Research Papers

J. Turbomach. 2017;140(2):021001-021001-9. doi:10.1115/1.4038177.

Experimentally evaluating gas turbine cooling schemes is generally prohibitive at engine conditions. Thus, researchers conduct film cooling experiments near room temperature and attempt to scale the results to engine conditions. An increasingly popular method of evaluating adiabatic effectiveness employs pressure sensitive paint (PSP) and the heat–mass transfer analogy. The suitability of mass transfer methods as a substitute for thermal methods is of interest in the present work. Much scaling work has been dedicated to the influence of the coolant-to-freestream density ratio (DR), but other fluid properties also differ between experimental and engine conditions. Most notably in the context of an examination of the ability of PSP to serve as a proxy for thermal methods are the properties that directly influence thermal transport. That is, even with an adiabatic wall, there is still heat transfer between the freestream flow and the coolant plume, and the mass transfer analogy would not be expected to account for the specific heat or thermal conductivity distributions within the flow. Using various coolant gases (air, carbon dioxide, nitrogen, and argon) and comparing with thermal experiments, the efficacy of the PSP method as a direct substitute for thermal measurements was evaluated on a cylindrical leading edge model with compound coolant injection. The results thus allow examination of how the two methods respond to different property variations. Overall, the PSP technique was found to overpredict the adiabatic effectiveness when compared to the results obtained from infrared (IR) thermography, but still reveals valuable information regarding the coolant flow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021002-021002-11. doi:10.1115/1.4038180.

The increased design space offered by additive manufacturing (AM) can inspire unique ideas and different modeling approaches. One tool for generating complex yet effective designs is found in numerical optimization schemes, but until relatively recently, the capability to physically produce such a design had been limited by manufacturing constraints. In this study, a commercial adjoint optimization solver was used in conjunction with a conventional flow solver to optimize the design of wavy microchannels, the end use of which can be found in gas turbine airfoil skin cooling schemes. Three objective functions were chosen for two baseline wavy channel designs: minimize the pressure drop between channel inlet and outlet, maximize the heat transfer on the channel walls, and maximize the ratio between heat transfer and pressure drop. The optimizer was successful in achieving each objective and generated significant geometric variations from the baseline study. The optimized channels were additively manufactured using direct metal laser sintering (DMLS) and printed reasonably true to the design intent. Experimental results showed that the high surface roughness in the channels prevented the objective to minimize pressure loss from being fulfilled. However, where heat transfer was to be maximized, the optimized channels showed a corresponding increase in Nusselt number.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021003-021003-11. doi:10.1115/1.4038178.

Blade-to-blade interactions in a low-pressure turbine (LPT) were investigated using highly resolved compressible large eddy simulations (LESs). For a realistic setup, a stator and rotor configuration with profiles typical of LPTs was used. Simulations were conducted with an in-house solver varying the gap size between stator and rotor from 21.5% to 43% rotor chord. To investigate the effect of the gap size on the prevailing loss mechanisms, a loss breakdown was conducted. It was found that in the large gap (LG) size case, the turbulence kinetic energy (TKE) levels of the stator wake close to the rotor leading edge were only one third of those in the small gap (SG) case, due to the longer distance of constant area mixing. The small time-averaged suction side separation on the blade, found in the LG case, disappeared in the SG calculations, confirming how stronger wakes can keep the boundary layer attached. The higher intensity wake impinging on the blade, however, did not affect the time-averaged losses calculated using the control volume approach of Denton. On the other hand, losses computed by taking cross sections upstream and downstream of the blade revealed a greater distortion loss generated by the stator wakes in the SG case. Despite the suction side separation suppression, the SG case gave higher losses overall due to the incoming wake turbulent kinetic energy amplification along the blade passage.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021004-021004-10. doi:10.1115/1.4038279.

