Double Wall Cooling of an Effusion Plate With Simultaneous Cross Flow and Impingement Jet Array Internal Cooling

The present experimental investigation considers a full coverage film cooling arrangement with different streamwise static pressure gradients. The film cooling holes in adjacent streamwise rows are staggered with respect to each other, with sharp edges and streamwise inclination angles of 20 deg with respect to the liner surface. Data are provided for turbulent film cooling, contraction ratios of 1 and 4, blowing ratios (BRs) (at the test section entrance) of 2.0, 5.0, and 10.0, a coolant Reynolds number of 12,000, freestream temperatures from 75 °C to 115 °C, a film hole diameter of 7 mm, and density ratios from 1.15 to 1.25. Nondimensional streamwise and spanwise film cooling hole spacings, X/D and Y/D, are 18 and 5, respectively. Data illustrating the effects of contraction ratio, BR, and streamwise location on local, line-averaged, and spatially averaged adiabatic film effectiveness data; and on local, line-averaged and spatially averaged heat transfer coefficient data are presented. Varying BR values are present along the length of the contraction passage, which contains the cooling hole arrangement, when contraction ratio is 4. Dependence on BR indicates important influences of coolant concentration and distribution. For example, line-averaged and spatially averaged adiabatic effectiveness data show vastly different changes with BR for the configurations with contraction ratios of 1 and 4. In addition, much larger effectiveness alterations are present as BR changes from 2.0 to 10.0, when significant acceleration is present and Cr = 4 (in comparison with the Cr = 1 data).

Within the present investigation, a louver slot is employed upstream of an array full coverage film cooling holes. Cooling air is supplied using a combination arrangement, with cross-flow and impingement together. The louver consists of a row of film cooling holes, contained within a specially-designed device which concentrates, and directs the coolant from a slot, so that it then advects as a layer downstream along the test surface. This louver-supplied coolant is then supplemented by coolant which emerges from different rows of downstream film cooling holes. The same coolant supply passage is employed for the louver row of holes, as well as for the film cooling holes, such that different louver and film cooling mass flow rates are set by different hole diameters for the two different types of cooling holes. The results are different from data provided by past investigations, because of the use and arrangement of the louver slot, and because of the unique coolant supply configurations. The experimental results are given for mainstream Reynolds numbers from 107000 to 114000. Full-coverage blowing ratios are constant with streamwise location along the test surface, and range from 3.68 to 5.70. Corresponding louver slot blowing ratios then range from 1.72 to 2.65. Provided are heat transfer coefficient and adiabatic effectiveness distributions, which are measured along the mainstream side of the test plate. Both types of data show less variation with streamwise development location, relative to results obtained without a louver employed, when examined at the same approximate effective blowing ratio, mainstream Reynolds number, cross flow Reynolds number, and impingement jet Reynolds number. When compared at the same effective blowing ratio or the same impingement jet Reynolds number, spanwise-averaged heat transfer coefficients are consistently lower, especially for the downstream regions of the test plate, when the louver is utilized. With the same type of comparisons, the presence of the louver slot results in significantly higher values of adiabatic film cooling effectiveness (spanwise-averaged), particularly at and near the upstream portions of the test plate. With such characteristics, dramatic increases in thermal protection are provided by the presence of the louver slot, the magnitudes of which vary with experimental condition and test surface location.

Film cooling performance was evaluated numerically for three mainstream Reynolds numbers and four blowing ratios (BR). A computational model based on finite volume discretization was used to solve an incompressible and transient flow over a NACA 4412 cascade vane. Several passive scalars were included in the model to evaluate the condition of adiabatic temperature and constant temperature for the surface vane. For the adiabatic temperature condition, the film effectiveness mainly depends on the jet trajectory and recirculation zones. For the constant temperature condition, the net heat flux reduction (NHFR) varies according to the boundary layer separation and reattachment. Consequently, misleading conclusions could be drawn if only one of the two approaches is adopted. For instance, the mainstream Reynolds number Re∞ = 3615 reached a maximum average effectiveness lower than 0.3 with an average NHFR of 0.25. However, for Re∞ = 10,845 the maximum average effectiveness was close to 0.45, but with negative average NHFR values. This finding demonstrates the need to explore new indicators like jet trajectory, convective coefficient and skin friction coefficient, as presented in this paper.

