Inlet Turbulence Intensity Level Research Articles

Experimental analysis of the impact of an acoustic excitation on a separated boundary layer behaviour

The paper examines the effect of acoustic excitation at broadband frequency range (100 - 650 Hz) on separated boundary layer developing on a flat plate subjected to an adverse pressure gradient. The experiment was conducted with a low inlet turbulence intensity level (Tu < 1%) in order to provide a cleaner environment that magnifies the effects of the excitation frequency. It was shown that acoustic excitation at sound pressure levels: SPL=125 and 135 dB, can lead to a more rapid increase in flow instability followed by an earlier laminar-turbulent (l-t) transition. The boundary layer reattachment point was shifted upstream and the size of the separation bubble was reduced in the flow with the acoustic excitation active.

Experimental and Numerical Assessment on S-shaped Diffuser performance with different Turbulence Intensity

The turbulence structure in air intake S-shaped diffuser is proven to be influencing parameter on the diffuser performance. In the present research work, experimental and numerical investigations have been undertaken to explore the effect of inlet turbulence intensity level on the performance of air intake S-shaped diffuser. Detailed measurements including pressure and velocity at the inlet and outlet planes and static pressure on the top and bottom walls were taken at three pre-selected Reynolds numbers of 4.8×104, 6.4×104 and 7.5×104. The predicted corresponding turbulence intensities at the experimented three Re, were 4.16%, 3.15% and 2.8%. ANSYS-FLUENT 15 software with Standard k-ε turbulence model has been used for numerical simulations. Numerical results of static pressure recovery, total pressure loss coefficient and wall static pressure recovery have been compared with the experimental results. The experimental results indicate that increasing Reynolds number at the inlet of S-shaped diffuser resulting in slight decrease in measured turbulence intensity. Also, as Re increased from 4.8x104 to 7.5x104, the static pressure recovery increased by 9% and the total pressure loss coefficient was reduced by 4.9%.

A Parametric Investigation of Vane Showerhead Film Cooling by Pressure-Sensitive Paint Technique

Abstract In this study, a parametric analysis of the thermal performance of a nozzle vane cascade with a showerhead cooling system made of four rows of cylindrical holes was carried out by using the pressure-sensitive paint (PSP) technique. Coolant-to-mainstream blowing ratio (BR), density ratio (DR), main flow isentropic exit Mach number (Ma2is), and turbulence intensity level (Tu1) were the considered parameters. The cascade was tested in an atmospheric wind tunnel at Ma2is values ranging from 0.2 to 0.6, with an inlet turbulence intensity level of 1.6% and 9%, at variable injection conditions of BR = 2.0, 3.0, 4.0. Moreover, the influence of the DR on the leading-edge film-cooling performance was investigated: the testing was carried out at DR = 1.0, using nitrogen as foreign gas, and DR = 1.5, with carbon dioxide serving as a coolant. In the near-hole region, higher BR and Ma2is resulted in higher effectiveness, while higher mainstream turbulence intensity reduced the thermal coverage in between the rows of holes, whatever the BR is. Further downstream along the vane pressure side, the effectiveness was negatively affected by rising the BR but positively influenced by lowering the mainstream turbulence intensity. Moreover, a decrease in the DR caused a reduction in the film-cooling performance, whose extent depends on the injection condition.

Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through four rows of cylindrical holes. 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, varying the inlet flow angle in the range ±20 deg. The aero-thermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20 deg.

The Effect of Hot Streaks on a High Pressure Turbine Vane Cascade with Showerhead Film Cooling

Hot streak migration in a linear vane cascade with showerhead film cooling was experimentally and numerically investigated at isentropic exit Mach number of Ma2is = 0.40, with an inlet turbulence intensity level of Tu1 = 9%. Two tangential positions of the hot streak center were taken into account: 0% of pitch (hot streak is aligned with the vane leading edge) and 45% of pitch. After demonstrating that computations correctly predict hot streak attenuation through the vane passage with no showerhead blowing, the numerical method was used to investigate hot streak interaction with showerhead film cooling, at blowing ratio of BR = 3.0, corresponding to a coolant-to-mainstream mass flow ratio of MFR = 1.15%. The effects of mixing and coolant interaction on the hot streak reduction were interpreted under the light of the superposition principle, whose accuracy was within 12% on the leading edge region, in the central section of the vane span.

