Trailing Edge Coolant Research Articles

An Investigation of Secondary Cooling Effect on the Endwall Surface With Vane Trailing Edge Coolant Injection

Abstract Considering the adequate use of coolant, the secondary cooling effect of trailing edge coolant injection on the downstream endwall surface is investigated in this study. Distributions of adiabatic cooling effectiveness on the endwall surface are obtained through pressure-sensitive paint (PSP) technique in a linear cascade at four mass flowrate ratios (MFR = 1%, 2%, 3%, and 4%) and two density ratios (DR = 1.0 and 1.5). The configurations of trailing edge pressure-side cutback (PC) and central cutback (CC) with three compound angles (β = 15 deg, 30 deg, and 45 deg) are implemented in this study. Reynolds-averaged Navier–Stokes simulations are performed to present the flow field. Results show that the geometry of the cutback slot has a significant effect on the endwall cooling performance. As the compound angle increases, the coolant coverage is expanded in width. A higher density ratio leads to a decrease in the area of coolant coverage, while the distribution is more uniform. Generally, with a higher mass flow ratio, the coolant coverage is greatly improved in area and value, which almost covers the entire downstream endwall surface.

Turbulent Flow Heat Transfer and Thermal Stress Improvement of Gas Turbine Blade Trailing Edge Cooling with Diamond-Type TPMS Structure

Additive manufacturing allows the fabrication of relatively complex cooling structures, such as triply periodic minimal surface (TPMS), which offers high heat transfer per unit volume. This study shows the turbulent flow heat transfer and thermal stress of the Diamond-TPMS topology in the gas turbine blade trailing edge channel. The thermal-fluid-solid analysis of the Diamond-TPMS structure, made of directionally solidified GTD111, at the nearly realistic gas turbine condition is executed, and the results are compared with the conventional pin fin array at the Reynolds number of 30,000. Compared to the baseline pin fin structure, the Diamond-TPMS model distributes flow characteristics more uniformly throughout the channel. The overall heat transfer enhancement, friction factor ratio, and thermal performance are increased by 145.3%, 200.9%, and 32.5%, respectively. The temperature, displacement, and thermal stress in the Diamond-TPMS model are also distributed more evenly. The average temperature on the external surface in the Diamond-TPMS model is lower than the baseline pin fin array by 19.9%. The Diamond-TPMS network in the wedge-shaped cooling channel helps reduce the volume displacement due to the material thermal expansion by 29.3%. Moreover, the volume-averaged von Mises stress in the Diamond-TPMS structure is decreased by 28.8%.

Comparison of Aerodynamic Loss of a Gas Turbine Vane with Various Trailing Edge Cooling Schemes

In this study, a comparison of aerodynamic loss of gas turbine vane with various trailing edge configurations was experimentally performed. For the trailing edge, three cooling schemes (the cutback slot, the central injection slot, and the pressure side hole) and three levels of trailing edge diameter (7.5, 10.0, and 12.5% of the throat width, respectively) were considered. The total pressure loss was measured in a five-vane linear cascade at the low-speed mainstream condition (20 m/s) using a Kiel probe and a two-axis traversing device. The measurement region was constrained to the near mid-span region, and the measurement step size was 1 mm. The results showed that the distribution of the total pressure loss was strongly affected by cooling schemes. The vane with the pressure side hole showed a similar distribution to that of the baseline in which a cooling scheme was not applied. It was also shown that the total pressure loss increased as the trailing edge diameter increased regardless of the cooling schemes. Except for the cutback slot cases, there was an inflection point at which the tendency of the area-averaged total pressure loss to the blowing ratio was changed. Through the present study, correlation equations for each cooling scheme were derived against the trailing edge diameter and the blowing ratio.

Heat Transfer in a Rib Turbulated Pin Fin Array for Trailing Edge Cooling

Abstract Ribs were added into a pin fin array for a uniquely new cooling concept enabled through additive manufacturing. Heat transfer measurements are obtained using the thermochromic liquid crystal technique in a trapezoidal duct with pin fins and rib turbulators. Three pin blockage ratios and four rib heights at Reynolds numbers between 40,000 and 106,000 were tested. The Nusselt number augmentation is generally higher at the longer base of the trapezoidal duct. The same high heat transfer trend is seen at the columns closer to the longer base of the trapezoidal duct than on the shorter base. Through the length of the duct, the flow shifts from the nose region to the larger opening on the opposite wall. Also, it is observed that increasing the blockage ratio as well as increasing the rib height has a positive impact on heat transfer as ribs act as additional extended surfaces and alter the near-wall flow field. The heat transfer augmentation of pins and ribs is found to not be equal to the sum of both: Both heat transfer mechanisms are highly non-linear, thus cannot be superimposed. The observed heat transfer augmentation of the combined cases exceeded over the rib and pin only cases by up to 100%, but the weighted friction factor also doubled. The combination of ribs and pins is an excellent concept to achieve more uniform cooling over an array at higher levels when pressure drop is not of concern.

