Academic literature on the topic 'Low blade temperatures'
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Journal articles on the topic "Low blade temperatures"
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Cheng, Wenjie, Boqin Gu, and Chunlei Shao. "A numerical study on the steady flow in molten salt pump under various conditions for improved hydraulic performance." International Journal of Numerical Methods for Heat & Fluid Flow 27, no. 8 (August 7, 2017): 1870–86. http://dx.doi.org/10.1108/hff-06-2016-0238.
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Purpose This paper aims to figure out the steady flow status in the molten salt pump under various temperatures and blade number conditions, and give good insight on the structure and temperature-dependent efficiencies of all pump cases. Finally, the main objective of present work is to get best working condition and blade numbers for optimized hydraulic performance. Design/methodology/approach The steady flow in the molten salt pump was studied numerically based on the three-dimensional Reynolds-Averaged Navier–Stokes equations and the standard k-ε turbulence model. Under different temperature conditions, the internal flow fields in the pumps with different blade number were systematically simulated. Besides, a quantitative backflow analysis method was proposed for further investigation. Findings With the molten salt fluid temperature, sharply increasing from 160°C to 480°C, the static pressure decreases gently in all pump cases, and seven-blades pump has the least backflow under low flow rate condition. The efficiencies of all pump cases increase slowly at low temperature (about 160 to 320°C), but there is almost no variation at high temperature, and obviously seven-blades pump has the best efficiency and head in all pump cases over the wide range of temperatures. The seven-blades pump has the best performance in all selected pump cases. Originality/value The steady flow in molten salt pumps was systematically studied under various temperature and blade number conditions for the first time. A quantitative backflow analysis method was proposed first for further investigation on the local flow status in the molten salt pump. A definition about the low velocity region in molten salt pumps was built up to account for whether the studied pump gains most energy. This method can help us to know how to improve the efficiencies of molten salt pumps.
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Bachelez, Andreas, and Steven A. Martinez. "Heat Generation by Two Different Saw Blades Used for Tibial Plateau Leveling Osteotomies." Journal of the American Animal Hospital Association 48, no. 2 (March 1, 2012): 83–88. http://dx.doi.org/10.5326/jaaha-ms-5698.
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During tibial plateau leveling osteotomy (TPLO) the saw blade produces frictional heat. The purpose of this study was to evaluate and compare heat generated by two TPLO blade designs (Slocum Enterprises [SE] and New Generation Devices [NDG]), with or without irrigation, on cadaveric canine tibias. Thirty-six paired tibias were used to continuously measure bone temperatures during osteotomy through both cortices (i.e., the cis and trans cortices). Each pair was assigned to either an irrigation or nonirrigation group during osteotomy, and each tibia within a pair was osteotomized using a different saw blade design. Saw blade temperatures were recorded and temperatures were compared for all combinations of blade type, cortex, and irrigation. In the cis cortex group, the SE blade generated more bone heat than the NGD blade (P=0.0258). Significant differences in temperature generation between saw blade types were seen only when the osteotomy site was not irrigated (P=0.0156). For all variables measured, bone and saw blade temperature generation was lower with irrigation (P<0.05). None of the osteotomies performed with either saw blade produced a critical duration of damaging temperature ranges in this study. Although saw blade design and irrigation influence heat generation during the TPLO, the potential for bone thermal damage during TPLO is low. The use of the NGD blade with irrigation is recommended.
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Arakere, N. K. "High-Temperature Fatigue Properties of Single Crystal Superalloys in Air and Hydrogen." Journal of Engineering for Gas Turbines and Power 126, no. 3 (July 1, 2004): 590–603. http://dx.doi.org/10.1115/1.1501075.