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.

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.

Experimental Investigation of the Unsteady Flow Behavior on a Film Cooling Flat Plate

An experimental study was conducted to investigate the unsteady behavior of film cooling for cylindrical holes on a flat plate wind tunnel. Tests have been carried out at low speed and low inlet turbulence intensity level, with blowing ratios varied in the range 0.5–1.5. Aerodynamic investigations have been performed through the measurement of discharge coefficients in order to characterize the blowing conditions. A Laser Doppler Velocimetry (LDV) was then used to study the boundary layer behavior just downstream of the holes for variable injection condition. A high resolution PIV system was finally used for flow visualization and flow field measurement to investigate the unsteady mixing process taking place between coolant and main flow. For low blowing ratios, the jet stays close to the surface; traces of coherent vortical structures are detected only far from injection, which appear as clockwise vortices. Increasing the blowing ratio leads to the appearance of counter-clockwise vortical structures due to the breakdown of the Kelvin Helmholtz instability of the shear layers.

Numerical predictions of detailed flow structural characteristics in a channel with angled rib turbulators

Turbulent air flows within a channel with angled rib turbulators (45 degrees) on one wall are numerically predicted using the numerical code ANSYS CFX with a Shear-stress transport (SST) κ-ω turbulence model, and a hexahedral grid with 7115346 cells and no wall function. Three-dimensional turbulent transport, and detailed flow structural characteristics are considered to provide new insight into the mechanisms which result in surface heat transfer augmentations. Time-averaged turbulent flow characteristics and surface Nusselt number distributions are presented for an inlet turbulence intensity level of 1.0 percent, and for Reynolds numbers based upon channel height of 18300 and 48000. Overall, the numerically-predicted results show that large-scale, secondary flow induces a collection of small-scale vortical flows in the channel, as a result of local interactions with individual rib turbulators. These are often associated with different sized and highly skewed vortex pairs, which also induce secondary advection and increased turbulent mixing near ribbed channel surfaces. Within the flow separation regions, just downstream of each rib, surface Nusselt number ratios which are locally lower than for other surface locations, and local secondary flows are generally and partially characterized as upwash flows, with large positive magnitudes of spanwise vorticity, large static pressure deficits, and large static pressure augmentation regions. As the shear layers (initially located above the recirculation zones) impinge onto the test surface, increased mixing develops, as well as local thinning of the reattaching boundary layers, which lead to local Nusselt numbers which are generally higher than for other locations along the test surface.

Effects of Inlet Turbulence and End-Wall Boundary Layer on Aerothermal Performance of a Transonic Turbine Blade Tip

Most of the previous researches of inlet turbulence effects on blade tip have been carried out for low speed situations. Recent work has indicated that for a transonic turbine tip, turbulent diffusion tends to have a distinctively different impact on tip heat transfer than for its subsonic counterpart. It is hence of interest to examine how inlet turbulence flow conditioning would affect heat transfer characteristics for a transonic tip. The present work is aimed to identify and understand the effects of both inlet freestream turbulence and end wall boundary layer on a transonic turbine blade tip aerothermal performance. Spatially-resolved heat transfer data are obtained at aerodynamic conditions representative of a high-pressure turbine, using the transient infrared thermography technique with the Oxford High-Speed Linear Cascade research facility. With and without turbulence grids, the turbulence levels achieved are 7%–9% and 1%, respectively. On the blade tip surface, no apparent change in heat transfer was observed with high and low inlet turbulence intensity levels investigated. On the blade suction surface, however, substantially different local heat transfer distributions for the suction side near tip surface have been observed, indicating a strong local dependence of the local vortical flow structure on the freestream turbulence. These experimentally observed trends have also been confirmed by CFD examinations using the Rolls-Royce HYDRA. A further CFD analysis suggests that the level of inflow turbulence alters the balance between the passage vortex associated secondary flow and the over tip leakage (OTL) flow. Consequently, an enhanced inertia of near wall fluid at a higher inflow turbulence weakens the cross-passage flow. As such, the weaker passage vortex leads the tip leakage vortex to move further into the mid passage, with the less spanwise coverage on the suction surface, as consistently indicated by the heat transfer signature. Different inlet end wall boundary layer profiles are employed in the computational study with HYDRA. All CFD results indicate the inlet boundary layer thickness has little impact on the heat transfer over the tip surface as well as the pressure side near-tip surface. However, noticeable changes in heat transfer are observed for the suction side near-tip surface. Similar to the inlet turbulence effect, such changes can be attributed to the interaction between the passage vortex and the OTL flow.

