Adiabatic Film Research Articles

An Experimental Study of Mist/Air Film Cooling With Fan-Shaped Holes on an Extended Flat Plate—Part 1: Heat Transfer

Motivated by the need to further improve film cooling in terms of both cooling effectiveness and coolant coverage area, the mist/air film cooling scheme is investigated through experiments using fan-shaped holes over an extended downstream length in this study. Both an existing wind tunnel and test facility, used in previous work, have been retrofitted. The first modification was extending the length of the flat plate test section to cover longer distances downstream of the injection holes, up to X/D = 100, in order to investigate whether mist cooling can be harnessed farther downstream where single-phase film cooling is not effective. X represents the axial distance downstream of the cooling hole of diameter D. The second modification was to incorporate fan-shaped (diffusion) holes which are proven to have a higher film cooling efficiency, than cylindrical holes. The objective is to investigate whether mist can further enhance the film cooling performance of the already highly effective fan-shaped holes. A phase Doppler particle analyzer (PDPA) system is employed to measure the droplet size, velocity, and turbulence information. An infrared camera and thermocouples are both used for temperature measurements. Part I is focused on the heat transfer result on the wall. The results show that, at low blowing ratios when the film is attached to the surface, the enhancement of the mist film cooling effectiveness, compared to the air-only case, on the centerline of the hole ranges from 40% in the near hole region to over 170% at X/D = 100. Due to the diffusive nature of the fan-shaped hole, the laterally averaged enhancement is on par with that on the centerline. The significant enhancement over the extended downstream distance from X/D = 40–100 is attributed to the evaporation time needed to evaporate all of the droplets. Each droplet acts as a cooling sink and flies over a distance before it completely vaporizes. This “distributed cooling” characteristic allows the water droplets to extend the cooling effects farther downstream from the injection location. At higher blowing ratios, when the cooling film is lifted off from the surface, the cooling enhancement drops below 40%. Although the enhancement in the near hole region X/D < 40 is about 20% lower than that achieved by using the cylindrical holes, the magnitudes of the mist adiabatic film cooling effectiveness using fan-shaped holes are still much higher than those of the cylindrical holes. Part II of this study is focused on analyzing the two-phase droplet multiphase flow behavior to explain the fundamental physics involved in the mist film cooling.

Double Wall Cooling of a Full-Coverage Effusion Plate, Including Internal Impingement Array Cooling

New experimental data are provided for full-coverage effusion cooling and impingement array cooling, as applied simultaneously onto the respective external and internal surfaces of a single instrumented test plate. For the effusion cooled surface, presented are spatially resolved distributions of surface adiabatic film cooling effectiveness, and surface heat transfer coefficients. For the impingement cooled surface, presented are spatially resolved distributions of surface Nusselt numbers. Impingement jet arrays at different jet Reynolds numbers, from 7930 to 18,000, are employed. Experimental data are given for spanwise and streamwise impingement hole spacing such that coolant jet hole centerlines are located midway between individual effusion hole entrances. For the effusion cooling, streamwise hole spacing and spanwise hole spacing (normalized by effusion hole diameter) are 15 and 4, respectively. Effusion hole angle is 25 deg, and effusion plate thickness is 3.0 effusion hole diameters. In regard to the impingement cooled cold-side surface of the effusion plate, associated surface Nusselt number variations provide evidence that impingement jets are turned and redirected as they cross the impingement passage, just prior to the entrance of coolant into individual effusion holes. In regard to the effusion cooled hot-side surface of the effusion plate, when compared at particular values of injectant and mainstream Reynolds numbers, streamwise location x/de and blowing ratio BR, significantly increased thermal protection is provided when the effusion coolant is provided by an array of impingement cooling jets (compared to a cross flow channel supply arrangement).