Experimental and Numerical Study of Turbulent Flow and Heat Transfer in a Wedge-shaped Channel with Guiding Pin Fins for Turbine Blade Trailing Edge Cooling

Gas turbine blades employ wedge-shaped channels for the trailing edge internal cooling. Detailed experimental and numerical analyses have been done on the heat transfer and turbulent flow structures in a wedge-shaped channel with arrays of streamlined pin fins for the Reynolds number range from 20,000 to 80,000. The study examines two different configurations of the streamlined guiding pin fins in the wedge-shaped channels, which show the potential of flow guiding and heat transfer improvement for the turbine blade trailing edge internal cooling. The experiments indicate that the streamlined shape of the guiding pin fins causes significantly reduced pressure losses compared to the baseline circular pin fin arrays. Depending on two different configurations the guiding pin fin arrays can achieve pressure loss reduction by up to about 37.3% and 8.7%, respectively, while obtaining appreciably higher average heat transfer on the bottom wall by up to about 8.8% and 12.0%, respectively, compared to the baseline circular pin fin arrays. Besides, the conjugate numerical simulations obtain the total Nusselt numbers of the different pin fin array configurations. Depending on two different configurations the guiding pin fin arrays increase the total Nusselt numbers by up to 11.5% and 13.8%, respectively, compared to the baseline circular pin fin arrays. Additionally, the heat transfer enhancement and pressure loss reduction present more significant promotions for the guiding pin fin arrays as the Reynolds number increases. Also, the numerical simulations reveal the detailed flow structures in the wedge-shaped channels with different pin fin configurations and demonstrate the flow guiding mechanism of the guiding pin fin arrays. Due to the flow optimizations by the guiding pin fin arrays in the wedge-shaped channels, the heat transfer and flow distribution uniformities are improved.

Analysis of ultra-low uncertainty gas turbine flow capacity measurement techniques

Accurate assessment of nozzle guide vane (NGV) capacity is essential for understanding engine performance data, and to achieve accurate turbine stage matching. In accelerated engine development programmes in particular, accurate and early assessment of NGV capacity is a significant advantage. Whilst the capabilities of computational methods have improved rapidly in recent years, the accuracy of absolute capacity prediction capability is lower than experimental techniques by some margin. Thus, experimental measurement of NGV capacity is still regarded as an essential part of many engine programmes. The semi-transient capacity measurement technique, developed and refined at the University of Oxford over the last 10 years, allows rapid and accurate measurement of engine component (typically fully cooled NGVs) capacity at engine-representative conditions of Mach and Reynolds numbers and coolant-to-mainstream pressure ratio. The technique has been demonstrated to offer considerable advantages over traditional (industrial steady-state) techniques in terms of accuracy, time and operating cost. Since the original facility was constructed, the facility has been modularised to allow for rapid interchange of test vane modules, and the instrumentation has been optimised to drive down the uncertainty in NGV capacity. In this paper, these improvements are described in detail, and a detailed uncertainty analysis is presented of the original facility, the current facility, and a proposed future facility in which the uncertainty of the measurement has been driven down to a practical limit. The bias errors of the three facilities are determined to be ±0.535%, ± 0.495% and ±0.301%, respectively (to 95% confidence). The corresponding precision uncertainties are ±0.028%, ±0.025% and ±0.025%, respectively. The extremely low precision uncertainty in particular allows very small changes in capacity to be resolved. This, combined with rapid interchangeability of test modules, allows studies of the sensitivity of capacity to secondary influences with much greater flexibility than was previously possible. Consideration is also given to the definition of vane capacity in systems with several streams at different conditions of inlet total pressure and temperature. A typical high pressure (HP) NGV has three distinct streams: a mainstream flow; coolant flow ejected from film cooling holes (distributed over the vane surface); and trailing edge coolant ejection. Whilst it is helpful for the coolant mass flow rates and inlet temperatures to be included in the definition, only a relatively small difference arises from the way in which this is achieved. Several definitions appear to share similar usefulness in terms of their robustness to changing inlet conditions of individual streams, but the favoured definition offers the possibility of isolating sensitivities to key effects such as trailing edge coolant ejection. This is achieved by explicitly expressing vane capacity as a function of two controlling pressure ratios. The overall purpose of this paper is to review and analyse in detail the current state-of-the-art in gas turbine flow capacity measurement.

