experimental and Computational Investigation

This paper presents experimental measurements and computational predictions of surface and end wall heat transfer for a high-pressure (HP) nozzle guide vane operating as part of a full HP turbine stage in an annular rotating turbine facility, with and without inlet temperature distortion (hot streaks). A detailed aerodynamic survey of the vane surface is also presented. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough, UK. This is a short-duration facility, which simulates engine-representative M, Re, nondimensional speed, and gas-to-wall temperature ratio at the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation combustor simulator, capable of simulating well-defined, aggressive temperature profiles in both the radial and circumferential directions. This work forms part of the pan-European research program, TATEF II. Measurements of HP vane and end wall heat transfer obtained with inlet temperature distortion are compared with results for uniform inlet conditions. Steady and unsteady computational fluid dynamics (CFD) predictions have also been conducted on vane and end wall surfaces using the Rolls-Royce CFD code HYDRA to complement the analysis of experimental results. The heat transfer measurements presented in this paper are the first of their kind in that the temperature distortion is representative of an extreme cycle point, and was simulated with good periodicity and with well-defined boundary conditions in the test turbine.

Heat transfer measurements on an intermediate-pressure nozzle guide vane tested in a rotating annular turbine facility, and the modifying effects of a non-uniform inlet temperature profile

Abstract In modern gas turbine engines there exist significant temperature gradients in the combustor exit flow. These gradients arise because both fuel and dilution air are introduced within the combustor as discrete jets. The effects of this non-uniform temperature field on the aerodynamics and heat transfer rate distributions of nozzle guide vanes and turbine blades is difficult to predict, although an increased understanding of the effects of temperature gradients would enhance the accuracy of estimates of turbine component life and efficiency. Low-frequency measurements of heat transfer rate have been conducted on an annular transonic intermediate-pressure (IP) nozzle guide vane operating downstream of a high-pressure (HP) rotating turbine stage. Measurements were conducted with both uniform and non-uniform inlet temperature profiles. The non-uniform temperature profile included both radial and circumferential gradients of temperature. Experiments were conducted in the isentropic light piston facility at QinetiQ Pyestock, a short-duration engine-size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas—wall temperature ratios are correctly modelled. Experimental heat transfer results are compared with predictions performed using boundary layer methods.

Analysis on the Effect of a Nonuniform Inlet Profile on Heat Transfer and Fluid Flow in Turbine Stages

This paper presents an investigation of the aerothermal performance of a modern unshrouded high pressure (HP) aeroengine turbine subject to non-uniform inlet temperature profile. The turbine used for the study was the MT1 turbine installed in the QinetiQ Turbine Test Facility (TTF) based in Farnborough (UK). The MT1 turbine is a full scale transonic HP turbine, and is operated in the test facility at the correct non-dimensional conditions for aerodynamics and heat transfer. Datum experiments of aero-thermal performance were conducted with uniform inlet conditions. Experiments with non-uniform inlet temperature were conducted with a temperature profile that had a non-uniformity in the radial direction defined by (T(max) - T(min))/(T) over bar = 0.355, and a non-uniformity in the circumferential direction defined by (T(max) - T(min))/(T) over bar = 0.14. This corresponds to an extreme point in the engine cycle, in an engine where the non-uniformity is dominated by the radial distribution. Accurate experimental area surveys of the turbine inlet and exit flows were conducted, and detailed heat transfer measurements were obtained on the blade surfaces and end-walls. These results are analysed with the unsteady numerical data obtained using the in-house HybFlow code developed at the University of Firenze. Two particular aspects are highlighted in the discussion: prediction confidence for state of the art computational fluid dynamics (CFD) and impact of real conditions on stator-rotor thermal loading. The efficiency value obtained with the numerical analysis is compared with the experimental data and a 0.8% difference is found and discussed. A study of the flow field influence on the blade thermal load has also been detailed. It is shown that the hot streak migration mainly affects the rotor pressure side from 20% to 70% of the span, where the Nusselt number increases by a factor of 60% with respect to the uniform case. Furthermore, in this work it has been found that a nonuniform temperature distribution is beneficial for the rotor tip, contrary to the results found in the open literature. Although the hot streak is affected by the pressure gradient across the tip gap, the radial profile (which dominates the temperature profile being considered) is not fully mixed out in passing through the HP stage, and contributes significantly to cooling the turbine casing. A design approach not taking into account these effects will underestimate to rotor life near the tip and the thermal load at mid-span. The temperature profile that has been used in both the experiments and CFD is the first simulation of an extreme cycle point (more than twice the magnitude of distortion all previous experimental studies): it represents an engine-take-off condition combined with the full combustor cooling.

A hot-streak (combustor) simulator suited to aerodynamic performance measurements

Abstract A new hot-streak (combustor) simulator has been designed and implemented in a turbine test facility at QinetiQ Farnborough (the QinetiQ Isentropic Light Piston Facility, ILPF) to study the impact of temperature distortion on high pressure (HP) turbine efficiency, aerodynamics, and heat transfer. The ILPF is an engine scale, short duration, rotating transonic turbine test facility, in which M, Re, Tg/Tw, and The hot-streak simulator is a second-generation design, in which cold gas is introduced into a hot mainstream though radial and circumferential slots upstream of the turbine stage. The simulator is rotatable, so the effect of clocking (relative circumferential position of hot streak and nozzle guide vane leading edge) can easily be investigated. An emphasis was placed on accurate measurement of turbine inlet enthalpy flux so that the impact of hot streaks on turbine efficiency could be investigated. The hot-streak simulator differs from all previous systems in that a pronounced radial and circumferential temperature profile has been generated, with a hot-streak to vane count of 1:1. The profile is very well matched (non-dimensionally) to the target profile, which is a combustor temperature profile measured in a modern operating engine at the most extreme point in the cycle. The most accurate area survey of a simulated temperature profile has been conducted to date, and this demonstrates that the simulator offers an exceptionally high degree of circumferential symmetry and run-to-run repeatability. The design and commissioning of the simulator is described, and the measured temperature profiles are compared with the target profile.

