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J. Turbomach. 2018;140(7):071001-071001-13. doi:10.1115/1.4039806.

Self-recirculating injection, which bleeds air from the downstream duct of the last blade row and injects air as a wall jet upstream of the first rotor blade row, is experimentally investigated after the design of its structure in single- and three-stage axial flow compressors. External injection and outlet bleed air are selected for comparison. Results show that self-recirculating injection can improve the stall margin by 13.67% and 13% on the premise of no efficiency penalty in single- and three-stage axial flow compressors with only 0.7% and 4.2% of the total injected momentum ratio recirculated near stall, respectively. The self-recirculating injection is the best among all the three cases if the influence on pressure rise coefficient and efficiency is comprehensively considered. Moreover, findings indicate that the upstream injection plays an important role in terms of stability-enhancement. The details of the flow field are captured using a collection of pressure transducers on the casing with circumferential and chordwise spatial resolution. A detailed comparative analysis of the endwall flow indicates that the self-recirculating injection can postpone the occurrence of stalling in the proposed compressor by delaying the forward movement of the interface between the tip leakage flow (TLF) and main stream flow (MF), weakening the unsteadiness of TLF (UTLF), and sharply decreasing the circumferentially propagating speed dominated by the UTLF that triggers the spike-type stall inception. Finally, the stall control concept on the stage that first generates stall inception using self-recirculating injection is proposed. This study helps to guide the design of self-recirculating injection in actual application.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071002-071002-11. doi:10.1115/1.4039821.
OPEN ACCESS

Exhaust diffusers significantly enhance the available power output and efficiency of gas and steam turbines by allowing lower turbine exit pressures. The residual dynamic pressure of the turbine outflow is converted into static pressure, which is referred to as pressure recovery. Since total pressure losses and construction costs increase drastically with diffuser length, it is strongly preferred to design shorter diffusers with steeper opening angles. However, these designs are more susceptible to boundary layer separation. In this paper, the stabilizing properties of tip leakage vortices generated in the last rotor row and their effect on the boundary layer characteristics are examined. Based on analytical considerations, for the first time, a correlation between the pressure recovery of the diffuser and the integral rotor parameters of the last stage, namely, the loading coefficient, flow coefficient, and reduced frequency, is established. Experimental data and scale-resolving simulations, carried out with the shear stress transport scale-adaptive simulation (SST-SAS) method, both show excellent agreement with the correlation. Blade tip vortex strength predominantly depends on the amount of work exchanged between fluid and rotor, which in turn is described by the nondimensional loading coefficient. The flow coefficient influences mainly the orientation of the vortex, which affects the interaction between vortex and boundary layer. The induced velocity field accelerates the boundary layer, essentially reducing the thickness of the separated layer or even preventing separation locally.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071003-071003-13. doi:10.1115/1.4040101.

Film-cooling effectiveness of rectangular diffusion holes under an inclination angle α = 45 deg, an orientation angle β = 45 deg, and a length-to-diameter ratio of L/D = 8.5 were, respectively, examined in a flat-plate experimental facility using the pressure sensitive paint (PSP) technique. Experiments were performed at a density ratio of DR = 1.38 and a mainstream turbulence intensity of Tu = 3.5%. The semicircle sidewall rectangular diffusion hole varied at three cross-sectional aspect ratios, i.e., AS = 3.4, 4.9, and 6.6. The tested results were compared with the baseline design with an inclination angle α = 30 deg, an orientation angle β = 0 deg, and a length-to-diameter ratio L/D = 6. A three-dimensional (3D) numerical simulation method was employed to analyze the flow field. The experimental results showed that the increased inclination angle converted the bi- or tri-peak effectiveness pattern of the baseline design to a single-peak pattern, weakened the lateral diffusion of coolant, and consequently decreased cooling effectiveness obviously. The decreased magnitude amplified with the increase of cross-sectional aspect ratio and blowing ratio. The adding of orientation angle seriously weakened the cooling effectiveness of the baseline design, and the blowing ratio and cross-sectional aspect ratio had almost no effect on overall cooling effectiveness. The elongated hole length provided a uniform distribution of lateral cooling effectiveness, which produced differential effects on the bi- or tri-peak pattern. The elongated hole length decreased the cooling effectiveness on the near hole region, but had less effects on overall cooling effectiveness, except the high blowing ratio.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071004-071004-12. doi:10.1115/1.4040111.

Blade row interactions drive the unsteady performance of high-pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake's path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue three-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady stator 2 performance, time-resolved interrogations of the stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between rotor 1 and rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the stator 1 wake and rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, stator 2 experiences significantly different inlet blockage and unsteadiness from the rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of stator 2.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071005-071005-11. doi:10.1115/1.4039807.