High resolution measurements of film cooling performance of simple and compound angle cylindrical holes with varying hole length-to-diameter ratio—Part I: Adiabatic film effectiveness

Adiabatic film cooling effectiveness on a flat plate surface downstream a row of simple and compound angle holes was investigated using high-resolution pressure sensitive paint (PSP) technique. A variation of flow parameters including density ratio and blowing ratio, and geometry parameters including compound angle and length-to-diameter ratio were examined. Blowing ratios (M) ranging from 0.3 to 2, density ratio (DR) are 1.0 and 1.5, length to diameter ratios (L/D) from 0.5 to 5 and two compound angle (β: 0°, 45°) were employed composing a test matrix of 140 test cases. Discharge coefficient was measured with varied blowing ratio and L/D values for simple and compound angle hole. Detailed local, laterally averaged, and area averaged film effectiveness are presented to illustrate the effect of length-to-diameter ratio and compound angle. It was found that the effect of density ratio on the film cooling effectiveness was associated with the blowing ratio. Increasing density ratio provides a negative effect when the M ≤ 0.5 for simple and compound angle hole, which a positive effect when the M > 0.5. Also found was that the film effectiveness was highest when L/D = 0.5 and 1, regardless of blowing ratio and compound angle. The film cooling effectiveness of simple angle hole increased as L/D increased from 2 to 5 when M ≤ 1, while showed a minimum value at L/D = 3.5 when M ≥ 1. The film cooling effectiveness of compound angle hole showed a minimum value at L/D = 3.5 regardless of blowing ratio. The scaling methods of film effectiveness with velocity ratio, blowing ratio, and momentum flux ratio were also discussed for simple angle and compound angle holes with varying L/D.

Internal and external cooling of a full coverage effusion cooling plate: Effects of double wall configuration and conditions

The present study provides new heat transfer data for both the surfaces of the full coverage effusion cooling plate within a double wall cooling test facility. To produce the cooling stream, a cold-side cross-flow supply for the effusion hole array is employed. Also utilized is a unique mainstream mesh heater, which provides transient thermal boundary conditions, after mainstream flow conditions are established. For the effusion cooled surface, presented are spatially-resolved distributions of surface adiabatic film cooling effectiveness, and surface heat transfer coefficients (measured using infrared thermography). For the coolant side, presented are spatially-resolved distributions of surface Nusselt numbers (measured using liquid crystal thermography). Of interest are the effects of streamwise development, blowing ratio, and Reynolds number. Streamwise hole spacing and spanwise hole spacing (normalized by effusion hole diameter) on the effusion plate are 15 and 4, respectively. Effusion hole diameter is 6.35 mm, effusion hole angle is 25°, and effusion plate thickness is 3 hole diameters. Considered are overall effusion blowing ratios from 2.9 to 7.5, with subsonic, incompressible flow, and constant freestream velocity with streamwise development, for two different mainstream Reynolds numbers. For the hot side (mainstream) of the effusion film cooling test plate, results for two mainflow Reynolds numbers of about 145000 and 96000 show that the adiabatic cooling effectiveness is generally higher for the lower Reynolds number for a particular streamwise location and blowing ratio. The heat transfer coefficient is generally higher for the low Reynolds number flow. This is due to altered supply passage flow behavior, which causes a reduction in coolant lift-off of the film from the surface as coolant momentum, relative to mainstream momentum, decreases. For the coolant side of the effusion test plate, Nusselt numbers generally increase with blowing ratio, when compared at a particular streamwise location and mainflow Reynolds number.

Film Cooling Performances at Different Turbulence Intensities Using Conjugate Heat Transfer Analysis

This study presents a numerical investigation of cooling performances of a modified vane of the film-cooled vane reported by Timko (NASA CR-168289) at different mainstream turbulence intensities (Tus). A 3D conjugate heat transfer (CHT) analysis with SST k-ω turbulence model in FLUENT V.15 is used. Three different mechanisms in CHT analysis, i.e. fluid flow, heat convection between solid surfaces and flowing fluid in an external mainstream and internal cooling passages, and heat conduction within the vane structure, are simultaneously considered. Numerical results are conducted in terms of overall cooling effectiveness at Tu=3.3, 10, and 20%. Comparison between overall cooling effectiveness and film effectiveness under adiabatic assumption is discussed at the three Tus, also. The findings of this research indicate the following phenomena: 1) overall cooling effectiveness decreases with Tu, and this effect on the pressure side (PS) is stronger than that on the suction side (SS) in general. 2) By comparison with adiabatic film effectiveness, the level of overall cooling effectiveness in most regions is higher and more uniform than that of adiabatic film effectiveness for all three Tus. 3) In the leading edge (LE), when Tu increases, near the exits of film holes overall cooling effectiveness deteriorates, but adiabatic film effectiveness improves. Furthermore, a large area with relatively low overall cooling effectiveness is able to move with Tu in the LE region.

