Pressure and Suction Surfaces Redesign for High Lift Low Pressure Turbines

Nowadays there is a big effort toward improving the low pressure turbine efficiency even to the extent of penalising other relevant design parameters. LP turbine efficiency influences SFC more than other modules in the engine. Most of the research has been oriented to reduce profile losses, modifying the suction surface, the pressure surface or the three-dimensional regions of the flow. To date, the pressure surface has received very little attention. The dependence of the profile losses on the behaviour of both pressure and suction surfaces has been investigated for the case of a high lift design that is representative of a modern civil engine LP turbine. The experimental work described in this paper consists on two different sets of experiments: the first one concluded an improved pressure surface definition and the second set was oriented to achieve further improvement in losses modifying the profile suction surface. Three profiles were designed and tested over a range of conditions. The first profile is a thin-solid design. This profile has a large pressure side separation bubble extending from near the leading edge to mid-chord. The second profile is a hollow design with the same suction surface as the first one but avoiding pressure surface separation. The third one is also a hollow design with the same pressure surface as the second profile but more aft loaded suction surface. The study is part of a wider on-going research programme covering the effects of the different design parameters on losses. The paper describes the experiments conducted in a low-speed linear cascade facility. It gathers together steady and unsteady loss measurements by wake traverse and surface pressure distributions for all the profiles. It is shown that thick profiles generate only around 90% of the losses of a thin-solid profile with the same suction surface. The results support the idea of an optimum axial position for the peak Mach number. Caution is recommended as profile aft loading would not be a completely secure method for reducing losses.

Design, Fabrication, and Testing of a Semi-Closed Cycle Gas Turbine Engine

A small semi-closed gas turbine was designed, fabricated, and tested to demonstrate the cycle the cycle feasibility with exhaust gas recirculation. The demonstrator unit comprised a low pressure spool compressor and turbine supercharging a high pressure spool compressor and turbine, whose exhaust passed through a recuperator, and was subsequently split, one half being recirculated to the high pressure spool compressor inlet via an intercooler, and the remaining half expanded across the low pressure spool turbine. The design and fabrication phases proceeded on schedule but commencement of engine development testing encountered mechanical difficulties. These were eventually resolved and shakedown testing of the demonstrator accomplished prior to final contractual delivery. The demonstration program was funded under a NASA LeRc contract NAS3-27396.

Investigation of Flow in a Steam Turbine Exhaust Hood With/Without Turbine Exit Conditions Simulated

Experimental and numerical investigations for the flow in an exhaust hood model of large steam turbines have been carried out in order to understand the complex 3D flow pattern existing in the hood and also to validate the CFD solver. The model is a typical design for 300/600 MW steam turbines currently in operation. Static pressure at the diffuser tip and hub endwalls and at hood outer casing is measured and nonuniform circumferential distributions of static pressure are noticed. Velocity field at the model exit is measured and compared with the numerical prediction. The multigrid multiblock 3D Navier-Stokes solver used for the simulations is based upon the TVD Lax-Wendroff scheme and the Baldwin-Lomax turbulence model. Good agreement between numerical results and experimental data is demonstrated. It is found that the flow pattern and hood performance are very different with or without the turbine exit flow conditions simulated.

Development of Fault Diagnosis and Failure Prediction Techniques for Small Gas Turbine Engines

A computer model of a gas turbine auxiliary power unit was produced to develop techniques for fault diagnosis and prediction of remaining life in small gas turbine engines. Due to the relatively low capital cost of small engines it is important that the techniques have both low capital and operating costs. Failing engine components were identified with fault maps, and an algorithm was developed for predicting the time to failure, based on the engine’s past operation. Simulating daily engine operation over a maintenance cycle tested the techniques for identification and prediction. The simulation included daily variations in ambient conditions, operating time, load, engine speed and operating environment, to determine the amount of degradation per day. The algorithm successfully adapted to the daily changes and corrected the operating point back to standard conditions to predict the time to failure.

Midspan Flow-Field Measurements for Two Transonic Linear Turbine Cascades at Off-Design Conditions

The paper presents detailed mid-span experimental results from two transonic linear turbine cascades. The blades for the two cascades were designed for the same service and differ mainly in their leading-edge geometries. One of the goals of the study was investigate the influence of the leading-edge metal angle on the sensitivity of the blade to positive off-design incidence. Measurements were made for incidence values of −10.0°, 0.0°, +4.5°, +10.0°, and +14.5° relative to design incidence. The exit Mach numbers varied roughly from 0.5 to 1.2 and the Reynolds numbers from about 4×105 to 106. The measurements include the midspan losses, blade loadings and base pressures. In addition, the axial-velocity-density ratio (AVDR) was extracted for each operating point The AVDR was found to vary from about 0.98 at −10.0° of incidence to about 1.27 at +14.5°. Thus, the data set also provides some evidence of the influence AVDR on axial turbine blade performance. Detailed experimental results for turbine blade performance at off-design incidence are very scarce in the open literature, particularly for transonic conditions. Among other things, the present results are intended to expand the database available in the open literature. To this end, the key aerodynamic results are presented in tabular form, along with the detailed geometry of the cascades. The results could be used in the development of new or improved correlations for use in the early stages of design. They could also be used to evaluate the ability of current CFD codes to capture reliably the variation in losses and other aerodynamic quantities with variations in blade incidence.