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Hot section components in high-performance aircraft and rocket engines are increasingly being made of single crystal nickel superalloys such as PWA1480, PWA1484, CMSX-4, and Rene N-4 as these materials provide superior creep, stress rupture, melt resistance, and thermomechanical fatigue capabilities over their polycrystalline counterparts. Fatigue failures in PWA1480 single crystal nickel-base superalloy turbine blades used in the space shuttle main engine fuel turbopump are discussed. During testing many turbine blades experienced stage II noncrystallographic fatigue cracks with multiple origins at the core leading edge radius and extending down the airfoil span along the core surface. The longer cracks transitioned from stage II fatigue to crystallographic stage I fatigue propagation, on octahedral planes. An investigation of crack depths on the population of blades as a function of secondary crystallographic orientation (β) revealed that for β=45+/−15 deg tip cracks arrested after some growth or did not initiate at all. Finite element analysis of stress response at the blade tip, as a function of primary and secondary crystal orientation, revealed that there are preferential β orientations for which crack growth is minimized at the blade tip. To assess blade fatigue life and durability extensive testing of uniaxial single crystal specimens with different orientations has been tested over a wide temperature range in air and hydrogen. A detailed analysis of the experimentally determined low cycle fatigue properties for PWA1480 and SC 7-14-6 single crystal materials as a function of specimen crystallographic orientation is presented at high temperature (75°F–1800°F) in high-pressure hydrogen and air. Fatigue failure parameters are investigated for low cycle fatigue data of single crystal material based on the shear stress amplitudes on the 24 octahedral and 6 cube slip systems for FCC single crystals. The max shear stress amplitude [Δτmax] on the slip planes reduces the scatter in the low cycle fatigue data and is found to be a good fatigue damage parameter, especially at elevated temperatures. The parameter Δτmax did not characterize the room temperature low cycle fatigue data in high-pressure hydrogen well because of the noncrystallographic eutectic failure mechanism activated by hydrogen at room temperature. Fatigue life equations are developed for various temperature ranges and environmental conditions based on power-law curve fits of the failure parameter with low cycle fatigue test data. These curve fits can be used for assessing blade fatigue life.
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Guijarro, Rubén, Alberto Tapetado, David Sánchez Montero, and Carmen Vázquez. "Cleaving of PMMA Microstructured Polymer Optical Fibers with 3- and 4-Ring Hexagonal Cladding Structures." Polymers 13, no. 9 (April 22, 2021): 1366. http://dx.doi.org/10.3390/polym13091366.
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The cleaving of a novel microstructured polymer optical fiber (mPOF) to obtain an acceptable connectorized fiber end-face is studied. The effect of the blade temperature and the speed of the cutting blade on the end-face is qualitatively assessed. Recently manufactured mPOFs with air-structured 3- and 4-ring hexagonal-like hole cladding structures with outer fiber diameters of around 250 μm are employed. Good quality end-faces can be obtained by cleaving mPOF fibers at room temperature for blade temperatures within the range 60–80 °C and at a low blade speed at 0.5 mm/s. The importance of the blade surface quality is also addressed, being a critical condition for obtaining satisfactory mPOF end-faces after cleaving. From our experiments, up to four fiber cuts with the same razor blade and blade surface can be carried out with acceptable and similar fiber end-face results.
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Wang, Xiaopeng, Wenchen Xu, Peng Xu, Haitao Zhou, Fantao Kong, and Yuyong Chen. "High Nb–TiAl Intermetallic Blades Fabricated by Isothermal Die Forging Process at Low Temperature." Metals 10, no. 6 (June 6, 2020): 757. http://dx.doi.org/10.3390/met10060757.
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In this study, the isothermal die forging process of high Nb–TiAl (Ti-44Al-8Nb-0.2W-0.2B-Y, at.%) alloy blades was simulated using the ABAQUS V6.11 software and the blades were fabricated successfully. The influence of a low forging temperature (lower than 1000 °C) and strain rate on the distributions of effective strain and stress were analyzed. The results indicate that the effective strain exhibits negative temperature sensitivity and positive strain rate sensitivity. The stress exponent (n = 3.02) and the apparent activation energy (Q = 293.381 kJ/mol) of the present alloy suggests that this as-forged high Nb–TiAl alloy exhibits good deformability at low temperatures. With the reduction in strain rate and the increase in forging temperature, the effective stress decreases. Finally, high-quality high Nb–TiAl alloy blades were fabricated using an isothermal die forging technology at a rate of 0.01 mm/s and temperature of 950 °C, chosen on the basis of the simulations results. Scanning electron microscopy (SEM) and electron back scatter diffraction (EBSD) results indicated that the center of the TiAl alloy blade possessed a duplex microstructure, consisting of remnant lamellar colonies and recrystallized γ/B2 grains. The refined α2 laths showed a typical forging flow line feature in the edge position, whereas the γ laths had broken down and recrystallized.