Detailed Flow and Heat Transfer Analyses in a Rib-Roughened Trailing-Edge Cooling Cavity With Impingement

A rig, simulating two adjacent cooling cavities on the trailing side of an airfoil, made up of two trapezoidal channels is tested. Eleven crossover holes on the partition wall between the two channels create the jets. Two exit flow arrangements are investigated—(a) jets, after interaction with the target surface, are turned toward the target channel exit axially and (b) jets exit from a row of racetrack-shaped slots along the target channel. Flow measurements are reported for individual holes and heat transfer coefficients on the eleven target walls downstream the jets are measured using liquid crystals under steady-state conditions. Smooth as well as ribbed target surfaces with four rib angles are tested. Correlations are developed for mass flow rate through each crossover hole, varying the number of crossover holes. Heat transfer coefficient variations along the target channel are reported for a range of 5000–50,000 local jet Reynolds numbers. Major conclusions are: (1) Correlations are developed to successfully predict the air flow rate through each crossover hole for partition walls with six to eleven crossover holes, based on the pressure drop across the holes, (2) impingement heat transfer coefficient correlates well with local jet Reynolds number for both exit flow arrangements, and (3) case of target channel flow exiting from the channel end, at higher jet Reynolds numbers, produce higher heat transfer coefficients than those in the case of flow exiting through a row of slots along the target channel opposite to the crossover holes.

Turbine Blade Surface Phantom Cooling from Upstream Nozzle Trailing-Edge Ejection

This paper presents upstream nozzle trailing-edge coolant ejection on downstream uncooled blades. A pressure-sensitive paint mass transfer technique provides detailed phantom cooling effectiveness distribution on modeled land-based turbine rotor blade surfaces. Without cavity purge and tip leakage flows, a uniform blade inlet temperature is adopted in the current study. Experiments have been completed in a low-speed wind-tunnel facility with a five-blade linear cascade. The inlet Reynolds numbers based on chord length are 100,000 and 200,000. Nozzle trailing-edge coolant ejection on a rotor blade is simulated by a spoked-wheel-type rotating facility with 32 hollow rods equipped with coolant ejection from 128 holes per rod. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Nozzle coolant discharge velocity to nozzle mainstream velocity ratio varies from 0.4 to 1.4. Velocity ratios from 0.4–0.6 are closest to typical engine conditions. Coolant to mainstream mass flow rate ratio effect is from 0.67 to 2.94%. Higher phantom cooling effectiveness occurs on suction and pressure surfaces at the velocity ratio of 0.4–0.6 and over 1.0, respectively. Velocity ratio effect impacts phantom cooling effectiveness distribution more than the mass flow rate ratio effect. Most of the trailing-edge coolant migrates toward blade inner and outer spans than the blade midspan.

Experimental/Numerical Investigation on the Effects of Trailing-Edge Cooling Hole Blockage on Heat Transfer in a Trailing-Edge Cooling Channel

