Labyrinth seals are extensively used in turbomachinery to prevent high pressure gas from flowing into a region of low pressure. Because of thermal expansions and centrifugal forces, the actual seal clearance can vary based on engine conditions. Pressure ratio, Reynolds number, tip geometry, and seal clearance all affected the sealing performance. This paper deals with its influence on the leakage flow and heat transfer coefficient through a thirteen teeth straight through labyrinth seal. Three gaps were experimentally investigated using a stationary test rig. The experiments covered a range of Reynolds numbers between 5000 and 50000 and pressure ratios between 1.0 and 2.7. Cavity pressure measurements along the seal were also performed in order to characterize each constriction. In addition, 2D PIV measurements were made on the plane containing the seal teeth to obtain a high local resolution of the velocity distribution and the flow field within the seal. Experimental results show a strong influence of clearance on both leakage loss and heat transfer as well as on the development of the flow fields. A simplified model to calculate the leakage mass flow rate is presented and validated comparing its prediction capability with experimental data. In order to improve the agreement between numerical and experimental results a correction of published correlations is proposed.
Short pin fins (pedestals) and long ribs (enlarged pedestals) are usually used in trailing edge cooling of turbine airfoils. To better reproduce the geometry of such cooling systems, test section of this experimental study is a wedge duct with inserts, to take into account effects of accelerating flow. A complete investigation of the system needs a separate evaluation of the heat transfer coefficients (HTC) for the pedestal surface and for the free endwall surface. In the present work an innovative technique made of different methods for these two different regions is presented, so cooling performances of each are evaluated. Thermochromic liquid crystal (TLC) transient technique is used to measure detailed heat transfer coefficients only on the endwall surface: a typical transparent (Plexiglas®) test article is used. Aluminum pedestals, employing a procedure based on a finite elements code, allow to measure the average heat transfer coefficient on the insert surface. To investigate the effects on transient technique of high conductive components, a comparison with Plexiglas inserts has been performed. Results show an underestimation of HTC with TLC transient technique only for aluminum long ribs at low Reynolds number values. Nevertheless the results seem to agree with other authors both for the endwall surface and for the entire cooling system.
The performance increase for Heavy Duty (HD) gas turbines requires very high R&D efforts to improve life and reliability of the hot components. In fact the maximum temperature increase is the most important factor for the HD gas turbines development. Blade cooling of hottest and most loaded turbine stages becomes a fundamental key for R&D process in gas turbine design. Several innovations have been realised in last years about blade cooling and many industrial manufacturers have recently proposed new coolants, like steam or water. In this paper a detailed comparison among different cooling systems and coolants is proposed. Therefore performances of the most important HD gas turbine models are predicted and applications of the various cooling solutions are evaluated. An advanced cooling scheme based on air cooling closed loop is also studied. The simulations have been made from the point of view of blade cooling design and of global plant performances prediction. Interesting results have been obtained to confirm the present trend of R&D of HD gas turbines. Finally a wide comparison between different cooling solutions for several gas turbine models is presented.
The design of modern gas turbines cooling systems cannot be separated from the use of computational fluid dynamics (CFD) and the accurate estimation of the effect of film cooling. Nevertheless, a complete modeling of film cooling holes within the computational domain requires an effort both from the point of view of the mesh creation and from computational time. It is here proposed a new way to model the film cooling (FCM), capable of representing the effect of the coolant at hole exit. This is possible due to the introduction of local source terms near the hole exit in a delimited portion of the domain, avoiding the meshing process of perforations. The goal is to provide a reliable and accurate tool to simulate film-cooled turbine blades and nozzles without having to explicitly mesh the holes. The model was subjected to an intensive validation campaign, composed of two phases. During the first one, FCM results are compared to experimental data and numerical results (obtained with complete cooling holes meshing) on a series of test cases reproducing flat plate cooling configurations for different coolant conditions (in terms of blowing and density ratio). In the second phase, a film-cooled vane test case has been studied, in order to consider a real injection system and flow conditions: FCM predictions are compared to an in-house developed correlative approach and full conjugate heat transfer (CHT) results. Finally, a comparison between FCM predictions and experimental data was performed on an actual nozzle of a GE Oil & Gas heavy-duty gas turbine, in order to prove the feasibility of the procedure. The presented film cooling model (FCM) proved to be a feasible and reliable tool, able to evaluate adiabatic effectiveness, simplifying the design phase avoiding the meshing process of perforations.
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