Experimental investigation of the effects of blowing conditions and Mach number on the unsteady behavior of coolant ejection through a trailing edge cutback

The present paper shows the results of an experimental investigation into the unsteadiness of coolant ejection at the trailing edge of a highly loaded nozzle vane cascade. The trailing edge cooling scheme features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is also ejected through two rows of cylindrical holes placed upstream of the cutback. Tests were performed with a low inlet turbulence intensity level (Tu1=1.6%), changing the cascade operating conditions from low speed (M2is=0.2) up to high subsonic regime (M2is=0.6), and with coolant to main stream mass flow ratio varied within the 0.5–2.0% range. Particle Image Velocimetry (PIV) and flow visualizations were used to investigate the unsteady mixing process taking place between coolant and main flow downstream of the cutback, up to the trailing edge. For all the tested conditions, the results show the presence of large coherent structures, which presence is still evident up to the trailing edge. Their shape and direction of rotation change with injection conditions, as a function of coolant to mainstream velocity ratio, strongly influencing the thermal protection capability of the injected coolant flow. The Mach number increase is only responsible for a positioning of such vortical structures closer to the wall, while the Strouhal number almost remains unchanged.

Effects of Trenched Holes on Film Cooling of a Contoured Endwall Nozzle Vane

The present paper investigates the effects of the application of trenched holes in the front part of a contoured film cooled endwall. Two trench configurations were tested, changing the trench depth. Tests have been carried out at low speed (M2is=0.2) and low inlet turbulence intensity level, with coolant mass flow rate ratio varied within the 0.5–2.5% range. Pressure probe traverses were performed downstream of the vane trailing edge to show the secondary flow field modifications and to evaluate trench additional losses. Endwall distributions of film cooling effectiveness have been obtained by the thermochromic liquid crystal (TLC) technique. For each injection condition, energy loss coefficient and film cooling effectiveness distributions were analyzed and compared with the ones obtained from rows of cylindrical holes. Laterally and area averaged effectiveness as well as pitch and mass averaged kinetic energy loss coefficients were computed to enlighten any change induced by the introduction of trenched holes. A uniform and high thermal coverage was obtained in the region just downstream of the trench, but it quickly decayed because of enforced mixing of coolant with main flow. Compared with the cylindrical hole configuration, trenches are able to provide a higher global cooling effectiveness, but a larger amount of coolant injection is required. The introduction of both trenches is responsible for a secondary thermodynamic loss increase of about 0.7% at low coolant injection rates. Increasing the blowing rates, the additional loss vanishes.

Film cooling of a contoured endwall nozzle vane through fan-shaped holes

The present paper shows the results of an experimental investigation into a nozzle vane cascade with endwall contouring and film cooling. In a film cooled endwall the front area, especially close to the leading edge, is one of the most critical regions to be cooled; in case of a contoured endwall the thermal protection becomes even more difficult because of the endwall curvature effect. Two film cooling hole geometries with four rows of holes were tested: cylindrical holes and conical expanded holes. Tests were performed at low speed ( M 2 is = 0.2) and low inlet turbulence intensity level, with coolant to main stream mass flow ratio varied within the 0.5–2.5% range. Aerodynamic and thermal performances were evaluated in every injection condition. Compared to cylindrical holes, results showed that shaped film cooling holes generate slightly higher secondary losses and provide a better thermal coverage, at the expense of a higher coolant flow injection.

