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Shaped film cooling holes are used extensively in gas turbines to reduce component temperatures. These holes generally consist of a metering section through the material and a diffuser to spread coolant over the surface. These two hole features are created separately using electrical discharge machining, and occasionally an offset can occur between the meter and diffuser due to misalignment. The current study examines the potential impact of this manufacturing defect to the film cooling effectiveness for a well-characterized shaped hole known as the 7-7-7 hole. Five meter-diffuser offset directions and two offset sizes were examined, both computationally and experimentally. Adiabatic effectiveness measurements were obtained at a density ratio of 1.2 and blowing ratios ranging from 0.5 to 3. The detriment in cooling relative to the baseline 7-7-7 hole was worst when the diffuser was shifted upstream (aft meter-diffuser offset), and least when the diffuser was shifted downstream (fore meter-diffuser offset). At some blowing ratios and offset sizes, the fore meter-diffuser offset resulted in slightly higher adiabatic effectiveness than the baseline hole, due to a reduction in the high-momentum region of the coolant jet caused by a separation region created inside the hole by the fore meter-diffuser offset. Steady RANS predictions did not accurately capture the levels of adiabatic effectiveness or the trend in the offsets, but it did predict the fore offset’s improved performance.Copyright © 2016 by ASME

Turbine vanes and blades are generally manufactured as single or double airfoil sections that must each be installed onto a turbine disk. Between each section, a gap at the endwalls through the blade passage is present, through which high pressure coolant is leaked. Furthermore, sections can become misaligned due to thermal expansion or centrifugal forces. Flow and heat transfer around the gap is complicated due to the interaction of the mainstream and the leakage flow. An experimental and computational study was undertaken to determine the physics of the leakage flow interaction for a realistic turbine blade endwall, and assess whether steady RANS CFD, commonly used for non-axisymmetric endwall design, can be used to accurately model this interaction. Computational models were compared against experimental observations of endwall heat transfer on a contoured endwall with a midpassage gap. Endwall heat transfer coefficients were determined experimentally by using infrared thermography to capture spatially-resolved surface temperatures on a uniform heat flux surface (heater) attached to the endwall. Predictions and measurements both indicated an increase in endwall heat transfer with increasing gap leakage flow, although the distribution of heat transfer coefficients along the gap was not well captured by CFD. A misalignment of the blade endwall causing a forward-facing step for the near-endwall flow resulted in a large highly turbulent recirculation region downstream of the step and high local heat transfer that was overpredicted by CFD. Conversely, a backward-facing step reduced turbulence and local heat transfer. The misprediction of local heat transfer around the gap is thought to be caused by unsteady interaction of the passage secondary flow and gap leakage flow, which cannot be well-captured by a steady RANS approach.

Film cooling and sprayed thermal barrier coatings (TBCs) protect gas turbine components from the hot combustion gas temperatures. As gas turbine designers pursue higher turbine inlet temperatures, film cooling and thermal barrier coatings are critical in protecting the durability of turbomachinery hardware. One obstacle to the synergy of these technologies is that TBC coatings can block cooling holes when applied to the components, causing a decrease in the film cooling flow area thereby reducing coolant flow for a given pressure ratio. In this study the effect of TBC blockages was simulated on film cooling holes for widely spaced cylindrical and shaped holes. At low blowing ratios for shaped holes the blockages were found to have very little effect on adiabatic effectiveness. At high blowing ratios, the area-averaged effectiveness of shaped and cylindrical holes decreased as much as 75% from blockage. The decrease in area-averaged effectiveness was found to scale best with the effective momentum flux ratio of the jet exiting the film cooling hole for the shaped holes.Copyright © 2014 by ASME