Hot and harsh environments, sometimes experienced by gas turbine airfoils, can create undesirable effects such as clogging of the cooling holes. Clogging of the cooling holes along the trailing edge of an airfoil on the tip side and its effects on the heat transfer coefficients in the cooling cavity around the clogged holes is the main focus of this investigation. Local and average heat transfer coefficients were measured in a test section simulating a rib-roughened trailing edge cooling cavity of a turbine airfoil. The rig was made up of two adjacent channels, each with a trapezoidal cross sectional area. The first channel supplied the cooling air to the trailing-edge channel through a row of racetrack-shaped slots on the partition wall between the two channels. Eleven crossover jets, issued from these slots entered the trailing-edge channel, impinged on eleven radial ribs and exited from a second row of race-track shaped slots on the opposite wall that simulated the cooling holes along the trailing edge of the airfoil. Tests were run for the baseline case with all exit holes open and for cases in which 2, 3, and 4 exit holes on the airfoil tip side were clogged. All tests were run for two crossover jet angles. The first set of tests were run for zero angle between the jet axis and the trailing-edge channel centerline. The jets were then tilted towards the ribs by five degrees. Results of the two set of tests for a range of jet Reynolds number from 10,000 to 35,000 were compared. The numerical models contained the entire trailing-edge and supply channels with all slots and ribs to simulate exactly the tested geometries. They were meshed with all-hexa structured mesh of high near-wall concentration. A pressure-correction based, multiblock, multigrid, unstructured/adaptive commercial software was used in this investigation. The realizable k-ε turbulence model in combination with enhanced wall treatment approach for the near wall regions were used for turbulence closure. Boundary conditions identical to those of the experiments were applied and several turbulence model results were compared. The numerical analyses also provided the share of each crossover and each exit hole from the total flow for different geometries. The major conclusions of this study were: (a) clogging of the exit holes near the airfoil tip alters the distribution of the coolant mass flow rate through the crossover holes and changes the flow structure. Depending on the number of clogged exit holes (from 3 to 6, out of 12), the tip-end crossover hole experienced from 35% to 49% reductions in its mass flow rate while the root-end crossover hole, under the same conditions, experienced an increase of the same magnitude in its mass flow rate. (b) Up to 64% reduction in heat transfer coefficients on the tip-end surface areas around the clogged holes were observed which might have devastating effects on the airfoil life. At the same time, a gain in heat transfer coefficient of up 40% was observed around the root-end due to increased crossover flows. (c) Numerical heat transfer results with the use of the realizable k-ε turbulence model in combination with enhanced wall treatment approach for the near wall regions were generally in a reasonable agreement with the test results. The overall difference between the CFD and test results was about 10%.

Heat Transfer and Friction Studies in a Tilted and Rib-Roughened Trailing-Edge Cooling Cavity with and without the Trailing-Edge Cooling Holes

Local and average heat transfer coefficients and friction factors were measured in a test section simulating the trailing-edge cooling cavity of a turbine airfoil. The test rig with a trapezoidal cross-sectional area was rib-roughened on two opposite sides of the trapezoid (airfoil pressure and suction sides) with tapered ribs to conform to the cooling cavity shape and had a 22-degree tilt in the flow direction upstream of the ribs that affected the heat transfer coefficients on the two rib-roughened surfaces. The radial cooling flow traveled from the airfoil root to the tip while exiting through 22 cooling holes along the airfoil trailing-edge. Two rib geometries, with and without the presence of the trailing-edge cooling holes, were examined. The numerical model contained the entire trailing-edge channel, ribs, and trailing-edge cooling holes to simulate exactly the tested geometry. A pressure-correction based, multiblock, multigrid, unstructured/adaptive commercial software was used in this investigation. Realizablek-εturbulence model in conjunction with enhanced wall treatment approach for the near wall regions was used for turbulence closure. The applied thermal boundary conditions to the CFD models matched the test boundary conditions. Comparisons are made between the experimental and numerical results.

Nusselt Correlations in a Trailing Edge Cooling System with Long Pedestals and Ribs

Detailed TLC measurements have been carried out to estimate the heat transfer coefficient in stationary conditions under high Reynolds numbers in a trailing edge cooling system of a high pressure gas turbine blade. The investigated geometry consists of a 30:1 PMMA scaled model reproducing the typical wedge-shaped discharge duct of a trailing edge cooling system with one row of 7 enlarged pedestals. The section of the channel upstream the pedestal region is used to guide the airflow from the radial hub inlet to the tangential trailing edge outlet; in this section three different surfaces have been studied: one is smooth and the other two are ribbed with +60° and -60° angled ribs respect to the radial direction. This work focuses on the pedestal outlet section, by giving a correlation of the variation of the Nusselt number as a function of the Reynolds number, from 10000 to 40000, in the different inter-pedestal regions along the radial direction. These correlations give the regionally averaged heat transfer coefficients, from the hub to the tip of the model for both smooth and ribbed cases. The interest of these results is the use for the design of the trailing edge blade cooling systems, under the investigated Reynolds number range, that it is the typical case of the industrial application.