Endwall Film Cooling Effects on Secondary Flows in a Contoured Endwall Nozzle Vane

The present paper investigates the effects of endwall injection of cooling flow on the aerodynamic performance of a nozzle vane cascade with endwall contouring. Tests have been performed on a seven vane cascade with a geometry typical of a real gas turbine nozzle vane. The cooling scheme consists of four rows of cylindrical holes. Tests have been carried out at low speed (Ma2is=0.2) with a low inlet turbulence intensity level (1.0%) and with a coolant to mainstream mass flow ratio varied in the range from 0% (solid endwall) to 2.5%. Energy loss coefficient, secondary vorticity, and outlet angle distributions were computed from five-hole probe measured data. Contoured endwall results, with and without film cooling, were compared with planar endwall data. Endwall contouring was responsible for a significant overall loss decrease, as a result of the reduction in both profile and planar side secondary flows losses; a loss increase on the contoured side was instead observed. Like as for the planar endwall, even for the contoured endwall, coolant injection modifies secondary flows, reducing their intensity, but the relevance of the changes is reduced. Nevertheless, for all the tested injection conditions, secondary losses on the contoured side are always higher than in the planar case, while contoured cascade overall losses are lower. A unique minimum overall loss injection condition was found for both tested geometries, which corresponds to an injected mass flow rate of about 1.0%.

Breakdown of Laminar Pipe Flow into Transitional Intermittency and Subsequent Attainment of Fully Developed Intermittent or Turbulent Flow

The breakdown of laminar pipe flow into transitional intermittency is predicted numerically here for the first time. Subsequent to transitional intermittency, a fully developed regime is achieved wherein the flow may be either intermittent or fully turbulent. Fully developed friction factors are predicted as a function of the Reynolds number throughout the intermittent regime. These predictions successfully bridge the gap between well-established laminar friction factors and turbulent friction factors. Definitive numerical information is provided about the locations in the pipe at which both laminar breakdown and fully developed attainment occur. These locations are a function of the Reynolds number. The streamwise changes in the velocity profiles reflect the complex evolution of the flow as it passes through the successive regimes. For design purposes, information is provided for the pressure drop that characterizes the evolving flow. The numerical results correspond to an inlet turbulence intensity level of 5%.

Mach Number and Freestream Turbulence Effects on Turbine Vane Aerodynamic Losses

The aerodynamic performance of a smooth. cambered turbine vane (which replicates one used in an operating gas turbine engine) is investigated in this paper. Three different Mach-number distributions are employed, which result in one transonic flow and two subsonic flows. All of these distributions match flow conditions in different industrial applications. A fine mesh grid and cross bars are used to augment the magnitudes of longitudinal turbulence intensity at the inlet of the test section. Wake-profile data are presented for two different locations downstream of the vane trailing edge (one axial chord length and 0.25 axial chord length). The contributions of varying Mach number and varying freestream turbulence intensity to aerodynamic losses, normalized kinetic energy profiles, normalized Mach-number profiles, integrated aerodynamies losses, and area-averaged loss coefficients are quantified. Results show that wake profiles are more sensitive to turbulence intensity variations at lower subsonic flow conditions than when transonic flow is present. Wake profiles are also broadened either as the exit Mach number increases or as the freestream turbulence intensity level increases. Higher losses in the freestream flow are present as the inlet turbulence intensity level increases. Also described are effects of increased turbulent diffusion, streamwise development, and profile asymmetry. Corresponding integrated aerodynamic losses and area-averaged loss coefficient Y A magnitudes increase with increasing Mach number or with increasing turbulence intensity level. Results additionally show larger loss magnitudes with flow turning and cambered airfoils, relative to symmetric airfoils, when compared at the same exit Mach number.