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Cosack, T., L. Pawlowski, S. Schneiderbanger, and S. Sturlese. "Thermal Barrier Coatings on Turbine Blades by Plasma Spraying With Improved Cooling." Journal of Engineering for Gas Turbines and Power 116, no. 1 (January 1, 1994): 272–76. http://dx.doi.org/10.1115/1.2906805.
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Turbine blades were coated with a thermal barrier coating system consisting of an MCrAlY bond coat about 100 μm thick deposited by Low-Pressure Plasma Spraying (LPPS) and a 300 μm thick ZrO2-7 wt. % Y2O3 top coat. The latter was manufactured by both Atmosphere and Temperature Controlled Spraying (ATCS) and Air Plasma Spraying (APS) using internal air cooling through the cooling holes of the turbine blades. Coated blades were submitted to thermal cycling tests in a burner rig with hot gas temperature of 1485°C. In the case of ATCS coated blades the number of cycles until the first spallation at the leading edge of the blade was between 350 and 2400. The number of cycles of the thermal barrier coatings sprayed with internal cooling was between 1200 and 1800. Furnace cycling tests were also carried out with ATCS coated blades at temperatures of 1100 and 1200°C. The results of thermal cycle tests and the investigations of the microstructure are discussed.
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Weber, H. E. "Wave Engine Aerothermodynamic Design." Journal of Engineering for Gas Turbines and Power 114, no. 4 (October 1, 1992): 790–96. http://dx.doi.org/10.1115/1.2906658.
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A method for aerothermodynamic preliminary design of a wave engine is presented. The engine has a centrifugal precompressor for the wave rotor, which feeds high and low-pressure turbines. Three specific wave engine designs are presented. Wave rotor blades are naturally cooled by the ingested air; thus combustion temperatures can be as high as 1900 K. Engine pressure ratios of over 25 are obtained in compact designs. It is shown that placing no nozzles at the end of the rotor blade passages yields the highest cycle efficiencies, which can be over 50 percent. Rotor blades are straight and easily milled, cast, or fabricated.
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Wilson, M., R. Pilbrow, and J. M. Owen. "Flow and Heat Transfer in a Preswirl Rotor–Stator System." Journal of Turbomachinery 119, no. 2 (April 1, 1997): 364–73. http://dx.doi.org/10.1115/1.2841120.
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Conditions in the internal-air system of a high-pressure turbine stage are modeled using a rig comprising an outer preswirl chamber separated by a seal from an inner rotor-stator system. Preswirl nozzles in the stator supply the “blade-cooling” air, which leaves the system via holes in the rotor, and disk-cooling air enters at the center of the system and leaves through clearances in the peripheral seals. The experimental rig is instrumented with thermocouples, fluxmeters, pitot tubes, and pressure taps, enabling temperatures, heat fluxes, velocities, and pressures to be measured at a number of radial locations. For rotational Reynolds numbers of Reφ ≃ 1.2 × 106, the swirl ratio and the ratios of disk-cooling and blade-cooling flow rates are chosen to be representative of those found inside gas turbines. Measured radial distributions of velocity, temperature, and Nusselt number are compared with computations obtained from an axisymmetric elliptic solver, featuring a low-Reynolds-number k–ε turbulence model. For the inner rotor-stator system, the computed core temperatures and velocities are in good agreement with measured values, but the Nusselt numbers are underpredicted. For the outer preswirl chamber, it was possible to make comparisons between the measured and computed values for cooling-air temperatures but not for the Nusselt numbers. As expected, the temperature of the blade-cooling air decreases as the inlet swirl ratio increases, but the computed air temperatures are significantly lower than the measured ones. Overall, the results give valuable insight into some of the heat transfer characteristics of this complex system.
