Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design.
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.
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.
Purpose To derive lobar ventilation in patients with chronic obstructive pulmonary disease (COPD) using a rapid time-series hyperpolarized xenon-129 (HPX) magnetic resonance imaging (MRI) technique and compare this to ventilation/perfusion single-photon emission computed tomography (V/Q-SPECT), correlating the results with high-resolution computed tomography (CT) and pulmonary function tests (PFTs). Materials and methods Twelve COPD subjects (GOLD stages I–IV) participated in this study and underwent HPX-MRI, V/Q-SPECT/CT, high-resolution CT, and PFTs. HPX-MRI was performed using a novel time-series spiral k-space sampling approach. Relative percentage ventilations were calculated for individual lobe for comparison to the relative SPECT lobar ventilation and perfusion. The absolute HPX-MRI percentage ventilation in each lobe was compared to the absolute CT percentage emphysema score calculated using a signal threshold method. Pearson’s correlation and linear regression tests were performed to compare each imaging modality. Results Strong correlations were found between the relative lobar percentage ventilation with HPX-MRI and percentage ventilation SPECT ( r = 0.644; p < 0.001) and percentage perfusion SPECT ( r = 0.767; p < 0.001). The absolute CT percentage emphysema and HPX percentage ventilation correlation was also statistically significant ( r = 0.695, p < 0.001). The whole lung HPX percentage ventilation correlated with the PFT measurements (FEV 1 with r = − 0.886, p < 0.001*, and FEV 1 /FVC with r = − 0.861, p < 0.001*) better than the whole lung CT percentage emphysema score (FEV 1 with r = − 0.635, p = 0.027; and FEV 1 /FVC with r = − 0.652, p = 0.021). Conclusion Lobar ventilation with HPX-MRI showed a strong correlation with lobar ventilation and perfusion measurements derived from SPECT/CT, and is better than the emphysema score obtained with high-resolution CT. Key Points • The ventilation hyperpolarized xenon-129 MRI correlates well with ventilation and perfusion with SPECT/CT with the advantage of higher temporal and spatial resolution. • The hyperpolarized xenon-129 MRI correlates with the PFT measurements better than the high-resolution CT with the advantage of avoiding the use of ionizing radiation. Electronic supplementary material The online version of this article (10.1007/s00330-018-5888-y) contains supplementary material,...
This paper presents an investigation of the aerothermal performance of a modern unshrouded high-pressure (HP) aero-engine turbine subject to nonuniform inlet temperature profile. The turbine used for this study was the MT1 turbine installed in the QinetiQ turbine test facility based in Farnborough (UK). The MT1 turbine is a full scale transonic HP turbine, and is operated in the test facility at the correct nondimensional conditions for aerodynamics and heat transfer. Datum experiments of aerothermal performance were conducted with uniform inlet conditions. Experiments with nonuniform inlet temperature were conducted with a temperature profile that had a nonuniformity in the radial direction defined by (Tmax−Tmin)/T¯=0.355, and a nonuniformity in the circumferential direction defined by (Tmax−Tmin)/T¯=0.14. This corresponds to an extreme point in the engine cycle, in an engine where the nonuniformity 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 analyzed 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 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 the rotor life near the tip and the thermal load at midspan. The temperature profile that has been used in both experiments and CFD is the first simulation of an extreme cycle point (more than twice the magnitude of distortion of all previous experimental studies): It represents an engine-take-off condition combined with the full combustor cooling. This research was part of the EU funded Turbine AeroThermal External Flows 2 program.
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