Deposition on film-cooled turbine components was studied in an accelerated test facility. The accelerated deposition facility seeds a natural-gas burning combustor with finely-ground coal ash particulate at 1180°C and 180 m/s (M = 0.25). Both cylindrical and shaped holes, with and without TBC coating, were studied over a range of blowing ratios from 0.5 to 4.0. Coolant density ratios were maintained at values from 2.1 to 2.4. Deposition patterns generated with the cylindrical film cooling holes indicated regions of low deposition in the path of the coolant, with heightened deposition between film holes. This distinctive pattern was more accentuated at higher blowing ratios. Optical temperature measurements of the turbine component surface during deposition showed elevated temperatures between coolant paths. This temperature non-uniformity became more accentuated as deposition increased, highlighting a mechanism for deposition growth that has been documented on in-service turbines as well. The shaped-hole components exhibited little or no deposition in the region just downstream of the holes, due to the distributed coolant film. Close cylindrical hole spacing of 2.25d displayed similar behavior to the shaped hole configuration.Copyright © 2008 by ASME

Experiments were conducted to determine the impact of synfuel deposits on film cooling effectiveness and heat transfer. Scaled up models were made of synfuel deposits formed on film-cooled turbine blade coupons exposed to accelerated deposition. Three distinct deposition patterns were modeled: a large deposition pattern (maximum deposit peak≅2 hole diameters) located exclusively upstream of the holes, a large deposition pattern (maximum deposit peak≅1.25 hole diameters) extending downstream between the cooling holes, and a small deposition pattern (maximum deposit peak≅0.75 hole diameter) also extending downstream between the cooling holes. The models featured cylindrical holes inclined at 30 deg to the surface and aligned with the primary flow direction. The spacing of the holes were 3, 3.35, and 4.5 hole diameters, respectively. Flat models with the same film cooling hole geometry and spacing were used for comparison. The models were tested using blowing ratios of 0.5–2 with a turbulent approach boundary layer and 0.5% freestream turbulence. The density ratio was approximately 1.1 and the primary flow Reynolds number at the film cooling row location was 300,000. An infrared camera was used to obtain the film cooling effectiveness from steady state tests and surface convective heat transfer coefficients using transient tests. The model with upstream deposition caused the primary flow to lift off the surface over the roughness peaks and allowed the coolant to stay attached to the model. Increasing the blowing ratio from 0.5 to 2 only expanded the region that the coolant could reach and improved the cooling effectiveness. Though the heat transfer coefficient also increased at high blowing ratios, the net heat flux ratio was still less than unity, indicating film cooling benefit. For the two models with deposition between the cooling holes, the freestream air was forced into the valleys in line with the coolant holes and degraded area-averaged coolant performance at lower blowing ratios. It is concluded that the film cooling effectiveness is highest when deposition is limited to upstream of the cooling holes. When accounting for the insulating effect of the deposits between the film holes, even the panels with deposits downstream of the film holes can yield a net decrease in heat flux for some cases.

Four series of tests were performed in an accelerated deposition test facility to study the independent effects of particle size, gas temperature, and metal temperature on ash deposits from two candidate power turbine synfuels (coal and petcoke). The facility matches the gas temperature and velocity of modern first stage high pressure turbine vanes while accelerating the deposition process. Particle size was found to have a significant effect on capture efficiency with larger particles causing significant thermal barrier coating (TBC) spallation durin ga4h accelerated test. In the second series of tests, particle deposition rate was found to decrease with decreasing gas temperature. The threshold gas temperature for deposition was approximately 960°C. In the third and fourth test series, impingement cooling was applied to the back side of the target coupon to simulate internal vane cooling. Capture efficiency was reduced with increasing mass flow of coolant air; however, at low levels of cooling, the deposits attached more tenaciously to the TBC layer. Postexposure analyses of the third test series (scanning electron microscopy and X-ray spectroscopy) show decreasing TBC damage with increased cooling levels. DOI: 10.1115/1.2903901

OPO performance in 2-micron-pumped OPGaAs samples grown on different templates was found to be comparable, but exhibited significant variation within and between samples. Measured signal and idler wavelength values agreed closely with theory.