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Research Papers

J. Turbomach. 2016;139(4):041001-041001-10. doi:10.1115/1.4034983.

As air traffic continues to increase in the subtropical areas where high moisture laden air is present at subfreezing conditions, engine icing probability increases. It has been shown that compressor stages rematch under icing conditions—front stages are choked, while rear stages throttle due to ice melting and evaporation. Such an analysis uses various empirical models to represent ice-breakup and water-splash processes as ice/water particles interact with rotors/stators. This paper presents a compressor stall sensitivity analysis around different splash models. The effect of droplet splash at both rotor and stator blades, blade solidity effect, and trailing edge shed effect is modeled. A representative ten-stage high-speed compressor section operating near design point (100% Nc) is used for the study. Results show that the temperature drop at high-pressure compressor (HPC) exit and the overall compressor operability are functions of evaporating stages, and droplet–blade interaction models influence them. A comprehensive compressor stability envelope has been evaluated for different models. It is observed that the droplet–blade interaction behavior influences overall compressor stability and the stall-margin predictions can vary by as much as 25% with different models. Therefore, there is a need for better calibration and continual improvement of empirical models to capture compressor interstage dynamics and stage rematching accurately under ice/water ingestion.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(4):041005-041005-23. doi:10.1115/1.4035070.

Fouling afflicts gas turbine operation from first time application. Filtration systems and washing operations work against air contaminants in order to limit the particles entering the compressor inlet and remove the existing deposits. In this work, a global overview of the operational experience of the manufacturer, the filtration systems, and the particle deposition of the compressor are reported. The data reported in this review have been collected from 60 years (1956–2015) of ASME Turbo Expo proceedings. This conference is recognized as the must-attend event for turbomachinery professionals. Through the years, many issues have been resolved by the contributions of this conference. Regarding the compressor fouling phenomenon, the contributions presented at the ASME Turbo Expo mark the high level of development in this field of research, thanks to the simultaneous presence of manufacturers, government, and academia attendees. The goal of the authors is to describe the technological evolution and challenges faced by manufacturers and researchers through the years, highlighting the state of the art in the knowledge of fouling, and defining the background on which further studies will be based.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(4):041006-041006-12. doi:10.1115/1.4035073.

The flow at the combustor turbine interface of power generation gas turbines with can combustors is characterized by high and nonuniform turbulence levels, lengthscales, and residual swirl. These complexities have a significant impact on the first vanes aerothermal performance and lead to challenges for an effective turbine design. To date, this design philosophy mostly assumed steady flow and thus largely disregards the intrinsic unsteadiness. This paper investigates the steady and unsteady effects of the combustor flow with swirl on the turbines first vanes. Experimental measurements are conducted on a high-speed linear cascade that comprises two can combustors and four nozzle guide vanes (NGVs). The experimental results are supported by a large eddy simulation (LES) performed with the inhouse computational fluid dynamics (CFD) flow solver TBLOCK. The study reveals the highly unsteady nature of the flow in the first vane and its effect on the heat transfer. A persistent flow structure of concentrated vorticity is observed. It wraps around the unshielded vane's leading edge (LE) at midspan and periodically oscillates in spanwise direction due to the interaction of the residual low-pressure swirl core and the vane's potential field. Moreover, the transient behavior of the horseshoe-vortex system due to large fluctuations in incidence is demonstrated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(4):041007-041007-10. doi:10.1115/1.4035043.

During the testing of development engines and components, intrusive instrumentation such as Kiel-head pitot probes and shrouded thermocouples are used to evaluate gas properties and performance. The size of these instruments can be significant relative to the blades, and their impact on aerodynamic efficiency must be considered when analyzing the test data. This paper reports on such parasitic losses for instruments mounted on the leading edge of a stator in a low-pressure turbine, with particular emphasis on understanding the impact of probe geometry on the induced loss. The instrumentation and turbine blades have been modeled in a low Mach number cascade facility with an upstream turbulence grid. The cascade was designed so that the leading edge probes were interchangeable in situ, allowing for rapid testing of differing probe geometries. Reynolds-averaged Navier–Stokes (RANS) calculations were performed to complement the experiments and improve understanding of the flow behavior. A horseshoe vortex-like system forms at the join of the probe body and blade leading edge, generating pairs of streamwise vortices which convect over the blade pressure and suction surfaces. These vortices promote mixing between the freestream and boundary layer fluid and promote the transition of the boundary layer from laminar to turbulent flow. The size and shape of the leading edge probes relative to the blade vary significantly between applications. Tests with realistic probe geometries demonstrate that the detailed design of the shroud bleed system can impact the loss. A study of idealized cylinders is performed to isolate the impact of probe diameter, aspect ratio, and incidence. Beyond a probe aspect ratio of two, parasitic loss was found to scale approximately with probe frontal area.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(4):041008-041008-14. doi:10.1115/1.4035076.

Experiments preformed in the JHU refractive index matched facility examine flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. High-speed imaging of cavitation performed at low pressures qualitatively visualizes vortical structures. Stereoscopic particle image velocimetry (SPIV) measurements provide detailed snapshots and ensemble statistics of the flow in a series of meridional planes. At prestall condition, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures shortly after rollup. The most prominent instability involves periodic formation of large-scale backflow vortices (BFVs) that extend diagonally upstream, from the suction side (SS) of one blade at midchord to the pressure side (PS) near the leading edge of the next blade. The 3D vorticity distributions obtained from data recorded in closely spaced planes show that the BFVs originate form at the transition between the high circumferential velocity region below the TLV center and the main passage flow radially inward from it. When the BFVs penetrate to the next passage across the tip gap or by circumventing the leading edge, they trigger a similar phenomenon there, sustaining the process. Further reduction in flow rate into the stall range increases the number and size of the backflow vortices, and they regularly propagate upstream of the leading edge of the next blade, where they increase the incidence angle in the tip corner. As this process proliferates circumferentially, the BFVs rotate with the blades, indicating that there is very little through flow across the tip region.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(4):041009-041009-12. doi:10.1115/1.4035074.

The continuing maturation of metal laser-sintering technology (direct metal laser sintering (DMLS)) presents the opportunity to derisk the engine design process by experimentally down-selecting high-pressure nozzle guide vane (HPNGV) cooling designs using laboratory tests of laser-sintered—instead of cast—parts to assess thermal performance. Such tests could be seen as supplementary to thermal-paint test engines, which are used during certification to validate cooling system designs. In this paper, we compare conventionally cast and laser-sintered titanium alloy parts in back-to-back experimental tests at engine-representative conditions over a range of coolant mass flow rates. Tests were performed in the University of Oxford Annular Sector Heat Transfer Facility. The thermal performance of the cast and laser-sintered parts—measured using new infrared processing techniques—is shown to be very similar, demonstrating the utility of laser-sintered parts for preliminary engine thermal assessments. We conclude that the methods reported in this paper are sufficiently mature to make assessments which could influence engine development programs.

Commentary by Dr. Valentin Fuster

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