This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The freestream turbulence intensity is measured to be 13.5% upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a two-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure-sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

This work focuses on the parametric study of film cooling effectiveness on the turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N 2 , CO 2 , and Argon/SF 6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5, to 2.0. There are eight cylindrical holes on the endwall inside the passage. The pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall crossflow passage.

The internal cooling passages of gas turbine blades mostly have varying aspect ratios from one passage to another. However, there are limited data available in the open literature that used a reduced cross section and aspect ratio (AR), after the tip turn. Therefore, the current study presents heat transfer and pressure drop of three different α = 45 deg profiled rib orientations, typical parallel (usual), reversed parallel (unusual), and crisscross patterns in a rotating two-pass rectangular channel with AR = 4:1 and 2:1 in the first radially outward flow and second radially inward flow passages, respectively. For each rib orientation, regional averaged heat transfer results are obtained for both the flow passages with the Reynolds number ranging from 10,000 to 70,000 for the first passage and 16,000 to 114,000 for the second passage with a rotational speed range of 0–400 rpm. This results in the highest rotation number of 0.39 and 0.16 for the first and second passage respectively. The effects of rib orientation, aspect ratio variation, 180-deg tip turn, and rotation number on the heat transfer and pressure drop will be addressed. According to the results, for usual, unusual and crisscross rib patterns, increasing rotation number causes the heat transfer to decrease on the leading surface and increase on the trailing surface for the first passage and vice versa for the second passage. The overall heat transfer enhancement of the usual and unusual rib patterns is higher than the crisscross one. In terms of the pressure losses, the crisscross rib pattern has the lowest and the usual rib pattern has the highest-pressure loss coefficients. When pressure loss and heat transfer enhancement are both taken into account together, the crisscross or unusual rib pattern might be an option to use in the internal cooling method. Therefore, the results can be useful for the turbine blade internal cooling design and heat transfer analysis.

This paper focuses on the measurements of heat transfer enhancement and pressure drop of different pin or fin configurations in a high aspect ratio (AR = 9.57/1.2) channel. Two different pin-fin shapes including circular pins and strip fins were studied. Different pin-fin spacings for circular pins (S/D = 2, 4) and strip fins (S/W = 8, 16) were investigated, respectively. In addition, the thickness effect of the strip fin was included in this study. The regionally averaged heat transfer measurement method was used to acquire the heat transfer coefficients on two opposite featured surfaces within the test channel. For each configuration, the tested Reynolds number was ranging from 20,000 to 80,000. The results indicate that the channel with circular pins has better heat transfer enhancement and higher pressure loss than their strip fins counterparts. However, the strip fins are considered better designs in terms of thermal performance. For the gas turbine designers aim at developing an improved internal cooling feature, this work demonstrates the great potential of the strip fins as a novel and effective cooling design compared with the conventional circular pins.

A realistic internal cooling system of a turbine blade includes both ribs and pin-fins inside the passages to enhance the heat transfer. However, the majority studies in the open literature assessing the heat transfer characteristics on a simplified cooling model by examining ribbed-roughen passages and pin-finned passage separately. This work presents the high-resolution heat transfer coefficients of a scaled realistic turbine blade internal cooling design. The cooling system, using a 3D-printed plastic material, consists of an S-shaped inlet, four serpentine passages (three U-bends) of variable aspect ratio, and the trailing edge ejection. Angled ribs are implemented inside the passages and the elongated fins and pins are used near the trailing edge. Two dust holes are realized on the blade tip, the injections are individually controlled to reflect the realistic coolant flowrate variation inside the entire internal cooling system. The tests are conducted at two Reynolds number, 45,000 and 60,000 based on the hydraulic diameter of the inlet passage. Transient heat transfer technique using thermochromic liquid crystal is applied to obtain the detailed heat transfer characteristic inside the cooling channel. The local and averaged Nusselt numbers are also compared with the correlations in the open literature. This paper provides gas turbine designers the difference of local heat transfer distributions between the realistic and simplified internal cooling designs.

Varying aspect ratio (AR) channels are found in modern gas turbine airfoils for internal cooling purposes. Corresponding experimental data are needed in understanding and assisting the design of advanced cooling systems. The present study features a two-pass rectangular channel with an AR = 4:1 in the first pass with the radial outward flow and an AR = 2:1 in the second pass with the radial inward flow after a 180 deg tip turn. Effects of rib coverage near the tip region are investigated using profiled 45 deg ribs (P/e = 10, e/D h ≈ 0.11, parallel and in-line) with three different configurations: less coverage, medium coverage, and full coverage. The Reynolds number (Re) ranges from 10,000 to 70,000 in the first passage. The highest rotation number achieved was Ro = 0.39 in the first passage and 0.16 in the second passage. Heat transfer coefficients on the internal surfaces were obtained by the regionally averaged copper plate method. The results showed that the rotation effects on both heat transfer and pressure loss coefficient are reduced with an increased rib coverage in the tip turn region. Different rib coverage upstream of the tip turn significantly changes the heat transfer in the turn portion. Heat transfer reduction (up to −27%) on the tip wall was seen at lower Ro. Dependence on the Reynolds number can be seen for this particular design. The combined geometric, rib coverage, and rotation effects should be taken into consideration in the internal cooling design.

