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Research Papers

J. Turbomach. 2017;139(6):061001-061001-13. doi:10.1115/1.4035212.

The effect of the diffuser vane setting angle on the performance of a centrifugal compressor was experimentally investigated. Seven different vaned diffusers were tested with the same impeller. In order to evaluate the vaned diffusers under the same flow range and diffuser inlet conditions, all the diffusers were designed to have the same throat area, the same diffuser height. and the same vane inlet position. The first tests were performed by a compressor with a volute. In this case, the diffuser outlet conditions were varied along the circumferential directions due to the nonaxisymmetric geometry of the volute. In the second tests, four of the seven vaned diffusers were tested using a compressor with a collector. In this case, the diffuser outlet conditions were more uniform along the circumferential directions compared to the tests using the compressor with the volute. The effect of the vane setting angle on the stage characteristics and diffuser performance as well as the effect of the circumferential distortion caused by the volute on the stage characteristics are presented. The impact of the incidence loss on the vaned diffuser performance is discussed using 1D vaned diffuser performance modeling. Comparisons of the 1D predictions and the tests results show the incidence loss has a strong influence on vaned diffuser performance.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061002-061002-12. doi:10.1115/1.4035232.

This paper presents a new unconventional philosophy for high-pressure (HP) vane design. It is proposed that the most natural design starting point for admitting and accelerating flow with minimum loss and secondary flow is a trumpet-shaped flow-path which gradually turns to the desired angle. Multiple trumpet-shaped inlets are seamlessly blended into the (annular or partitioned) combustor walls resulting in a highly lofted flow-path, rather than a traditional flow-path defined by distinct airfoil and endwall surfaces. We call this trumped-shaped inlet the fully lofted oval vane (FLOvane). The purpose of this paper is to describe the FLOvane concept and to present back-to-back CFD analyses of two current industrial gas turbines with conventional and FLOvane-modified designs. The resulting designs diverge significantly from conventional designs in terms of both process and final geometric form. Computational fluid dynamic predictions for the FLOvane-modified designs show improved aerodynamic performance characteristics, reduced heat load, improved cooling performance, improved thermal–mechanical life, and improved stage/engine efficiency. The mechanisms for improved performance include reduction of secondary flows, reduced mixing of coolant flow with the mainstream flow, reduced skin friction, and improved coolant distribution. In the two current industrial gas turbine engines, the absolute (percentage point) improvement in stage isentropic efficiency when the FLOvane design was included was estimated at 0.33% points and 0.40% points without cooling flow reduction, and 1.5% points in one case and much more is expected for the other case when cooling flow reductions were accounted for.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061003-061003-11. doi:10.1115/1.4035162.

Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces: a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using pressure sensitive paint (PSP) technique on the endwall of a scaled up, midrange industrial turbine vane geometry with the multiple rows of discrete film cooling (DFC) holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 × 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate free-stream turbulence (FST) level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR), were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface with the corresponding spanwise averaged values along the axial direction.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061004-061004-7. doi:10.1115/1.4035211.

The steady improvement of aircraft engine performance has led toward more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 × 106 and 6 × 106). The steady and unsteady heat flux was retrieved using thin-film two-layered gauges. The complexity of the tests required the development of an original heat transfer postprocessing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061005-061005-11. doi:10.1115/1.4035233.

The operating range of a compressor is usually limited by the rapid growth of three-dimensional (3D) separations in the endwall flow region. In contrast, the freestream region is not usually close to its diffusion limit and has little effect on overall range. In light of these two distinct flow regions, this paper considers how velocity triangles in the endwall region should be designed to give a more balanced spanwise failure across the span of a blade row. In the first part of this paper, the sensitivity of 3D separations in a single blade row to variations in realistic multistage inlet conditions and endwall geometry is investigated. It is shown that a blade's 3D separation size is largely controlled by the dynamic pressure within the incoming endwall “repeating stage” boundary layer and not the detailed local geometry within the blade row. In the second part of this paper, the traditional design process is “flipped.” Instead of redesigning a blade's endwall geometry to cope with a particular inlet profile into the blade row, the endwall region is redesigned in the multistage environment to “tailor” the inlet profile into downstream blade rows, giving the designer a new extra degree-of-freedom. This extra degree-of-freedom is exploited to balance freestream and endwall operating range, resulting in a compressor having an increased operating range of ∼20%. If this increased operating range is traded with reduced blade count, it is shown that a design efficiency improvement of ∼0.5% can be unlocked.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061006-061006-9. doi:10.1115/1.4035521.

Effects of a large rotor tip gap on the performance of a one and a half stage axial compressor are investigated in detail with a numerical simulation based on large eddy simulation (LES) and available particle image velocimetry (PIV) data. This paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade's tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade's tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all the operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Nonsynchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061007-061007-13. doi:10.1115/1.4035274.