Non-Intrusive Temperature Measurements in a Transonic Axial Compressor Using Spontaneous Rotational Raman Spectroscopy

Spontaneous rotational Raman spectroscopy was used to measure static temperature in a transonic axial compressor. This laser technique has the general advantages of such techniques, i.e. possibility of performing in-situ measurements with high spatial resolution and temporal resolution, and non-intrusiveness. The experimental investigations were performed just downstream the rotor for the design mass flow rate. The results are obtained from Raman scattering of air by using a high repetition rate pulsed Nd:YAG laser and a spectrally and spatially resolving detector system that consists of a triple spectrograph and a CCD camera. Raman Spectra presented here are integrated over 20 s. Profiles of temperature at various locations relative to the casing are presented and the measured temperatures are compared with theoretical values. Of great importance for the results is the accuracy of the technique, and a sensitivity analysis is performed to test the temperature on uncertainties in experimental parameters. It was shown that the absolute accuracy of temperature could be as low as ± 1 K in the 300 - 400 K operating range of interest.

Evaluation of Unsteady CFD Methods by Their Application to a Transonic Propfan Stage

The accuracy in predicting the unsteady aerodynamic blade-row-interaction of two state-of-the-art Navier-Stokes codes is evaluated within the current paper. The general flow features of the test case — a transonic research propfan stage — are described in brief as far as necessary to understand the detailed comparisons. The calculated unsteady velocity and flow angle distributions at various axial planes of the stage are compared to data from unsteady laser measurements. The general flow features of the propfan are very well reproduced by the numerical methods and a good agreement is also obtained in comparison to the measured data. One important outcome of the comparison is the good agreement of both numerical methods with the unsteady fluctuations measured in the experiment.

Improving the Efficiency of the Trent 500 HP Turbine Using Non-Axisymmetric End Walls: Part II — Experimental Validation

Part I of this paper described how the HP turbine model rig of the Rolls-Royce Trent 500 was redesigned by applying non-axisymmetric end walls to both the vane and blade passages, whilst leaving the turbine operating point and overall flow conditions unaltered. This paper describes the results obtained from testing of the model rig and compares them with those obtained for the datum design (with conventional axisymmetric end walls). Measured improvements in the turbine efficiency are shown to be in line with those expected from the previous linear cascade research at Durham University, see Harvey et al. [1] and Hartland et al. [2]. These improvements are observed at both design and off-design conditions. Hot wire traverses taken at the exit of the rotor show, unexpectedly, that the end wall profiling has caused changes across the whole of the turbine flow field. This result is discussed making reference to a preliminary 3-D CFD analysis. It is concluded that the design methodology described in part I of this paper has been validated, and that non-axisymmetric end wall profiling is now a major new tool for the reduction of secondary loss in turbines (and potentially all axial flow turbomachinery). Further work, though, is needed to fully understand the stage (and multistage) effects of end wall profiling.

Numerical Investigation of Airfoil Clocking in a Three-Stage Low Pressure Turbine

A quasi–three–dimensional, blade–to–blade, time–accurate, viscous solver w as used for a three–stage LP turbine study Due to the low Reynolds number, transitional computations were performed. Unsteady analyses were then carried out by varying the circumferential relative position of consecutive vanes and blade rows to study the effects of clocking on the turbine’s performance. A clocking strategy developed in order to limit the number of configurations to be analyzed is discussed. The optimum analytically–determined clocking position is illustrated for two different operating conditions, referred to as cruise and takeoff. The effects of clocking on wake interaction mechanisms and unsteady blade loadings is presented and discussed. For low Reynolds number turbine flows, the importance of taking transition into account in clocking analysis is demonstrated by a comparison with a fully turbulent approach.

An Experimental and Numerical Investigation Into the Mechanisms of Rotating Instability

This paper reports on an experimental and numerical investigation aimed at understanding the mechanisms of rotating instabilities in a low speed axial flow compressor. The phenomena of rotating instabilities in the current compressor were first identified with an experimental study. Then, an unsteady numerical method was applied to confirm the phenomena and to interrogate the physical mechanisms behind them. The experimental study was conducted with high-resolution pressure measurements at different clearances, employing a double phase-averaging technique. The numerical investigation was performed with an unsteady 3-D Navier-Stokes method that solves for the entire blade row. The current study reveals that a vortex structure forms near the leading edge plane. This vortex is the result of interactions among the classical tip-clearance flow, axially reversed endwall flow, and the incoming flow. The vortex travels from the suction side to the pressure side of the passage at roughly half of the rotor speed. The formation and movement of this vortex seem to be the main causes of unsteadiness when rotating instability develops. Due to the nature of this vortex, the classical tip-clearance flow does not spill over into the following blade passage. This behavior of the tip-clearance flow is why the compressor operates in a stable mode even with the rotating instability, unlike traditional rotating stall phenomena.