The strive to achieve increasingly higher efficiencies in gas turbine power generation has led to a continued rise in the turbine inlet temperature. As a result, novel cooling approaches for gas turbine blades are necessary to maintain them within the material’s thermal mechanical performance envelope. Various internal and external cooling technologies are used in different parts of the blade airfoil to provide the desired levels of cooling. Among the different regions of the blade profile, the trailing edge (TE) presents additional cooling challenges due to the thin cross section and high thermal loads. In this study, a new wavy geometry for the TE has been proposed and analyzed using steady state numerical simulations. The wavy TE structure resembled a sinusoidal wave running along the span of the blade. The troughs on both pressure side and suction side contained the coolant exit slots. As a result, consecutive coolant exit slots provided an alternating discharge between the suction side and the pressure side of the blade. Steady state conjugate heat transfer simulations were carried out using CFX solver for four coolant to mainstream mass flow ratios of 0.45%, 1%, 1.5% and 3%. The temperature distribution and film cooling effectiveness in the TE region were compared to two conventional geometries, pressure side cutback and centerline ejection which are widely used in vanes and blades for both land-based and aviation gas turbine engines. Unstructured mesh was generated for both fluid and solid domains and interfaces were defined between the two domains. For turbulence closer, SST-kω model was used. The wall y+ was maintained < 1 by using inflation layers at all the solid fluid interfaces. The numerical results depicted that the alternating discharge from the wavy TE was able to form protective film coverage on both the pressure and suction side of the blade. As a result, significant reduction in the TE metal was observed which was up to 14% lower in volume averaged temperature in the TE region when compared to the two baseline conventional configurations.

Turbine vane endwalls are highly susceptible to intensive heat load due to their large exposed area and complex flow field especially for the first-stage of the vane. Therefore, a suitable film cooling design that properly distributes the given amount of coolant is critical to keep the vane endwall from failure at the same time to maintain a good balance between manufacturing cost, performance, and durability. This work is focused on film cooling effectiveness evaluation on full-scale heavy duty turbine vane endwall and the performance comparison with different film cooling pattern designs in the literature. The area of interest of this study is on the inner endwall (hub) of turbine vane. Tests were performed in a three-vane annular sector cascade under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09 and the freestream turbulence intensity is 12%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are investigated. The conduction-error free Pressure sensitive paint (PSP) technique is utilized to evaluate the local flow behavior as well as the film cooling performance. The presented results are expected to provide the gas turbine engine designer a direct comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.

This report describes the detailed experimental study to characterize the local heat transfer coefficient distribution over the internal cooling passages of a simplified generic airfoil. The airfoil is manufactured through additive manufacturing based on actual geometry and dimensions (1X scale model) of row one airfoil, applicable in large gas turbine system. At the mainbody section, the serpentine channel consists of three passages without any surface features or vortex generators. Both the leading edge and trailing edge sections are subjected to direct impingement. The trailing edge section is divided into three chambers, separated by two rows of blockages. This study employs the well-documented transient liquid crystal technique, where the local heat transfer coefficient on both pressure and suction sides is deduced. The experiments were performed at varying Reynolds number, ranging from approximately 31,000–63,000. The heat transfer distribution on the pressure side and suction side is largely comparable in the first and third pass, except for the second pass. Highest heat transfer occurs at the trailing edge region, which is ultimately dominated by impingement due to the presence of three rows of blockages. A cursory numerical calculation is performed using commercially available software, ANSYS CFX to obtain detailed flow field distribution within the airfoil, which explains the heat transfer behavior at each passage. The flow parameter results revealed that the pressure ratio is strongly proportional with increasing Reynolds number.

Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane because the complex flow characteristics and intensive heat load that comes from the exit of the combustion chamber. A better understanding is needed on how the coolant flow interacts with the mainstream and the resulting cooling effect in the real engine especially for the first-stage vane. An authentic flow channel and condition should be achieved. In this study, three full-scale turbine vanes are used to construct an annular-sector cascade. The film-cooling design is attained through numerous layback fan-shaped and cylindrical holes dispersed on the vane and both end-walls. With the three-dimensional vane geometry and corresponding wind tunnel design, the true flow field can thus be simulated as in the engine. This study targets the film-cooling effectiveness on the inner end-wall (hub) of turbine vane. Tests are performed under the mainstream Reynolds number 3.5 × 10 5 ; the related inlet Mach number is 0.09 and the free stream turbulence intensity is 8%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are examined. Pressure-sensitive paint (PSP) technique is utilized to capture the detail contour of film-cooling effectiveness on the inner end-wall and demonstrate the coolant trace. The presented results serve a comparison basis for other sets of vanes with different cooling designs. The results are expected to strengthen the promise of PSP technique on evaluating the film-cooling performance of the engine geometries.