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Ion, Ion, Anibal Portinha, Jorge Martins, Vasco Teixeira, and Joaquim Carneiro. "Analysis of the energetic/environmental performances of gas turbine plant: Effect of thermal barrier coatings and mass of cooling air." Thermal Science 13, no. 1 (2009): 147–64. http://dx.doi.org/10.2298/tsci0901147i.
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Zirconia stabilized with 8 wt.% Y2O3 is the most common material to be applied in thermal barrier coatings owing to its excellent properties: low thermal conductivity, high toughness and thermal expansion coefficient as ceramic material. Calculation has been made to evaluate the gains of thermal barrier coatings applied on gas turbine blades. The study considers a top ceramic coating Zirconia stabilized with 8 wt.% Y2O3 on a NiCoCrAlY bond coat and Inconel 738LC as substrate. For different thickness and different cooling air flow rates, a thermodynamic analysis has been performed and pollutants emissions (CO, NOx) have been estimated to analyze the effect of rising the gas inlet temperature. The effect of thickness and thermal conductivity of top coating and the mass flow rate of cooling air have been analyzed. The model for heat transfer analysis gives the temperature reduction through the wall blade for the considered conditions and the results presented in this contribution are restricted to a two considered limits: (1) maximum allowable temperature for top layer (1200?C) and (2) for blade material (1000?C). The model can be used to analyze other materials that support higher temperatures helping in the development of new materials for thermal barrier coatings.
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Ford, D. A., K. P. L. Fullagar, H. K. Bhangu, M. C. Thomas, P. S. Burkholder, P. S. Korinko, K. Harris, and J. B. Wahl. "Improved Performance Rhenium Containing Single Crystal Alloy Turbine Blades Utilizing PPM Levels of the Highly Reactive Elements Lanthanum and Yttrium." Journal of Engineering for Gas Turbines and Power 121, no. 1 (January 1, 1999): 138–43. http://dx.doi.org/10.1115/1.2816301.
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Turbine inlet temperatures have now approached 1650°C (3000°F) at maximum power for the latest large commercial turbofan engines, resulting in high fuel efficiency and thrust levels approaching or exceeding 445 kN (100,000 lbs.). High reliability and durability must be intrinsically designed into these turbine engines to meet operating economic targets and ETOPS certification requirements. This level of performance has been brought about by a combination of advances in air cooling for turbine blades and vanes, computerized design technology for stresses and airflow, and the development and application of rhenium (Re) containing, high γ’ volume fraction nickel-base single crystal superalloys, with advanced coatings, including prime-reliant ceramic thermal barrier coatings (TBCs). Re additions to cast airfoil superalloys not only improve creep and thermomechanical fatigue strength but also environmental properties, including coating performance. Re slows down diffusion in these alloys at high operating temperatures [1]. At high gas temperatures, several issues are critical to turbine engine performance retention, blade life, and integrity. These are tip oxidation in particular for shroudless blades, internal oxidation for lightly cooled turbine blades, and TBC adherence to both the airfoil and tip seal liner. It is now known that sulfur (S) at levels, <10 ppm but >0.2 ppm in these alloys reduces the adherence of α alumina protective scales on these materials or their coatings by weakening the Van der Waal’s bond between the scale and the alloy substrate. A team approach has been used to develop an improvement to CMSX-41 alloy which contains 3 percent Re, by reducing S and phosphorus (P) levels in the alloy to <2 ppm, combined with residual additions of lanthanum (La) + yttrium (Y) in the range 10-30 ppm. Results from cyclic, burner rig dynamic oxidation testing at 1093°C (2000°F) show thirteen times the number of cycles to initial alumina scale spallation for CMSX-4 [La + Y] compared to standard CMSX-4. A key factor for application acceptance is of course manufacturing cost. The development of improved low reactivity prime coats for the blade shell molds along with a viable, tight dimensional control yttrium oxide core body are discussed. The target is to attain grain yields of single crystal CMSX-4 (ULS) (La + Y) turbine blades and casting cleanliness approaching standard CMSX-4. The low residual levels of La + Y along with a sophisticated homogenisation/solutioning heat treatment procedure result in full solutioning with essentially no residual γ/γ’ eutectic phase, Ni (La, Y) low melting point eutectics, and associated incipient melting pores. Thus, full CMSX-4 mechanical properties are attained. The La assists with ppm chemistry control of the Y throughout the single crystal turbine blade castings through the formation of a continuous lanthanum oxide film between the molten and solidifying alloy and the ceramic core and prime coat of the shell mold. Y and La tie up the <2 ppm but >0.2 ppm residual S in the alloy as very stable Y and La sulfides and oxysulfides, thus preventing diffusion of the S atoms to the alumina scale layer under high temperature, cyclic oxidising conditions. La also forms a stable phosphide. CMSX-4 (ULS) (La + Y) HP shroudless turbine blades will commence engine testing in May 1998.