Turbine design engineers have to ensure that film cooling can provide sufficient protection to turbine blades from the hot mainstream gas, while keeping the losses low. Film cooling hole design parameters include inclination angle (α), compound angle (β), hole inlet geometry, and hole exit geometry. The influence of these parameters on aerodynamic loss and net heat flux reduction is investigated, with loss being the primary focus. Low-speed flat plate experiments have been conducted at momentum flux ratios of IR = 0.16, 0.64, and 1.44. The film cooling aerodynamic mixing loss, generated by the mixing of mainstream and coolant, can be quantified using a three-dimensional analytical model that has been previously reported by the authors. The model suggests that for the same flow conditions, the aerodynamic mixing loss is the same for holes with different α and β but with the same angle between the mainstream and coolant flow directions (angle κ). This relationship is assessed through experiments by testing two sets of cylindrical holes with different α and β: one set with κ = 35 deg, and another set with κ = 60 deg. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimize κ to obtain the lowest loss, but maximize β to achieve the best heat transfer performance. A suggestion on improving the loss model is also given. Five different hole geometries (α = 35.0 deg, β = 0 deg) were also tested: cylindrical hole, trenched hole, fan-shaped hole, D-Fan, and SD-Fan. The D-Fan and the SD-Fan have similar hole exits to the fan-shaped hole but their hole inlets are laterally expanded. The external mixing loss and the loss generated inside the hole are compared. It was found that the D-Fan and the SD-Fan have the lowest loss. This is attributed to their laterally expanded hole inlets, which lead to significant reduction in the loss generated inside the holes. As a result, the loss of these geometries is ≈ 50% of the loss of the fan-shaped hole at IR = 0.64 and 1.44.

Film cooling adiabatic effectiveness and heat transfer coefficients for cylindrical holes embedded in a 1d transverse trench on the suction side of a simulated turbine vane were investigated to determine the net heat flux reduction. For reference, measurements were also conducted with standard inclined, cylindrical holes. Heat transfer coefficients were determined with and without upstream heating to isolate the hydrodynamic effects of the trench and to investigate the effects of the thermal approach boundary layer. Also, the effects of a tripped versus an untripped boundary layer were explored. For both the cylindrical holes and the trench, heat transfer augmentation was much greater for the untripped approach flow. A further increase in heat transfer augmentation was caused by use of upstream heating, with as much as a 180% augmentation for the trench. The tripped approach flow led to much lower heat transfer augmentation than the untipped case. The net heat flux reduction for the trench was found to be significantly higher than for the row of cylindrical holes.

Effects of coolant supply arrangement on double wall cooling: Hot-side effusion performance and cold-side Nusselt numbers at different initial blowing ratios

An advanced, high-effectiveness film cooling design, the antivortex hole (AVH) has been investigated by several research groups and shown to mitigate or counter the vorticity generated by conventional holes and increase film effectiveness at high blowing ratios and low freestream turbulence levels. The effects of increased turbulence on an AVH geometry were previously investigated in a preliminary steady computational fluid dynamics (CFD) study by Hunley et al. on the film effectiveness and net heat flux reduction (NHFR) at high blowing ratio. The current paper presents the results of an extended numerical parametric study, which attempts to separate the effects of turbulence intensity and length scale on film cooling performance of the AVH concept at high blowing ratio (2.0) and density ratio (2.0). In this extended study, steady Reynolds-averaged Navier–Stokes (RANS) analysis was performed with turbulence intensities of 5, 10, and 20% and length scales based on cooling hole diameter of Λx/dm = 1, 3, and 6. Increasing turbulence intensity was shown to increase the centerline, span-averaged, and area-averaged adiabatic film cooling effectiveness and NHFR. Larger turbulent length scales in the steady RANS analysis were shown to have little to no effect on the centerline, span-averaged, and area-averaged adiabatic film cooling effectiveness and NHFR at lower turbulence levels, but moderate effect at the highest turbulence levels investigated. Heat transfer results were in good agreement with the findings from adiabatic cases from previous work. Unsteady RANS results also provided supplementary flow visualization for the AVH film cooling flow under varying turbulence levels.