Jet mixing inside a slot has the potential to be applied for cooling in various components in gas turbines. Based on the premise that more uniform flow at the exit of the slot would achieve more uniform cooling downstream of the slot, an experimental program was conducted to focus on investigating the flow mixing behavior inside the slots. Various parameters including orientation angle, inclination angle, slot width, effect of primary flow and slot depth were systematically examined. An array of seven jets with a jet-to-jet spacing of five diameters and velocity ratio of unity was used in the experiment. Experiments were performed by changing the orientation angles of the jets with two inclination angles (90° and 30°). The results indicated that the flow distribution at the slot exit becomes more uniform as the orientation angle was increased from 0° to 60°. Wider slot width brought about high non-uniformity and was clearly undesirable. The optimum slot depth was found to range from 2 to 2.8 times the jet diameter for slots perpendicular to the primary flow (i.e. with 90° inclination angle). The flow field was the most uniform where the shear layers of the two adjacent jets met. Regions of low velocity and high pressure caused non-uniform velocity distribution downstream where the jets met, so deeper slots do not necessarily render better uniformity. When the orientation angle increased from 0 to 30° and 60°, the jets bent toward the wall in the inclined direction caused by the Coanda effect. The presence of the primary flow forced the jet to spread faster in the lateral direction and caused some nonuniform flow pockets at the slot exit. The compound angle configuration (60° jet orientation and 30° slot inclination angles) was discovered as the best choice with four merits: (i) The acceptable uniform flow can be reached as shallow as h/d = 1.0 (slot depth over jet diameter), (ii) The range of the acceptable uniform flow was the broadest (form y/d = 1.0 to y/d = 2.8). (iii) Flow uniformity was the best at the exit, (iv) The total pressure loss was the lowest.

This paper experimentally investigates the effect of rotation on heat transfer in typical turbine blade serpentine coolant passage with ribbed walls at low Mach numbers. To achieve the low Mach number (around 0.01) condition, pressurized Freon R-134a vapor is utilized as the working fluid. The flow in the first passage is radial outward, after the 180° tip turn the flow is radial inward to the second passage, and after the 180° hub turn the flow is radial outward to the third passage. The effects of rotation on the heat transfer coefficients were investigated at rotation numbers up to 0.6 and Reynolds numbers from 30,000 to 70,000. Heat transfer coefficients were measured using the thermocouples-copper-plate-heater regional average method. Heat transfer results are obtained over a wide range of Reynolds numbers and rotation numbers. An increase in heat transfer rates due to rotation is observed in radially outward passes; a reduction in heat transfer rate is observed in the radially inward pass. Regional heat transfer coefficients are correlated with Reynolds numbers for non-rotation and with rotation numbers for rotating condition, respectively. The results can be useful for understanding real rotor blade coolant passage heat transfer under low Mach number, medium-high Reynolds number and high rotation number conditions.

Flow in the internal three-pass serpentine rib turbulated passages of an advanced high pressure rotor blade is simulated on a 1:1 scale in the laboratory. Tests to measure the effect of rotation on the Nusselt number are conducted at rotation numbers up to 0.4 and Reynolds numbers from 75,000 to 165,000. To achieve this similitude, pressurized Freon R134a vapor is utilized as the working fluid. Experimental heat transfer coefficient measurements are made using the copperplate regional average method. Regional heat transfer coefficients are correlated with rotation numbers. An increase in heat transfer rates due to rotation is observed in radially outward passes; a reduction in heat transfer rate is observed in the radially inward pass. Strikingly, a significant deterioration in heat transfer is noticed in the “hub” region — between the radially inward second pass and the radially outward third pass. This heat transfer reduction is critical for turbine cooling designs.

The present study investigates a new class of in-wall mesh cooling techniques intended to produce significant wetted surface area heat transfer enhancement products, h A A wetted , with low to moderate pressure losses. Cooling networks (meshes) are presented using round pins and rounded diamond shaped pins with H/D p ratios of 0.2 and S/D p ratios of 1.5, as well as less dense rounded diamond pins of smaller diameter with H/D p of 0.3 and S/D p of 2.14. These geometries differ substantially from conventional pin fin arrays in which H/D p ≥ 1 and S/D p ratios are typically about 2.5. The objective of mesh cooling is to provide highly effective cooling inside component walls while maintaining very thin outer protective skin thickness. Special attention is paid to the combination of techniques including pin meshes, turbulators, and concavity arrays that can provide multiple design solutions for heat transfer and pressure loss goals. Average channel wetted-area heat transfer capabilities exceeding 3 times that of smooth channels have been demonstrated using actual surface area enhancements of no more than 20%. This cooling capability increase is also realized with a large decrease in channel material solidity, up to 30%, compared to conventional pin fin arrays. Friction factor enhancements in the range of 15 to 25 times that of a smooth channel turbulent flow accompany this performance. This investigation represents the first of its kind to demonstrate feasible in-wall cooling methods that may be realized in practice for investment cast turbine blades.

Over tip leakage in high-pressure turbines contributes to aerodynamic losses and migration of hot gasses towards the tip resulting in increased thermal distress. Consequently, turbine designers continue to search for improved blade tip concepts that offer the promise of reducing tip leakage. The present paper deals with the CFD analyses of some such tip configurations. The geometries, patented by GE, are variants of a conventional squealer tip and include (i) a pressure side tip shelf with vertical squealer tip wall, and (ii) a pressure side tip shelf with an inclined squealer tip wall. It is found that the inclined shelf results in separation of flow leaking over the tip resulting in reduced leakage and improved efficiency. The inclined shelf also shows a reduced efficiency derivative with clearance.