Experimental and Computational Study of the Effect of Momentum-Flux Ratio on High-Pressure Nozzle Guide Vane Endwall Cooling Systems

High-pressure (HP) nozzle guide vane (NGV) endwalls are often characterized by highly three-dimensional (3D) flows. The flow structure depends on the incoming boundary layer state (inlet total pressure profile) and the (static) pressure gradients within the vane passage. In many engine applications, this can lead to strong secondary flows. The prediction and design of optimized endwall film cooling systems is therefore challenging and is a topic of current research interest. A detailed experimental investigation of the film effectiveness distribution on an engine-realistic endwall geometry is presented in this paper. The film cooling system was a fairly conventional axisymmetric double-row configuration. The study was performed on a large-scale, low-speed wind tunnel using infrared (IR) thermography. Adiabatic film effectiveness distributions were measured using IR cameras, and tests were performed across a wide range of coolant-to-mainstream momentum-flux and mass flow ratios (MFRs). Complex interactions between coolant film and vane secondary flows are presented and discussed. A particular feature of interest is the suppression of secondary flows (and associated improved adiabatic film effectiveness) beyond a critical momentum flux ratio. Jet liftoff effects are also observed and discussed in the context of sensitivity to local momentum flux ratio. Full coverage experimental results are also compared to 3D, steady-state computational fluid dynamics (CFD) simulations. This paper provides insights into the effects of momentum flux ratio in establishing similarity between cascade conditions and engine conditions and gives design guidelines for engine designers in relation to minimum endwall cooling momentum flux requirements to suppress endwall secondary flows.

Development of PSP Technique for Vane Film Cooling Investigations

The present paper reports on the development of dual-luminophor Pressure Sensitive Paints (PSP) technique at University of Bergamo and on its application to the measurement of adiabatic film cooling effectiveness. An in-house calibration was carried out in order to assess the influence of temperature on PSP signal. Binary PSP have been then applied on a film cooled flat plate model for validation purposes and then used to measure the adiabatic film cooling effectiveness distribution over the leading and the trailing edge region of a high loaded high pressure nozzle guide vane. In both cases, PSP results have been compared with Liquid Crystal (TLC) data. Unlike the heat-transfer based techniques (i.e. TLC), binary PSP results are not affected by conduction phenomena, allowing to comply with adiabatic condition over test duration. PSP allowed to obtain highly resolved and high quality distributions even in the near hole exit region and where compressibility effects are relevant.

On the Flow Structures and Adiabatic Film Effectiveness for Simple and Compound Angle Hole With Varied Length-to-Diameter Ratio by Large Eddy Simulation and Pressure-Sensitive Paint Techniques

The effects of hole length-to-diameter ratio and compound angle on flat plate film cooling effectiveness are investigated from an experimental and numerical view. Film cooling effectiveness measurements are performed for seven blowing ratios (M) ranging from 0.3 to 2, five-hole length-to-diameter ratios (L/D) from 0.5 to 5, and two compound angles (β: 0 deg and 45 deg) using pressure-sensitive paint (PSP) technique. Results indicate that discrete holes with L = 0.5 and 1 show highest film cooling effectiveness regardless of compound angle. Round hole generally shows an increasing trend as L increases from 2 to 5, while compound angle hole shows a complex trend concerning with blowing ratios (BRs) and length-to-diameter ratios. Compound angle enhances film cooling effectiveness with high blowing ratios and length-to-diameter ratios. In a parallel effort, large eddy simulation (LES) approach is employed to solve the flow field and visualize vortex structures of intube and mainstream regions. It is demonstrated that the counter-rotating vortex pair (CRVP) which is observed in the time-averaged flow field is originated in different vortex structures with varying blowing ratios and length-to-diameter ratios. Scalar field transportation features are also investigated to clarify how different vortex structures affect the temperature distribution and the film cooling effectiveness.