This study investigates the effects of blowing ratio, density ratio, and spanwise pitch on the flat plate film cooling from two rows of compound angled cylindrical holes. Two arrangements of two-row compound angled cylindrical holes are tested: (a) the first row and the second row are oriented in staggered and same compound angled direction ( β = +45 deg for the first row and +45 deg for the second row); (b) the first row and the second row are oriented in inline and opposite direction ( β = +45 deg for the first row and −45 deg for the second row). The cooling hole is 4 mm in diameter with an inclined angle of 30 deg. The streamwise row-to-row spacing is fixed at 3d, and the spanwise hole-to-hole (p) is varying from 4d, 6d to 8d for both designs. The film cooling effectiveness measurements were performed in a low-speed wind tunnel in which the turbulence intensity is kept at 6%. There are 36 cases for each design including four blowing ratios (M = 0.5, 1.0, 1.5, and 2.0), three density ratios (DR = 1.0, 1.5, and 2.0), and three hole-to-hole spacing (p/d = 4, 6, and 8). The detailed film cooling effectiveness distributions were obtained by using the steady-state pressure-sensitive paint (PSP) technique. The spanwise-averaged cooling effectiveness are compared over the range of flow parameters. Some interesting observations are discovered including blowing ratio effect strongly depending on geometric design; staggered arrangement of the hole with same orientation does not yield better effectiveness at higher blowing ratio. Currently, film cooling effectiveness correlation of two-row compound angled cylindrical holes is not available, so this study developed the correlations for the inline arrangement of holes with opposing angles and the staggered arrangement of holes with same angles. The results and correlations are expected to provide useful information for the two-row flat plate film cooling analysis.

The present study features a two-pass rectangular channel with an aspect ratio (AR) = 4:1 in the first pass and an AR = 2:1 in the second pass after a 180-deg tip turn. In addition to the smooth-wall case, ribs with a profiled cross section are placed at 60 deg to the flow direction on both the leading and trailing surfaces in both passages ( P/e = 10, e/D h ∼ 0.11, parallel and in-line). Regionally averaged heat transfer measurement method was used to obtain the heat transfer coefficients on all internal surfaces. The Reynolds number (Re) ranges from 10,000 to 70,000 in the first passage, and the rotational speed ranges from 0 to 400 rpm. Under pressurized condition (570 kPa), the highest rotation number achieved was Ro = 0.39 in the first passage and 0.16 in the second passage. The results showed that the turn-induced secondary flows are reduced in an accelerating flow. The effects of rotation on heat transfer are generally weakened in the ribbed case than the smooth case. Significant heat transfer reduction (∼30%) on the tip wall was seen in both the smooth and ribbed cases under rotating condition. Overall pressure penalty was reduced for the ribbed case under rotation. Reynolds number effect was found noticeable in the current study. The heat transfer and pressure drop characteristics are sensitive to the geometrical design of the channel and should be taken into account in the design process.

The effects of upstream injection angle on film cooling effectiveness of a turbine vane end wall with various endwall film-hole designs were examined by applying pressure-sensitive paint (PSP) measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create an additional controllable upstream film to cool the vane end wall. Three potential injection angles were studied: 30 deg, 40 deg, and 50 deg. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, midchord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio (MFR) 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream injection angle.

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 (AOI) 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 = 2–4%) 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.

The combined effects of upstream purge flow, slashface leakage flow, and discrete hole film cooling on turbine blade platform film cooling effectiveness were studied using the pressure sensitive paint (PSP) technique. As a continued study, discrete cylindrical holes were replaced by laidback fan-shaped (10-10-5) holes, which generally provide better film coverages on the endwall. Experiments were done in a five-blade linear cascade. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. A wide range of parameters was evaluated in this study. The coolant-to-mainstream mass flow ratio (MFR) was varied from 0.5%, 0.75%, to 1% for the upstream purge flow. For the platform film cooling holes and slashface gap, average blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1 (close to low-temperature experiments) to 1.5 (intermediate DR) and 2 (close to engine conditions) were also examined. Purge flow swirl effect was studied particularly at a typical swirl ratio (SR) of 0.6. Area-averaged film cooling effectiveness results were compared between cylindrical and fan-shaped holes. The results indicate that the fan-shaped holes provide superior film coverage than cylindrical holes for platform film cooling especially at higher blowing ratios and momentum flux ratios.