Mach Number/Surface Roughness Effects on Symmetric Transonic Turbine Airfoil Aerodynamic Losses

The effects of surface roughness and Mach number on the aerodynamic performance of turbine airfoils are investigated in compressible, high-speed flows with exit freestream Mach numbers of 0.6, 0.8, and 0.9 and three different inlet turbulence intensity levels of 0.9, 5.5, and 16.2%. Three symmetric airfoils, each with the same shape and exterior dimensions, are employed with different rough surfaces. The nonuniform, irregular, three-dimensional roughness is characterized using the equivalent sand grain roughness size. Increased surface roughness size and increasing freestream Mach number produce larger magnitudes of nondimensional total pressure loss coefficients. Magnitudes of integrated aerodynamic losses change by a much larger amount as either the freestream Mach number or turbulence intensity are altered, when the airfoil is roughened (compared to smooth airfoil results). This is partially a result of the thicker boundary layers, which develop over the roughened surfaces giving greater blockage and less expansion of the flow through the airfoil passage. As a result, total pressure losses indicate that oblique shock waves are present at the trailing edge of the airfoil (with transonic passage flow) when the airfoil is smooth, which are not present when the airfoil surfaces become roughened.

Effects of Surface Roughness and Turbulence Intensity on the Aerodynamic Losses Produced by the Suction Surface of a Simulated Turbine Airfoil

The effects of surface roughness on the aerodynamic performance of turbine airfoils are investigated with different inlet turbulence intensity levels of 0.9%, 5.5% and 16.2%. Three symmetric airfoils, each with the same shape and exterior dimensions, are employed with different rough surfaces. The nonuniform, irregular, 3-D roughness is characterized using the equivalent sand grain roughness size. Mach numbers along the airfoil range from 0.4 to 0.7. Chord Reynolds numbers based on inlet and exit flow conditions are 0.54×106 and 1.02×106, respectively. The contributions of varying surface roughness and turbulence intensity level to aerodynamic losses, Mach number profiles, normalized kinetic energy profiles, and Integrated Aerodynamics Losses (IAL) are quantified. Results show that effects of changing the surface roughness condition on IAL values are substantial, whereas the effects of different inlet turbulence intensity levels are generally relatively small.

Nusselt Numbers and Flow Structure on and Above a Shallow Dimpled Surface Within a Channel Including Effects of Inlet Turbulence Intensity Level

Abstract Experimental results from a channel with shallow dimples placed on one wall are given for Reynolds numbers based on channel height from 3,700 to 20,000, levels of longitudinal turbulence intensity from 3% to 11% (at the entrance of the channel test section), and a ratio of air inlet stagnation temperature to surface temperature of approximately 0.94. The ratio of dimple depth to dimple print diameter δ∕D is 0.1, and the ratio of channel height to dimple print diameter H∕D is 1.00. The data presented include friction factors, local Nusselt numbers, spatially averaged Nusselt numbers, a number of time-averaged flow structural characteristics, flow visualization results, and spectra of longitudinal velocity fluctuations which, at a Reynolds number of 20,000, show a primary vortex shedding frequency of 8.0Hz and a dimple edge vortex pair oscillation frequency of approximately 6.5Hz. The local flow structure shows some qualitative similarity to characteristics measured with deeper dimples (δ∕D of 0.2 and 0.3), with smaller quantitative changes from the dimples as δ∕D decreases. A similar conclusion is reached regarding qualitative and quantitative variations of local Nusselt number ratio data, which show that the highest local values are present within the downstream portions of dimples, as well as near dimple spanwise and downstream edges. Local and spatially averaged Nusselt number ratios sometimes change by small amounts as the channel inlet turbulence intensity level is altered, whereas friction factor ratios increase somewhat at the channel inlet turbulence intensity level increases. These changes to local Nusselt number data (with changing turbulence intensity level) are present at the same locations where the vortex pairs appear to originate, where they have the greatest influences on local flow and heat transfer behavior.

Heat Transfer and Flow on the First-Stage Blade Tip of a Power Generation Gas Turbine: Part 1—Experimental Results

A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first-stage blade tip surface for a geometry typical of large power generation turbines (&gt;100 MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment that develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high-pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57×106, and total turning of about 110 deg. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5 or 9 percent. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp edge and rounded edge tip geometries at each of the inlet turbulence intensity levels. [S0889-504X(00)01902-4]