Axial compressors may be operated under off-design incidences due to variable operating conditions. Therefore, a successful design requires accurate performance and stability limits predictions under a wide operating range. Designers generally rely both on correlations and on Reynolds-averaged Navier–Stokes (RANS), the accuracy of the latter often being questioned. The present study investigates profile losses in an axial compressor linear cascade using both RANS and wall-resolved large eddy simulation (LES), and compares with measurements. The analysis concentrates on “loss buckets,” local separation bubbles and boundary layer transition with high levels of free stream turbulence, as encountered in real compressor environment without and with periodic incoming wakes. The work extends the previous research with the intention of furthering our understanding of prediction tools and improving our quantification of the physical processes involved in loss generation. The results show that while RANS predicts overall profile losses with good accuracy, the relative importance of the different loss mechanisms does not match with LES, especially at off-design conditions. This implies that a RANS-based optimization of a compressor profile under a wide incidence range may require a thorough LES verification at off-design incidence.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071006-071006-9. doi:10.1115/1.4039842.

The ability to understand and predict the pressure losses of orifices is important in order to improve the air flow within the secondary air system. This experimental study investigates the behavior of the discharge coefficient for circular orifices with inlet cross flow which is a common flow case in gas turbines. Examples of this are at the inlet of a film cooling hole or the feeding of air to a blade through an orifice in a rotor disk. Measurements were conducted for a total number of 38 orifices, covering a wide range of length-to-diameter ratios, including short and long orifices with varying inlet geometries. Up to five different chamfer-to-diameter and radius-to-diameter ratios were tested per orifice length. Furthermore, the static pressure ratio across the orifice was varied between 1.05 and 1.6 for all examined orifices. The results of this comprehensive investigation demonstrate the beneficial influence of rounded inlet geometries and the ability to decrease pressure losses, which is especially true for higher cross flow ratios where the reduction of the pressure loss in comparison to sharp-edged holes can be as high as 54%. With some exceptions, the chamfered orifices show a similar behavior as the rounded ones but with generally lower discharge coefficients. Nevertheless, a chamfered inlet yields lower pressure losses than a sharp-edged inlet. The obtained experimental data were used to develop two correlations for the discharge coefficient as a function of geometrical as well as flow properties.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071007-071007-12. doi:10.1115/1.4039942.

Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The midchord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient (HTC). The film cooling along the midchord of the blade uses multiple rows of shaped diffusion holes. The trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071008-071008-11. doi:10.1115/1.4040030.

The effect of inlet distortion from curved intake ducts on jet engine fan stability is an important consideration for next-generation passenger aircraft such as the boundary layer ingestion (BLI) “silent aircraft.” Highly complex inlet flows which occur can significantly affect fan stability. Future aircraft designs are likely to feature more severe inlet distortion, pressing the need to understand the important factors influencing design. This paper presents the findings from a large computational fluid dynamics (CFD) investigation into which aspects of inlet distortion cause the most significant reductions in stall margin and, therefore, which flow patterns should be targeted by mitigating technology. The study considers the following aspects of distortion commonly observed in intakes: steady vortical distortion due to secondary flow, unsteady vortical distortion due to vortex shedding and mixing, static pressure distortion due to curved streamlines, and low momentum endwall flow due to thickened boundary layers or separation. Unsteady CFD was used to determine the stall points of a multipassage transonic rotor geometry with each of the inlet distortion patterns applied. Interesting new evidence is provided, which suggests that low momentum flow in the tip region, rather than distortion in the main body of the flow, leads to damaging instability.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(7):071009-071009-11. doi:10.1115/1.4039433.

Applications such as boundary-layer-ingesting (BLI) fans and compressors in turboprop engines require continuous operation with distorted inflow. A low-speed axial fan with incompressible flow is studied in this paper. The objectives are to (1) identify the physical mechanisms which govern the fan response to inflow distortions and (2) determine how fan performance scales as the type and severity of inlet distortion varies at the design flow coefficient. A distributed source term approach to modeling the rotor and stator blade rows is used in numerical simulations in this paper. The model does not include viscous losses so that changes in diffusion factor are the primary focus. Distortions in stagnation pressure and temperature as well as swirl are considered. The key findings are that unless sharp pitchwise gradients in the diffusion response, strong radial flows, or very large distortion magnitudes are present, the response of the blade rows for strong distortions can be predicted by scaling up the response to a weaker distortion. In addition, the response to distortions which are composed of nonuniformities in several inlet quantities can be predicted by summing up the responses to the constituent distortions.

Commentary by Dr. Valentin Fuster

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