An Experimental Study on Aerodynamic Performance of Turbine Nozzle Guide Vanes With Trailing-Edge Span-Wise Ejection

The present experimental study is to examine the influence of trailing-edge coolant ejection with the span-wise inclination on the aerodynamic loss of turbine nozzle guide vanes. This study uses a cascade of five vanes located in the test section of a low-speed wind tunnel. The vanes have the profile of high-pressure nozzle guide vanes, and the central vane is equipped with the internal cooling and the trailing-edge coolant ejection. The coolant is ejected through trailing-edge slots that are inclined in the span-wise direction at angles varying from 0 deg to 45 deg in 15 deg increments. The results indicate an optimum ejection rate, at which the aerodynamic loss is minimum. There is a little variation in loss as the span-wise inclination is varied when the ratio of coolant to mainstream gas mass flow rate is less than 1.5%. For higher coolant flow rates, however, the loss increases with increases in the span-wise ejection angle.

Influence of Unsteady Wake With Trailing Edge Coolant Ejection on Turbine Blade Film Cooling

Detailed film cooling effectiveness distributions along a modeled turbine rotor blade under the combined effects of an upstream trailing edge unsteady wake with coolant ejection are presented using the pressure sensitive paint (PSP) mass transfer analogy method. The experiment is conducted in a low speed wind tunnel facility with a five blade linear cascade. The exit Reynolds number based on the axial chord is 370,000. Unsteady wakes and trailing edge coolant jets are produced by a spoked wheel-type wake generator with hollow rods equipped with several coolant ejections from holes. The coolant-to-mainstream density ratios for both the blade and trailing edge coolant ejection range from 1.5 to 2.0 for simulating realistic engine conditions. Blade blowing ratio studies are 0.5 and 1.0 on the suction surface and 1.0 and 2.0 on the pressure surface. The trailing edge jet blowing ratio and Strouhal numbers are 1.0 and 0.12, respectively. The results show that the unsteady wake reduces the overall film cooling effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances the overall effectiveness. The results also show that the overall filming cooling effectiveness increases by using heavier coolant for trailing edge ejection and for blade surface film cooling.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part II: Independent Influences of Vane Trailing Edge and Purge Cooling

The independent influences of vane trailing edge and purge cooling are studied in detail for a one-and-one-half stage transonic high-pressure turbine operating at design-corrected conditions. This paper builds on the conclusions of Part I, which investigated the combined influence of all cooling circuits. Heat-flux measurements for the airfoil, platform, tip, and root of the turbine blade, as well as the shroud and the vane side of the purge cavity, are used to track the influence of cooling flow. By independently varying the coolant flow rate through the vane trailing edge or purge circuit, the region of influence of each circuit can be isolated. Vane trailing edge cooling is found to create the largest reductions in blade heat transfer. However, much of the coolant accumulates on the blade suction surface and little influence is observed for the pressure surface. In contrast, the purge cooling is able to cause small reductions in heat transfer on both the suction and pressure surfaces of the airfoil. Its region of influence is limited to near the hub, but given that the purge coolant mass flow rate is 1/8 that of the vane trailing edge, it is impressive that any impact is observed at all. The cooling contributions of these two circuits account for nearly all of the cooling reductions observed for all three circuits in Part I, indicating that the vane inner cooling circuit that feeds most of the vane film-cooling holes has little impact on the downstream blade heat transfer. Time-accurate pressure measurements provide further insight into the complex interactions in the purge region that govern purge coolant injection. While the pressures supplying the purge coolant and the overall coolant flow rate remain fairly constant, the interactions of the vane pressure field and the rotor pressure field create moving regions of high pressure and low pressure at the exit of the cavity. This results in pulsing regions of injection and ingestion.

Numerical and Experimental Investigation of Turning Flow Effects on Innovative Pin Fin Arrangements for Trailing Edge Cooling Configurations

