An official website of the United States government
While modern gas turbine engines operate at hot gas path velocities approaching the speed of sound, few facilities have studied the effects that the flow’s compressibility can have on the adiabatic effectiveness. A new facility at the University of Texas at Austin has been developed to investigate these high Mach number effects and how to appropriately scale laboratory film cooling experiments to engine conditions. This study investigates two film cooling hole geometries, a baseline 7-7-7 shaped film cooling hole and a recent design which has been numerically optimized for increased effectiveness. Both holes are tested at mainstream Mach numbers of 0.25 and 0.50 in a flat plate test section. The optimized hole outperforms the effectiveness of the baseline geometry at all blowing ratios tested, matching the trend in the results of previous studies on these geometries. However, there is a marked decrease in film cooling hole performance as the Mach number is increased.
more »
« less
Considerations for Compressible Film Cooling: A Computational Study of the Effects of Transonic Flows and Varying Mainstream Mach Number
Recent evidence suggests that film cooling flows with engine realistic mainstream Mach number have declined performance in comparison to those with conventional low-speed laboratory conditions. Consideration and understanding of these effects are fundamental to improving film cooling research. The proposed computational study investigates the film cooling performance of a 7-7-7 shaped film cooling hole with respect to varying mainstream Mach number, with constant Reynolds number. The cases studied include mainstream Mach numbers from 0.15–0.75, with a fixed, engine realistic, hole Reynolds number of Red = 10, 100. Significant results are then evaluated against varying stagnation temperature ratio and blowing ratio. The results showed that at a blowing ratio of 1.75, the adiabatic effectiveness declines significantly with high mainstream Mach number. The decreased performance is due to supersonic flows and shocks in the film cooling hole that disrupt flow in the diffuser section of the hole. These characteristics are observed across all stagnation temperature ratios considered. In addition to these insights, the study discusses the importance of proper thermal scaling and definition of adiabatic effectiveness when operating at high mainstream Mach number. It is demonstrated that the effects of high-speed flow challenge the efficacy of the conventional parameters used to characterize film cooling performance.
more »
« less
- Award ID(s):
- 1936676
- PAR ID:
- 10466388
- Date Published:
- Journal Name:
- ASME Turbo Expo 2023: Turbomachinery Technical Conference and Exposition
- Format(s):
- Medium: X
- Sponsoring Org:
- National Science Foundation
More Like this
-
-
NA (Ed.)Presented are experimental and analytic procedures for the measurement and determination of film cooling performance parameters for a transonic flow environment along a turbine blade tip, which account for the influences of viscous dissipation. Such viscous dissipation magnitudes are vital to ascertain appropriate driving temperatures for convective heat transfer within high velocity, compressible environments. A key step in this procedure is the separation of adiabatic surface temperature magnitudes due to the thermal field from adiabatic surface temperature values associated with flow effects related to viscous dissipation. These latter adiabatic surface temperature magnitudes, from flow effects only, are determined with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface. After the separation is implemented, adiabatic surface temperature variations from film cooling thermal effects only are provided relative to local baseline values with no film cooling. Resulting local adiabatic film cooling effectiveness and local heat transfer coefficients are subsequently determined which are based entirely upon locally measured temperatures (without the use of global temperature parameters), which is the most appropriate perspective for presentation of spatially-resolved surface heat transfer characteristics. Associated data are provided for the surface of a blade tip with a squealer rim and recess, with and without film cooling. An array of five film cooling holes, located along the upper pressure side of a two-dimensional turbine blade, are employed to provide the film coolant, with a local blowing ratio BR of 3.02. Ratios of tip gap to true blade span are 0.66 percent, and 1.16 percent. Experimental procedures employed to obtain the spatially-resolved surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.more » « less
-