Experimental investigations are performed on the suction side of a cooled turbine guide vane. Transient infrared thermography is used to evaluate film-cooling performance of cylindrical and fan-shaped holes in a test facility representing engine conditions. Adiabatic film effectiveness and net heat flux reduction results due to coolant injection through double and multiple rows in the presence and absence of an upstream showerhead are presented. Two double staggered rows at different positions have been cross compared: one at a relatively high convex curvature region and the other close to the maximum throat velocity. A combination of the two double rows is considered to be multiple rows. The tested blowing ratios are in the interval of [0.6–1.2] and [0.3–1.2] for double and multiple rows, respectively. The showerhead cooling is maintained at a nominal blowing ratio. The findings suggest that the choice of best cooling hole shape for film-cooling design can be highly influenced by the number of cooling rows to be used and also the presence (or absence) of showerhead cooling. It is worth noting that the outcome may differ depending on the quantity of interest (i.e., adiabatic film effectiveness or net heat flux reduction).

A novel turbine film cooling hole shape has been conceived and designed at NASA Glenn Research Center. This “anti-vortex” design is unique in that it requires only easily machinable round holes, unlike shaped film cooling holes and other advanced concepts. The hole design is intended to counteract the detrimental vorticity associated with standard circular cross-section film cooling holes. This vorticity typically entrains hot freestream gas and is associated with jet separation from the turbine blade surface. The anti-vortex film cooling hole concept has been modeled computationally for a single row of 30 degree angled holes on a flat surface using the 3D Navier-Stokes solver Glenn-HT. A blowing ratio of 1.0 and density ratios of 1.05 and 2.0 are studied. Both film effectiveness and heat transfer coefficient values are computed and compared to standard round hole cases for the same blowing rates. A net heat flux reduction is also determined using both the film effectiveness and heat transfer coefficient values to ascertain the overall effectiveness of the concept. An improvement in film effectiveness of about 0.2 and in net heat flux reduction of about 0.2 is demonstrated for the anti-vortex concept compared to the standard round hole for both blowing ratios. Detailed flow visualization shows that as expected, the design counteracts the detrimental vorticity of the round hole flow, allowing it to remain attached to the surface.

Analysis of Multiphase Heat Transfer of TA2/Q235B Clad Plate Subjected to Impinging Liquid Jet Cooling

Aerothermal behavior and performance of turbine blade squealer tip has been extensively studied, but previously mainly confined to very low speed flows. Some recent research efforts have revealed some distinctive tip heat transfer as well as aerodynamic characteristics in a transonic flow, many of which are confined to a linear cascade setting. There is a need to identify driving mechanisms for such configurations and to quantify corresponding performance characteristics in general, and in the process address how those drivers would relate to some specific issues of interest such as: 1) the differences between a cascade and a rotor tip, thus the impact of relative casing movement? 2) The difference between a transonic and a high subsonic flow condition? 3) the difference between a squealer tip and flat tip? In this study, the vortical structures and the loss mechanism of the Over Tip Leakage (OTL) flow in a high-subsonic stage and a transonic stage are analyzed using a proposed vortex model and Denton’s loss model. The two stages are specifically designed to work in the same Reynolds number with the efficiency comparable to the one predicted in the Smith’s chart. Two counter-rotating vortices, the Pressure-side Separation Vortex (PSV) and the Casing-driven Cavity Vortex (CCV), are identified as key drivers as inside a squealer cavity strongly impacting aerothermal performance (OTL mass flow and the heat transfer distribution). The cavity flow pattern results from the balancing between the CCV primarily driven by the relative casing movement, and the PSV originating from the flow separation around the pressure side rim. Consequently a detoured flow path is created within the squealer cavity when the leakage flow separating on the pressure side is entrained by the two vortices to the squealer bath and finally leaves the cavity along the suction side of the squealer wall. The entrainment of the leakage flow induces a signature of a high HTC strip on the squealer floor. Without the relative casing movement, the CCV is absent in the majority of the cavity so that the leakage flow can directly cross the tip which increases the leakage mass flow rate and reduce the HTC value on the cavity floor. Due to the distinct difference in vortical structures, the change of the HTC with and without relative casing movement can be interpreted by the vortex model. The vortex model is also able to explain the variation of the leakage mass flow rate and the HTC contours with the increasing squealer height. From the best knowledge of the authors, this is the first-of-its kind effort in using in the cavity vortices to systematically and quantitatively study the effect of the relative casing movement and the aerothermal performance of the tip with varying squealer height. In counting the aero losses for the two stages, three sources are attributed to the total loss generation: the baseline loss which assumes there is no leakage, the in-tip loss which is generated within the tip gap and the out-tip loss which accounts for the mixing loss between the OTL flow and the main passage flow. The out-tip loss is found to be the major source of the loss caused by the tip leakage flow. Denton’s loss model is used to calculate the out-tip loss quantitatively. The model predicted loss agrees in trend with the one calculated from the CFD results. According to the model, the out-tip loss is related to the main flow Mach number, the OTL mass flow rate and the velocity difference between the main flow and the OTL flow. From the distribution of these terms along the tip, the loss is mainly generated at the rear portion of the tip. The reduction of the leakage mass flow rate by the squealer is mainly at the zones where the CCV and the PSV are leaving the squealer cavity. The present analysis and results indicate that the squealer in the transonic stage tends to be less effective in reducing the out-tip loss compared to the subsonic one. The region of the tip where the CCV has already left the cavity with the PSV remaining shows the highest entropy generation rate, due to locally both the high mainstream velocity and the high OTL mass flow given the PSV being left as the only sealing vortex there.