The effect of the array configuration of circular pin fins is investigated from a numerical and experimental point of view reproducing a typical cooling scheme of a real high pressure aero-engine blade. The airstream enters the domain of interest radially from the hub inlet and exits axially from the trailing edge (TE) outlet section. More than 100 turbulators are inserted in the wedge-shaped TE duct to enhance the heat transfer: A reference array implementing seven rows of staggered pins is compared with an innovative pentagonal arrangement. Investigations were made considering real engine flow conditions: Both numerical calculations and experimental measurements were performed fixing Re=18,000 and Ma=0.3 in the TE throat section. The effect of the tip mass flow rate was also taken into account, investigating 0% and 25% of the TE mass flow rate. The experimental activity was aimed at obtaining detailed heat transfer coefficient maps over the internal pressure side (PS) surface by means of the transient technique with thermochromic liquid crystals. Particle image velocimetry measurements were performed and surface flow visualizations were made by means of the oil and dye technique on the PS surface. Steady-state Reynolds averaged Navier–Stokes simulations were performed with two different computational fluid dynamics (CFD) codes: the commercial software Ansys CFX®11.0 and an in-house solver based on the opensource toolbox OpenFOAM®, to compare the performance and predictive capabilities. Turbulence was modeled by means of the k−ω shear stress transport (SST) model with a hybrid near-wall treatment allowing strong clustering of the wall of interest as well as quite coarse refinement on the other viscous surfaces.

Experimental Investigation of Innovative Internal Trailing Edge Cooling Configurations with Pentagonal Arrangement and Elliptic Pin Fin

This paper describes a heat transfer experimental study of four different internal trailing edge cooling configurations based on pin fin schemes. The aim of the study is the comparison between innovative configurations and standard ones. So, a circular pin fin configuration with an innovative pentagonal scheme is compared to a standard staggered scheme, while two elliptic pin fin configurations are compared to each other turning the ellipse from the streamwise to the spanwise direction. For each configuration, heat transfer and pressure loss measurements were made keeping the Mach number fixed at 0.3 and varying the Reynolds number from 9000 to 27000. In order to investigate the overall behavior of both endwall and pedestals, heat transfer measurements are performed using a combined transient technique. Over the endwall surface, the classic transient technique with thermochromic liquid crystals allows the measurement of a detailed heat transfer coefficient (HTC) map. Pin fins are made of high thermal conductivity material, and an inverse data reduction method based on a finite element code allows to evaluate the mean HTC of each pin fin. Results show that the pentagonal arrangement generates a nonuniform HTC distribution over the endwall surface, while, in terms of average values, it is equivalent to the staggered configuration. On the contrary, the HTC map of the two elliptic configurations is similar, but the spanwise arrangement generates higher heat transfer coefficients and pressure losses.

Aerodynamic Loss Characteristics of a Turbine Blade With Trailing Edge Coolant Ejection: Part 1—Effect of Cut-Back Length, Spanwise Rib Spacing, Free-Stream Reynolds Number, and Chordwise Rib Length on Discharge Coefficients

The internal fluid mechanics losses generated between the blade plenum chamber and a reference point located just downstream of the trailing edge are investigated for a turbine blade trailing edge cooling system. The discharge coefficient Cd is presented as a function of the free-stream Reynolds number, cut-back length, spanwise rib spacing, and chordwise rib length. The results are presented in a wide range of coolant to free-stream mass flow rate ratios. The losses from the cooling system show strong free-stream Reynolds number dependency, especially at low ejection rates, when they are correlated against the coolant to free-stream pressure ratio. However, when Cd is correlated against a coolant to free-stream mass flow rate ratio, the Reynolds number dependency is eliminated. The current data clearly show that internal viscous losses due to varying rib lengths do not differ significantly. The interaction of the external wall jet in the cutback region with the free-stream fluid is also a strong contributor to the losses. Since the discharge coefficients do not have Reynolds number dependency at high ejection rates, Cd experiments can be performed at a low free-stream Reynolds number. Running a discharge coefficient experiment at low Reynolds number (or even in still air) will sufficiently define the high blowing rate portion of the curve. This approach is extremely time efficient and economical in finding the worst possible Cd value for a given trailing edge coolant system.

Investigation of the Unsteady Rotor Aerodynamics in a Transonic Turbine Stage

The paper focuses on the unsteady pressure field measured around the rotor midspan profile of the VKI Brite transonic turbine stage. The understanding of the complex unsteady flow field is supported by a quasi-three-dimensional unsteady Navier–Stokes computation using a k-ω turbulence model and a modified version of the Abu-Ghannam and Shaw correlation for the onset of transition. The agreement between computational and experimental results is satisfactory. They both reveal the dominance of the vane shock in the interaction. For this reason, it is difficult to identify the influence of vane-wake ingestion in the rotor passage from the experimental data. However, the computations allow us to draw some useful conclusions in this respect. The effect of the variation of the rotational speed, the stator–rotor spacing, and the stator trailing edge coolant flow ejection is investigated and the unsteady blade force pattern is analyzed.