Tangential endwall contouring (TEWC) is intended to improve the turbomachinery blading efficiency. This paper summarizes the experimental and numerical investigation of a test turbine with endwall contoured vanes and blades. Constant section (2D) airfoils as well as optimized compound lean (3D) high pressure steam turbine blading in baseline and endwall contoured configurations have been examined. Brush seals (BSs) are implemented within the casing sided cavities to minimize the leakage flow near the tip endwalls, where the contouring is also applied. The pressure and temperature data that are recorded in three axial measuring planes are plotted to visualize the change in flow structure. This shows that the efficiency is increased for 2D airfoils by means of endwall contouring. However, the efficiency of the first stage suffers, and the endwall contouring is still beneficial for the overall performance of the engine. Both phenomena (an efficiency loss in stage one and an improvement of the performance in stage two) have also been measured for the optimized 3D configurations; thus, it can be expected that the endwall contouring has also a beneficial impact on the performance of multirow turbines. The numerical investigations demonstrate in detail, how the secondary flow phenomena are influenced by end-wall contouring and a description of the changes in vortex formations as well as blade loading are given for the various configurations. It has been found that for steady computational fluid dynamics (CFD) simulations the use of stage interfaces suppresses the positive effects of the endwall contour onto the downstream blade row.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061008-061008-12. doi:10.1115/1.4035509.

Blowing from the tip of a turbine blade was studied experimentally to determine if total pressure loss could be reduced. Experiments were done with a linear cascade in a low-speed wind tunnel. Total pressure drop through the blade row and secondary velocity fields in the passage between two blades were measured. Cases were documented with various blowing hole configurations on flat and squealer tipped blades. Blowing normal to the tip was not helpful and in some cases increased losses. Blowing from the bottom of a squealer cavity provided little benefit. With a flat tip, blowing from holes located near and inclined toward the pressure side generally reduced total pressure drop by reducing the effect of the tip leakage vortex. Holes near the axial location of maximum loading were most helpful, while holes closer to the leading and trailing edges were not as effective. Higher jet velocity resulted in larger total pressure drop reduction. With a tip gap of 1.5% of axial chord, jets with a velocity 1.5 times the cascade inlet velocity had a significant effect. A total pressure drop reduction of the order 20% was possible using a jet mass flow of about 0.4% of the main flow. Jets were most effective with smaller tip gaps, as they were more able to counter the leakage flow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061009-061009-11. doi:10.1115/1.4035275.

Film cooling is often adopted, where coolant jets are ejected to form a protective layer on the surface against the hot combustor gases. The bending of jets in crossflow results in counter rotating vortex pair (CRVP), which is a cause for high jet lift-off and poor film cooling effectiveness in the near field. There are efforts to mitigate this detrimental effect of CRVP, and thus, to improve the film cooling performance. In the present study, the effects of both downwash and upwash type of vortex generator (VG) on film cooling are numerically analyzed. A series of discrete holes on a flat plate with 35 deg streamwise orientation and connected to a common delivery plenum is used here, where the vortex generators are placed upstream of the holes. The blowing ratio and the density ratio are considered as 0.5 and 1.2, respectively, with a Reynolds number based on freestream velocity and diameter of hole being 15,885. The computations are performed by ANSYS fluent 13.0 using k–ε realizable turbulence model. The results show that vortices generated by downwash vortex generator (DWVG) counteracts the effect of CRVP preventing the jet lift-off, which results in increased effectiveness in streamwise as well as in spanwise directions. However, upwash vortex generator (UWVG) augments the effect of CRVP, resulting in poor performance of film cooling.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(6):061010-061010-12. doi:10.1115/1.4035276.

In this paper, we describe the design, modeling, and experimental testing of a film cooling scheme employed on an unshrouded high-pressure (HP) rotor casing. The casing region has high thermal loads at both low and high frequency, with the flow being dominated by the potential field of the rotor and over-tip leakage flows. Increasingly high turbine entry temperatures necessitate internal and film cooling of the casing to ensure satisfactory service life and performance. There are, however, very few published studies presenting computational fluid dynamics (CFD) and experimental data for cooled rotor casings. Experimental testing was performed on a film-cooled rotor casing in the Oxford Turbine Research Facility (OTRF)—a rotating transonic facility of engine scale. Unsteady CFD of an HP rotor blade row with a film-cooled casing was undertaken, uniquely with a domain utilizing a sliding interface in the tip gap. A high density array of thin film heat flux gauges (TFHFGs) was used to obtain time-resolved and time-mean results of adiabatic wall temperature and film cooling effectiveness on the film-cooled rotor casing between −30% and +125% rotor tip axial chord. Results are compared to CFD predictions, and mechanisms for interaction of the coolant with the rotor tip are proposed and discussed. Acoustic effects within casing coolant holes due to the passing of the rotor are demonstrated on a 3D CFD geometry, supporting conclusions drawn in earlier work by the authors on the importance of this effect in a casing film cooling system.

Commentary by Dr. Valentin Fuster

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