The effect of effusion holes inclination angle on the adiabatic film cooling effectiveness in a three-sector gas turbine combustor rig with a realistic swirling flow

The introduction of Lean Burn concept as basic Low-NOx scheme for future aero-engines is heavily affecting the aero-thermal design of combustors. A great amount of air is admitted through the injection system with relevant swirl components, producing very complex flow structures (recirculations, vortex breakdown) for flame stabilization. As a consequence a reduced quantity of air is available for liner cooling, pushing the adoption of high effectiveness cooling schemes. Effusion cooling represents one of the first choices due to its low weight and a relatively easy manufacturability. Liner metal temperature is kept low by the combined protective effect of coolant film, heat removal inside holes and an improved cold-side convection. In lean burn systems the evolution of film protection can be heavily influenced by the swirl flow interaction with combustor walls.The subject of this work is to investigate the effects of the realistic flow field of a lean burn injector on the adiabatic film cooling effectiveness on an effusion cooled combustor liner. A dedicated three-sector rig was designed with the aim of measuring film effectiveness with Pressure Sensitive Paint technique. Three effusion cooling geometries with different inclination angles were tested at various levels of pressure drops across the perforation, resulting in different blowing ratio values. It was also taken into consideration several flow rate levels of starter film realized by spent dome cooling air, injected through a dedicated plain slot. The analysis of film effectiveness measurements were supported by flow field investigation in the near wall region carried out by means of Particle Image Velocimetry.Results pointed out the relevant impact of combustor flow field on the adiabatic film cooling effectiveness as well as a significant role of the inclination angle, recommending a careful revision of standard design practices based on one dimensional flow assumption and suggesting possible holes arrangement optimization.

Numerical investigation on the effects of the divided steps on film cooling performance

Described in this paper is a numerical investigation on the concept for enhancing the film cooling performance by placing divided steps upstream the film cooling hole. Five different divided steps are investigated, including the cases of steps with 2.5mm, 5mm, 10mm, 15mm divided spacing in the middle, and 2.5mm divided spacing in both lateral sides. The effects of different divided spacing on film cooling performance are conducted. The film cooling performance on a flat plate and the film cooling with an undivided step are also presented as a contrast. The film cooling performance is evaluated at the density of 0.97 with the blowing ratios ranging from 1.0 to 2.0. Results obtained show that the film cooling performance with the steps divided in the middle is greatly improved when compared with the undivided step. Both the entrainment of coolant and the anti-kidney vortexes are observed in the case of film cooling with divided steps and the film cooling performance is improved by the combined effect of the entrainment and the anti-kidney vortexes. The total pressure loss for film cooling with divided steps is lower than that with undivided step. The film coverage effect and the adiabatic film cooling effectiveness for case with 5mm divided in the middle of the step are all the best among all the cases.

Numerical study and GMDH-type neural networks modeling of plasma actuator effects on the film cooling over a flat plate

In this paper, the film cooling effectiveness for cylindrical and fan-shaped holes is numerically investigated in a 3-D model on the surface of a flat plate with and without the attendance of a plasma actuator. The electric force produced by a plasma actuator is employed as a body force into a finite volume based Navier–Stokes solver. Incompressible, viscous and turbulent flow has been simulated by the Fluent CFD code with the low-Re k-ε turbulence model. The results show that the fan-shaped holes provide higher film cooling effectiveness than cylindrical holes. The influences of flow, geometric and electrical variables on the performance of film cooling have been discussed. Finally, the group method of data handling (GMDH)-type neural networks has been used for presenting a relationship for area weighted average adiabatic film cooling effectiveness. Predicted results for film cooling effectiveness showed that GMDH model can provide beneficial information for simulation of film cooling process.

Adiabatic Film Cooling Effectiveness Measurements Throughout Multirow Film Cooling Arrays