In order to predict oscillating loads on a structure, time-linearized methods are fast enough to be routinely used in design and optimization steps of a turbomachine stage. In this work, frequency-domain time-linearized Navier–Stokes computations are proposed to predict the unsteady separated flow generated by an oscillating bump in a transonic nozzle. The influence of regressive pressure waves on the aeroelastic stability is investigated. This case is representative of flutter of a compressor blade submitted to downstream stator potential effects. The influence of frequency is first investigated on a generic oscillating bump to identify the most unstable configuration. Introducing backward traveling pressure waves, it is then showed that aeroelastic stability of the system depends on the phase shift between the wave's source and the bump motion. Finally, feasibility of active control through backward traveling pressure waves is evaluated. The results show a high stabilizing effect even for low amplitude, opening new perspectives for the active control of choke flutter.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021005-021005-13. doi:10.1115/1.4038280.

Thermal closure of the engine casing is widely used to minimize undesirable blade tip leakage flows thus improving jet engine performance. This may be achieved using an impingement cooling scheme on the external casing wall, provided by manifolds attached to the outside of the engine. The assembly tolerance of these components leads to variation in the standoff distance between the manifold and the casing, and its effects on casing contraction must be understood to allow build tolerance to be specified. For cooling arrangements with promising performance, the variation in closure with standoff distance of z/d = 1–6 were investigated through a mixture of extensive numerical modeling and experimental validation. A cooling manifold, typical of that adopted by several engine companies, incorporating three different arrays of short cooling holes (chosen from previous study by Choi et al. (2016, “The Relative Performance of External Casing Impingement Cooling Arrangements for Thermal Control of Blade Tip Clearance,” ASME J. Turbomach., 138(3), p. 031005.)) and thermal control dummy flanges were considered. Typical contractions of 0.5–2.2 mm are achieved from the 0.02–0.35 kg/s of the current casing cooling flows. The variation in heat transfer coefficient observed with standoff distance is much lower for the sparse array investigated compared to previous designs employing arrays typical of blade cooling configurations. The reason for this is explained through interrogation of the local flow field and resultant heat transfer coefficient. This implies that acceptable control of the circumferential uniformity of case cooling can be achieved with relatively large assembly tolerance of the manifold relative to the casing.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021006-021006-8. doi:10.1115/1.4038275.

In film cooling flows, it is important to know the temperature distribution resulting from the interaction between a hot main flow and a cooler jet. However, current Reynolds-averaged Navier–Stokes (RANS) models yield poor temperature predictions. A novel approach for RANS modeling of the turbulent heat flux is proposed, in which the simple gradient diffusion hypothesis (GDH) is assumed and a machine learning (ML) algorithm is used to infer an improved turbulent diffusivity field. This approach is implemented using three distinct data sets: two are used to train the model and the third is used for validation. The results show that the proposed method produces significant improvement compared to the common RANS closure, especially in the prediction of film cooling effectiveness.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021007-021007-12. doi:10.1115/1.4038316.

Rotor blade vibrations observed in the Darmstadt transonic compressor rig are investigated in this paper. The vibrations are nonsynchronous and occur in the near stall (NS) operating region. Rotor tip flow fluctuations traveling near the leading edge (LE) against the direction of rotation (in the rotor relative frame of reference) with about 50% blade tip speed are found to be the reason for the occurrence of the vibrations. The investigations show that the blockage at the rotor tip is an important factor for the aeroelastic stability of the compressor in the NS region. It is found that by application of a recirculating tip injection (TI) casing treatment, the aeroelastic stability increases as a result of reduced blockage in the rotor tip region.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021008-021008-9. doi:10.1115/1.4038277.

Effusion cooling has been a popular technology integrated into the design of gas turbine combustor liners. A staggering amount of research was completed that quantified performance with respect to operating conditions and cooling hole geometric properties; however, most of these investigations did not address the influence of the manufacturing process on the hole shape. This study completed an adiabatic wall numerical analysis using the realizable k–ϵ turbulence model of a laser-drilled hole that had a nozzled profile with an area ratio of 0.24 and five additional cylindrical, nozzled, diffusing, and fileted holes that yielded the same hole mass flow rate at representative engine conditions. The traditional methods for quantifying blowing ratio yielded the same value for all holes that was not useful considering the substantial differences in film cooling performance. It was proposed to define hole mass flux based on the outlet y-cross-sectional area projected onto the inclination angle plane. This gave blowing ratios that correlated to better and worse cooling performance for the diffusing and nozzled holes, respectively. The diffusing hole delivered the best film cooling due to having the lowest effluent velocity and greatest amount of in-hole turbulent production, which coincided with the worst discharge coefficient.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021009-021009-12. doi:10.1115/1.4038278.