A novel, double hole film cooling configuration is investigated as an alternative to traditional cylindrical and fanshaped, laidback holes. This experimental investigation utilizes a Stereo-Particle Image Velocimetry (S-PIV) to quantitatively assess the ability of the proposed, double hole geometry to weaken or mitigate the counter-rotating vortices formed within the jet structure. The three-dimensional flow field measurements are combined with surface film cooling effectiveness measurements obtained using Pressure Sensitive Paint (PSP). The double hole geometry consists of two compound angle holes. The inclination of each hole is = 35°, and the compound angle of the holes is = ± 45° (with the holes angled toward one another). The simple angle cylindrical and shaped holes both have an inclination angle of = 35°. The blowing ratio is varied from M = 0.5 to 1.5 for all three film cooling geometries while the density ratio is maintained at DR = 1.0. Time averaged velocity distributions are obtained for both the mainstream and coolant flows at five streamwise planes across the fluid domain (x/d = -4, 0, 1, 5, and 10). These transverse velocity distributions are combined with the detailed film cooling effectiveness distributions on the surface to evaluate the proposed double hole configuration (compared to the traditional hole designs). The fanshaped, laidback geometry effectively reduces the strength of the kidney-shaped vortices within the structure of the jet (over the entire range of blowing ratios considered). The three-dimensional velocity field measurements indicate the secondary flows formed from the double hole geometry strengthen in the plane perpendicular to the mainstream flow. At the exit of the double hole geometry, the streamwise momentum of the jets is reduced (compared to the single, cylindrical hole), and the geometry offers improved film cooling coverage. However, moving downstream in the steamwise direction, the two jets form a single jet, and the counter-rotating vortices are comparable to those formed within the jet from a single, cylindrical hole. These strong secondary flows lift the coolant off the surface, and the film cooling coverage offered by the double hole geometry is reduced.more » « less
-
This work presents recent advancements in the study of film cooling in hypersonic flows, considering experimental and numerical investigations, with the aim to characterize the wall-cooling performance in different coolant injection and baseflow conditions in a Mach number range 2–7.7. The study explores the mutual interaction between the injected coolant film and the boundary-layer flow, with emphasis on the effects of wall blowing on the boundary-layer characteristics, stability, and transition to turbulence, as well as the effect of transition on wall-cooling performance. Considered flow configurations include cases of effusion cooling in both wall-normal or slightly inclined and wall-parallel blowing, different types of coolant, cases of favorable pressure gradient compared to zero pressure gradient, as well as transpiration cooling cases at different blowing ratios and surface geometries. For the transpiration cooling case, experiments in different hypersonic wind tunnel facilities are presented for flat plate and cone geometries, with coolant injected through C/C porous samples, whereas numerical simulations of modeled porous injection are presented for a flat plate and a blunt cone, showing results for the boundary-layer receptivity with coolant injection and the associated effects on transition and cooling performance. A summary of the main findings is provided along with a critical analysis based on a comparative study to evaluate the effect of each configuration, injection strategy, and key parameters on the boundary-layer flow and the feedback on wall-cooling performance. Conclusions are drawn about potential directions of study for the further development and optimization of the film cooling technique for future hypersonic vehicles.more » « less
-
We employ numerically implicit subgrid-scale modeling provided by the well-known streamlined upwind/Petrov–Galerkin stabilization for the finite element discretization of advection–diffusion problems in a Large Eddy Simulation (LES) approach. Whereas its original purpose was to provide sufficient algorithmic dissipation for a stable and convergent numerical method, more recently, it has been utilized as a subgrid-scale (SGS) model to account for the effect of small scales, unresolvable by the discretization. The freestream Mach number is 2.5, and direct comparison with a DNS database from our research group, as well as with experiments from the literature of adiabatic supersonic spatially turbulent boundary layers, is performed. Turbulent inflow conditions are generated via our dynamic rescaling–recycling approach, recently extended to high-speed flows. Focus is given to the assessment of the resolved Reynolds stresses. In addition, flow visualization is performed to obtain a much better insight into the physics of the flow. A weak compressibility effect is observed on thermal turbulent structures based on two-point correlations (IC vs. supersonic). The Reynolds analogy (u′ vs. t′) approximately holds for the supersonic regime, but to a lesser extent than previously observed in incompressible (IC) turbulent boundary layers, where temperature was assumed as a passive scalar. A much longer power law behavior of the mean streamwise velocity is computed in the outer region when compared to the log law at Mach 2.5. Implicit LES has shown very good performance in Mach 2.5 adiabatic flat plates in terms of the mean flow (i.e., Cf and UVD+). iLES significantly overpredicts the peak values of u′, and consequently Reynolds shear stress peaks, in the buffer layer. However, excellent agreement between the turbulence intensities and Reynolds shear stresses is accomplished in the outer region by the present iLES with respect to the external DNS database at similar Reynolds numbers.more » « less
- Free Publicly Accessible Full Text
-
Accepted Manuscript1.0
- Conference Paper:
- https://doi.org/10.1115/GT2023-104049