Adiabatic film cooling effectiveness measurements are obtained using pressure-sensitive paint (PSP) on a flat film cooled surface. The effects of blowing ratio and hole spacing are investigated for four multirow arrays comprised of eight rows containing 52 holes of 3.8 mm diameter with 20 deg inclination angles and hole length-to-diameter ratio of 11.2. The four arrays investigated have two different hole-to-hole spacings composed of cylindrical and diffuser holes. For the first case, lateral and streamwise pitches are 7.5 times the diameter. For the second case, pitch-to-diameter ratio is 14 in lateral direction and 10 in the streamwise direction. The holes are in a staggered arrangement. Adiabatic effectiveness measurements are taken for a blowing ratio range of 0.3–1.2 and a density ratio of 1.5, with CO2 injected as the coolant. A thorough boundary layer analysis is presented, and data were taken using hotwire anemometry with air injection, with boundary layer, and turbulence measurements taken at multiple locations in order to characterize the boundary layer. Local effectiveness, laterally averaged effectiveness, boundary layer thickness, momentum thickness, turbulence intensity, and turbulence length scale are presented. For the cylindrical holes, at the first row of injection, the film jets are still attached at a blowing ratio of 0.3. By a blowing ratio of 0.5, the jet is observed to lift off, and then impinge back onto the test surface. At a blowing ratio of 1.2, the jets lift off, but reattach much further downstream, spreading the coolant further along the test surface. A thorough uncertainty analysis has been conducted in order to fully understand the presented measurements and any shortcomings of the measurement technique. The maximum uncertainty of effectiveness and blowing ratio is 0.02 counts of effectiveness and 3%, respectively.

Aerothermal Performance of a Nozzle Vane Cascade With a Generic Nonuniform Inlet Flow Condition—Part II: Influence of Purge and Film Cooling Injection

In the present paper, the influence of the presence of an inlet flow nonuniformity on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out with platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. An obstruction was installed upstream of the cascade at variable tangential and axial position to generate a flow nonuniformity. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition. Aerothermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Results show a relevant negative impact of inlet flow nonuniformity on the cooled cascade aerodynamic and thermal performance. Higher film cooling effectiveness and lower aerodynamic losses are obtained when the inlet flow nonuniformity is located at midpitch, while lower effectiveness and higher losses are obtained when it is aligned to the vane leading edge. Moving the nonuniformity axially or changing its blockage only marginally influences the platform thermal protection.

Experimental Investigation of Endwall Heat Transfer With Film and Impingement Cooling

The switch from diffusive combustion to premixed combustion in a modern gas turbine changes the combustor exit temperature profile to a more uniform one. This will directly affect the cooling of the first stage vane especially the endwall region. A typical endwall configuration with matched nondimensional parameters to the engine condition was investigated experimentally in this study. Two endwall cooling arrangements at four different coolant to mainstream mass flow ratios (MFR) were tested in a linear cascade. Detailed measurements of pressure distribution, heat transfer coefficient, adiabatic film cooling effectiveness, and overall effectiveness of the endwall were performed. The temperature-sensitive paint (TSP) and pressure-sensitive paint (PSP) were used to acquire these parameters. The conjugate heat transfer characteristic of endwall with film cooling and impingement cooling was discussed. Moreover, the influence of coolant mass flow rate on conjugate heat transfer of endwall was analyzed. One- and two-dimensional methods for overall effectiveness prediction based on experimental data for separate parameters and correlations were also studied.

Experimentally Measured Effects of Incidence Angle on the Adiabatic and Overall Effectiveness of a Fully Cooled Turbine Airfoil With Shaped Showerhead Holes

Manufacturing and assembly variation can lead to shifts in the inlet flow incidence angles of a rotating turbine airfoil row. Understanding the sensitivity of the adiabatic film cooling effectiveness to a range of inlet conditions is necessary to verify the robustness of a cooling design. In order to investigate the effects of inlet flow incidence angles, adiabatic and overall effectiveness data were measured in a low speed linear cascade at 0 deg and 10 deg of the designed operating condition. Tests were completed at an inlet Reynolds number of Re = 120,000 and a turbulence intensity of Tu = 5% at the leading edge of the test article. Particle image velocimetry was used to verify the incident flow angle for each angle studied. The test section was first adjusted so that the pressure distribution and stagnation line of the airfoil matched those predicted by an aerodynamic computational fluid dynamics (CFD) model. IR thermography was then used to measure the adiabatic effectiveness levels of the fully cooled airfoil model with nine rows of shaped holes of varying construction and feed delivery. Measurements were taken over a range of blowing ratios and at a density ratio of DR = 1.23. This process was repeated for the two incidence angles measured, while the inlet pressure to the airfoil model was held constant for these incidence angle changes. Differences in laterally adiabatic effectiveness across the airfoil model were most evident in the showerhead, with changes as large as 0.2. The effect persisted most strongly at s/D = ±35 downstream of the stagnation row of holes, but was visible over the whole viewable area of 160 s/D. The effect was due to the stagnation line affecting the film at the showerhead row. Due to this effect, the showerhead was investigated in detail, including the effects of the stagnation line shift as well as the influence of the incidence angle on the overall effectiveness of the showerhead region. It was found that the stagnation line has the tendency to dramatically increase the near-hole adiabatic effectiveness levels when positioned within the breakout footprint of the hole. The effect persisted for the overall effectiveness study, since the hole spacing for this particular configuration was wide enough that the through hole convection was not completely dominant. This is the first study to present measured effectiveness values over both the pressure- and suction-side surfaces of a fully cooled airfoil for appreciably off-nominal incidence angles as well as examine adiabatic and overall effectiveness levels for a conical stagnation row of holes.