Turbine vane endwalls are highly susceptible to intensive heat load due to their large exposed area and complex flow field especially for the first stage of the vane. Therefore, a suitable film cooling design that properly distributes the given amount of coolant is critical to keep the vane endwall from failure at the same time to maintain a good balance between manufacturing cost, performance, and durability. This work is focused on film cooling effectiveness evaluation on full-scale heavy-duty turbine vane endwall and the performance comparison with different film cooling pattern designs in the literature. The area of interest (AOI) of this study is on the inner endwall (hub) of turbine vane. Tests were performed in a three-vane annular sector cascade under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09 and the freestream turbulence intensity is 12%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2–4%) and density ratios (DR = 1.0, 1.5), are investigated. The conduction-error free pressure-sensitive paint (PSP) technique is utilized to evaluate the local flow behavior as well as the film cooling performance. The presented results are expected to provide the gas turbine engine designer a direct comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021010-021010-9. doi:10.1115/1.4038281.

Secondary flows in vane passages sweep off the endwall and onto the suction surface at a location typically close to the throat. These endwall/vane junction flows often have an immediate impact on heat transfer in this region and also move any film cooling off the affected region of the vane. The present paper documents the impact of secondary flows on suction surface heat transfer acquired over a range of turbulence levels (0.7–17.4%) and a range of exit chord Reynolds numbers (500,000–2,000,000). Heat transfer data are acquired with both an unheated endwall boundary condition and a heated endwall boundary condition. The vane design includes an aft loaded suction surface and a large leading edge diameter. The unheated endwall boundary condition produces initially very high heat transfer levels due to the thin thermal boundary layer starting at the edge of heating. This unheated starting length effect quickly falls off with the thermal boundary layer growth as the secondary flow sweeps up onto the vane suction surface. The heat transfer visualization for the heated endwall condition shows no initial high heat transfer level near the edge of heating on the vane. The heat transfer level in the region affected by the secondary flows is largely uniform, except for a notable depression in the affected region. This heat transfer depression is believed due to an upwash region generated above the separation line of the passage vortex, likely in conjunction with the counter rotating suction leg of the horseshoe vortex. The extent and definition of the secondary flow-affected region on the suction surface are clearly evident at lower Reynolds numbers and lower turbulence levels when the suction surface flow is largely laminar. The heat transfer in the plateau region has a magnitude similar to a turbulent boundary layer. However, the location and extent of this secondary flow-affected region are less perceptible at higher turbulence levels where transitional or turbulent flow is present. Also, aggressive mixing at higher turbulence levels serves to smooth out discernable differences in the heat transfer due to the secondary flows.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(2):021011-021011-10. doi:10.1115/1.4038317.

Computational fluid dynamics (CFD) has been widely used for compressor design, yet the prediction of performance and stage matching for multistage, high-speed machines remains challenging. This paper presents the authors' effort to improve the reliability of CFD in multistage compressor simulations. The endwall features (e.g., blade filet and shape of the platform edge) are meshed with minimal approximations. Turbulence models with linear and nonlinear eddy viscosity models are assessed. The nonlinear eddy viscosity model predicts a higher production of turbulent kinetic energy in the passages, especially close to the endwall region. This results in a more accurate prediction of the choked mass flow and the shape of total pressure profiles close to the hub. The nonlinear viscosity model generally shows an improvement on its linear counterparts based on the comparisons with the rig data. For geometrical details, truncated filet leads to thicker boundary layer on the filet and reduced mass flow and efficiency. Shroud cavities are found to be essential to predict the right blockage and the flow details close to the hub. At the part speed, the computations without the shroud cavities fail to predict the major flow features in the passage, and this leads to inaccurate predictions of mass flow and shapes of the compressor characteristic. The paper demonstrates that an accurate representation of the endwall geometry and an effective turbulence model, together with a good quality and sufficiently refined grid, result in a credible prediction of compressor matching and performance with steady-state mixing planes.

Commentary by Dr. Valentin Fuster

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