A combustor-representative swirl simulator for a transonic turbine research facility

Tighter aircraft emissions regulations have let to considerable improvement in gas turbine combustion in the past few decades. Modern combustors employ aggressive swirlers to increase mixing and to improve flame stability during the combustion process. The flow at combustor exit can therefore have high residual swirl. The impact of this swirl on the aerodynamic and heat transfer characteristics of the HP turbine stage has not yet received much attention. In order to investigate the effects of swirl on the HP turbine stage, an inlet swirl simulator has been designed and commissioned in an engine scale, short duration, rotating transonic turbine facility. The test facility simulates engine representative Mach number, Reynolds number, non-dimensional speed and gas-to-wall temperature ratio at the turbine inlet. The target swirl profile at turbine stage inlet was based upon extreme exit swirl conditions for a modern low-NO x combustor with peak yaw and pitch angles over ±40°. A number of candidate swirler designs were considered during a pilot study that was conducted in a subsonic wind tunnel to achieve suitable swirler design. The swirl simulator was developed based upon the pilot study results, which achieved a good match to the target profile after commissioning in the facility. This article mainly deals with the design and development of the swirl generator. It presents the experimental and computational results of the pilot study, followed by the description of the installation and commissioning of the swirl simulator on the test facility. Novel instrumentation was required to survey the swirl profile, which is also described. A comparison of the measured and computational aerodynamic results with and without swirl, at 10 per cent and 90 per cent span of HP nozzle guide vane is also presented. The comparison highlights significant impact of swirl on the vane incidence angle, and therefore a considerable affect on the loading distribution of the vane.

Turbine Efficiency Measurement System for the QinetiQ Turbine Test Facility

A turbine efficiency measurement system has been developed and installed on the turbine test facility (TTF) at QinetiQ Farnborough. The TTF is an engine-scale short-duration (0.5s run time) rotating transonic facility, which can operate as either single stage (HP vane and rotor) or 112 stage (HP stage with IP or LP vane). The current MT1 HP stage is highly loaded and unshrouded and is therefore relevant to current design trends. Implementation of the efficiency measurement system forms part of the EU Turbine Aero-Thermal External Flows (TATEF II) program. The following aspects of the efficiency measurement system are discussed in this paper: mass-flow rate measurement, power measurement by direct torque measurement, turbine inlet and exit area traverse measurement systems, computation of efficiency by mass weighting, and uncertainty analysis of the experimentally determined turbine efficiency. The calibration of the mass-flow rate and torque measurement systems are also discussed. Emphasis was placed on the need for a low efficiency precision uncertainty, so that changes in efficiency associated with turbine inlet temperature distortion and swirl can be resolved with good accuracy. Measurements with inlet flow distortion form part of the TATEF II program and will be the subject of forthcoming publications.

On a novel annular sector cascade technique

An advanced technique for establishing pressure boundary conditions in annular sector cascade experiments has been developed. This novel technique represents an improvement over previous methods and provides the first means by which annular sector boundary conditions that are representative of those which develop in an annular cascade can be established with a high degree of satisfaction. The technique will enable cascade designers to exploit the obvious advantages of annular sector cascade testing: the reduced cost of both facility manufacture and facility operation and the use of engine parts in place of two-dimensional counterparts. By employing an annular sector of deswirl vanes downstream of the annular sector of test vanes, the radial pressure gradient established in the swirling flow downstream of the test vanes is not disturbed. The deswirl vane exit flow-which has zero swirl velocity-can be exhausted without unsteadiness, and without the risk of separation, into a plenum at constant pressure. The pressure ratio across the annular sector of test vanes can be tuned by adjusting the throat area at the deswirl vane exit plane. Flow conditioning systems which utilize the Oxford deswirl vane technology have previously been used to set pressure boundary conditions downstream of fully annular cascades in both model and engine scale (the Isentropic Light Piston Facility at Farnhorough) experimental research facilities (Povey, T., Chana, K. S., Oldfield, M. L. G., Jones, T. V., and Owen, A. K., 2001, Proceedings of the ImechE Advances in Fluid Machinery Design Seminar, London, June 13; Povey, T., Chana, K. S., Jones, T. V., and Oldfield, M. L. G., 2003, Advances of CFD in Fluid Machinery Design, ImechE Professional Engineering, London, pp. 65-94). The deswirl vane is particularly suited to the control of highly whirling transonic flows. It has been demonstrated by direct comparison of aerodynamic measurements from fully annular and annular sector experiments that the use of a deswirl vane sector for flow conditioning at the exit of an annular sector cascade represents an attractive novel solution to the boundary condition problem. The annular sector technique is now described.

New technique for the fabrication of miniature thin film heat flux gauges

This paper details the improvements made to the design and fabrication of thin-film heat flux gauges at Oxford. These improvements have been driven by the desire to improve measurement accuracy and resolution in short duration wind-tunnel experiments.A thin-film heat flux gauge (TFHFG) measures heat flux by recording the temperature history of thin film resistive temperature sensors sputtered onto an insulating substrate. The heat flux can then be calculated using Fourier’s law of heat conduction.A new fabrication process utilising technology from the manufacture of flexible printed circuit boards is outlined, which enables the production of significantly smaller and more robust gauges than those previously used.