A novel sand-dune-inspired design for improved film cooling performance

An experimental study was performed to explore a novel film cooling design with a Barchan-Dune-Shaped Injection Compound (BDSIC) concept for improved film cooling performance over a test plate. Inspired by the unique crescent-shaped Barchan dunes commonly seen in deserts, the BDSIC concept was proposed to enable coolant streams to be discharged tangentially from hollowed Barchan-dune-shaped (BDS) shells and stay attached to the surface of the test plate. During the experiments, while a high-resolution Particle Image Velocimetry (PIV) system was used to conduct detailed flow field measurements to quantify the dynamic mixing process between the coolant stream and the mainstream flows over the test plate, a Pressure Sensitive Paint (PSP) technique was used to map the corresponding adiabatic film cooling effectiveness over the surface of interest based on a mass-flux analog to traditional temperature-based cooling effectiveness measurements. The effects of various important design parameters, including the shape of the BDS shells, the blowing ratio of the coolant streams, and the span-wise distance (i.e., the pitch) between the neighboring BDSICs, on the film cooling performance were evaluated in great detail based the quantitative PSP and PIV measurements. It was found that, in comparison with that of conventional film cooling design with coolant stream injected from circular hole (i.e., the baseline case), the film cooling effectiveness over the test plate would be enhanced significantly (i.e., 30–500%) by using the BDSIC design, especially for the test case with elongated BDS shells. As expected, the BDSIC design was found to cause a higher aerodynamic loss at relatively low blowing ratios, due to a greater friction loss associated with its complex shape. However, since the unique BDSIC design could keep the coolant stream to stay attached more firmly on the surface of the test plate (i.e., reducing the aerodynamic loss associated with the penetration of the coolant streams into the mainstream flow), the aerodynamic loss of the BDSIC design was found to become even smaller than that of the conventional film cooling design with circular hole design at relatively high blowing ratios.

Structure analysis of adiabatic film cooling effectiveness in the near field of a single inclined jet: Measurement using fast-response pressure-sensitive paint

In the present study, the film cooling effectiveness in the near field (x/D<4) of a single inclined film-cooling jet was measured using fast-response pressure sensitive paint (fast-PSP) and a low-frame-rate CCD camera. Previous experimental data demonstrated considerable variation in this region, and good agreement was established beyond it (x/D>4). The blowing ratios M=0.5 and 1.0 were used. The coolant fluid was nitrogen and the air was the mainstream fluid, and both were kept at the same temperature. A fast-PSP measurement technique was used to determine the variations of the film cooling effectiveness with time. The contours of the time-averaged film cooling effectiveness demonstrated that the coolant spread on the surface throughout the near-hole region at M=0.5, while at M=1.0, the coolant jet detached from the surface immediately behind the hole and reattached downstream around 1.5D behind the hole’s trailing edge. The spatial distribution of the film cooling effectiveness fluctuations and its cross-correlation pattern convincingly reflected the substantial influence of the energetic unsteady flow structures in the jet and cross-flow interaction. Subsequently, the Proper Orthogonal Decomposition (POD) method was used to identify the coherent parts of the film cooling effectiveness, which are regarded as the signatures of the convective large-scale vortical structures above the wall. At M=0.5, the near-hole region was subjected to the dominant influence of the counter-rotating vortex pair (CRVP), characterized by the first two POD modes, which contained up to 40% of the fluctuation energy. Two signatures of large-scale symmetric structures were identified with similar energy levels. The second two POD modes corresponding to the horseshoe vortex near the leading edge of the hole were identified and contained around 10% of the fluctuation energy. Phase-dependent variations of the large-scale convective signatures in relation to the quasi-periodic CRVP and horseshoe vortex were separately detected. At M=1.0, the signatures of the CRVP and the horseshoe vortex were also seen in the POD modes, though they were relatively difficult to distinguish.