A numerical study of anti-vortex film-cooling holes designs in a 1-1/2 turbine stage using LES

This experiment investigates the effects of blowing ratio on the film cooling performance of a showerhead-cooled first-stage turbine vane at low freestream turbulence (Tu = 2%) and an integral length scale normalized by vane pitch (Λx/P) of 0.05. The exit Reynolds number based on vane true chord is 1.1 × 106. The effect of freestream turbulence at high Mach number (Mex = 0.76) and blowing ratio (BR = 0, 1.5, 2.0) is also explored by comparing results with high freestream turbulence measurements (Tu = 16%) previously performed in the same cascade. To characterize film cooling performance, platinum thin-film gauges were used to measure Nusselt number and film cooling effectiveness distributions at the midspan of the vane. Net heat flux reduction is also addressed. The primary effects of coolant injection were augmentation of Nusselt number and reduction of adiabatic wall temperature on the vane surface. In general, increasing blowing ratio showed increases in Nusselt number augmentation over the vane surface and an increase in film cooling effectiveness as well. Both Nusselt number and film cooling effectiveness trends were influenced by a strong favorable pressure gradient and resulting flow acceleration on the suction surface. Comparing low freestream turbulence results with high freestream turbulence measurements showed that large-scale, high freestream turbulence can decrease heat transfer coefficient and film cooling effectiveness downstream of injection.

This paper presents an experimental investigation of the film cooling performance of a New Hybrid Film Cooling Scheme using Thermochromic Liquid Crystal technique. The new scheme has been designed to improve the film cooling performance of gas turbine airfoils. The scheme includes two consecutive film hole configurations with interior bending. The cooling performance of the new scheme was analyzed across blowing ratios of 0.5, 1.0 and 1.5, at a density ratio of 0.94. The results showed that the new scheme enhanced the local and the laterally averaged film cooling performance in terms of effectiveness, and net heat flux reduction in compared to other film hole configurations. The bending effect of the new scheme throttled the secondary flow causing it to spread widely over the downstream surfaces, hence enhancing the film cooling performance at low and high blowing ratios. The hybrid scheme provided an average heat transfer ratio near unity over the downstream surfaces at low and high blowing ratios.

This study uses transient IR-thermography to evaluate the effect of showerhead cooling and hole position on the performance of single-row cooling hole on the pressure side of a guide vane under engine representative conditions. The investigation includes both cylindrical and fan-shaped holes at two blowing conditions: 0.6 and 1.8. The influence of cooling hole alignment for these hole shapes in the performance of multiple row configurations was also studied in the presence of showerhead. For this purpose, double- and triple-row cases in staggered and non-staggered arrangements were considered for two blowing conditions, similar to the single row. The results are presented in terms of both adiabatic film effectiveness, AFE, and net heat flux reduction, NHFR. The showerhead effect was shown to be profound with regard to both AFE and NHFR for the cooling hole close to it. This holds for both hole shapes and blowing ratios. The overall film cooling performance, NHFR, of the rows further downstream of the showerhead and close to the trailing edge were affected marginally by the showerhead. The later cooling row showed superior performance compared to the other rows for fan-shaped holes in both presence and absence of shower-head at a low blowing ratio. For multiple row configurations, in general fan-shaped holes can maintain higher AFE in staggered alignment, while cylindrical holes benefit from consequent jet interaction between rows of cooling in a non-staggered arrangement. This holds for both investigated blowing ratios and double- and triple-rows. When considering NHFR, the results indicate that fan-shaped holes are less affected by the hole alignment. Cylindrical holes, however, can maintain superior performance in non-staggered alignment for all investigated cases except triple row under low blowing condition. The results also suggest that a double-row configuration in the presence of showerhead will benefit from an additional row mainly at high blowing ratios.