Effect of ribbed and smooth coolant cross-flow channel on film cooling

The influence of ribbed and unribbed coolant cross-flow channel on film cooling was investigated with the coolant supply being either a plenum-coolant feed or a coolant cross-flow feed. Validation experiments were conducted with comparison to numerical results using different RANS turbulence models showed that the RNG k–ε turbulence model and the RSM model gave closer predictions to the experimental data than the other RANS models. The results indicate that at a low blowing ratio of M=0.5, the coolant supply channel structure has little effect on the film cooling. However, at a high blowing ratio of M=1.0, the adiabatic wall film cooling effectiveness is significantly lower with the plenum feed than with the cross-flow feed, especially for the cases with ribs. The film cooling with the plenum model is better at M=0.5 than at M=1.0. The film cooling with the cross-flow model is better at a blowing ratio of M=0.5 in the near hole region, while further downstream, it is better at M=1.0. The results also show that the coolant cross-flow channel with V-shaped ribs has the best adiabatic film cooling effectiveness.

Film Cooling Performance of Tripod Antivortex Injection Holes Over the Pressure and Suction Surfaces of a Nozzle Guide Vane

Film cooling performance of the antivortex (AV) hole has been well documented for a flat plate. The goal of this study is to evaluate the same over an airfoil at three different locations: leading edge suction and pressure surface and midchord suction surface. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low-speed linear cascade wind tunnel. Steady-state infrared (IR) technique was employed to measure the adiabatic film cooling effectiveness. The study has been divided into two parts: the initial part focuses on the performance of the antivortex tripod hole compared to the cylindrical (CY) hole on the leading edge. Effects of blowing ratio (BR) and density ratio (DR) on the performance of cooling holes are studied here. Results show that the tripod hole clearly provides higher film cooling effectiveness than the baseline cylindrical hole case with overall reduced coolant usage on the both pressure and suction sides of the airfoil. The second part of the study focuses on evaluating the performance on the midchord suction surface. While the hole designs studied in the first part were retained as baseline cases, two additional geometries were also tested. These include cylindrical and tripod holes with shaped (SH) exits. Film cooling effectiveness was found at four different blowing ratios. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations, thus providing cooling in the important trailing edge portion of the airfoil.

Comparison of RANS and Detached Eddy Simulation Modeling Against Measurements of Leading Edge Film Cooling on a First-Stage Vane

The performance of a showerhead arrangement of film cooling in the leading edge region of a first-stage nozzle guide vane was experimentally and numerically evaluated. A six-vane linear cascade was tested at an isentropic exit Mach number of Ma2s = 0.42, with a high inlet turbulence intensity level of 9%. The showerhead cooling scheme consists of four staggered rows of cylindrical holes evenly distributed around the stagnation line, angled at 45 deg toward the tip. The blowing ratios tested are BR = 2.0, 3.0, and 4.0. Adiabatic film cooling effectiveness distributions on the vane surface around the leading edge region were measured by means of thermochromic liquid crystals (TLC) technique. Since the experimental contours of adiabatic effectiveness showed that there is no periodicity across the span, the computational fluid dynamics (CFD) calculations were conducted by simulating the whole vane. Within the Reynolds-averaged Navier–Stokes (RANS) framework, the very widely used realizable k–ε (Rke) and the shear stress transport k–ω (SST) turbulence models were chosen for simulating the effect of the BR on the surface distribution of adiabatic effectiveness. The turbulence model which provided the most accurate steady prediction, i.e., Rke, was selected for running detached eddy simulation (DES) at the intermediate value of BR = 3. Fluctuations of the local temperature were computed by DES, due to the vortex structures within the shear layers between the main flow and the coolant jets. Moreover, mixing was enhanced both in the wall-normal and spanwise directions, compared to RANS modeling. DES roughly halved the prediction error of laterally averaged film cooling effectiveness on the suction side of the leading edge. However, neither DES nor RANS provided the expected decay of effectiveness progressing downstream along the pressure side, with 15% overestimation of ηav at s/C = 0.2.