Dissertations / Theses on the topic "Low blade temperatures"
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Gillespie, David R. H. "Intricate internal cooling systems for gas turbine blading." Thesis, University of Oxford, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365831.
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Antill, Marc. "The effect of repair welds on the high temperature low cycle fatigue behaviour of nickel base superalloy turbine blades." Thesis, University of Bristol, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.297923.
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Lin, Hsin-I., and 林欣熠. "Low-Temperature Growth of Photocatalytic TiO2 on Plastic Fan Blade for Air Purification and its Mechanical Performance." Thesis, 2009. http://ndltd.ncl.edu.tw/handle/00301190415378974848.
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碩士逢甲大學
材料科學所
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Arc ion plating (AIP) beneficial from high cathode ionization rate, simple procedure, low process temperature, high film deposition rate, strong film adhesion and environmental friendly, are employed in this study to establish the coating technique for depositing photocatalytic titanium dioxide layer on poly-butylene terephthalate (PBT) surface. The correlation among deposition parameter, microstructure, crystalline structure and mechanical performance of TiO2 coating were discussed as well. Furthermore, the photodecomposition efficiency of methanol for TiO2-coated fan device and its kinetic behavior to the decomposition of methanol were also measured for evaluating the feasibility of this filter-free air purification fan device. The experimental results indicated that by adjusting the coating parameters including total pressure, cathode current and deposition time, the TiO2 coating with major anatase phase and minor rutile phase strucutre could be successfully fabricated. The deposition rate with single cathodic source could reach 6.0 �慆/h. For mechanical performance, pencil hardness values of TiO2 coated specimens are between 4H to 5H. The coating adhesion by tape test grades 5B, the highest rank of specification. These all indicate that the AIP depositing technique could provide satisfactory mechanical performance of the TiO2 coating. The photodecomposition efficiency of fan blade (without and with TiO2-coating) to methanol was revealed without and with 382.2 nm UV-LED illuminating over 12 hr. Moreover, two separated slopes were found for the curve of methanol concentration as a function of decomposition time, indicating two different activation energies during decomposing methanol. It is believed that methanol gas absorbed on the fan blade and the inner wall of chamber surface rise the difficult in decomposing methanol gas. However, the TiO2-coated fan blade deposited under 0.25 Pa oxygen pressure, 80 A cathode current and 25 min deposition time with UV-LED illuminating show that the optima decomposition time is 3.05 h. Based on the calculation of this particular case, it is found that activation energy of photodecomposition give two individual value, 5.6 KJ/mole and 16.0 KJ/mole, respectively, far lower than the value for other reported catalytic materials could provide (100~500 KJ/mole). The reduced reaction activation energy and ease of chemical reaction is obtained. It is believed that the CO intermediate absorbed on the TiO2 surface retards photocatalytic reaction and consequently two separated activation energies. Taking all coating parameters into consideration, not only the crystallinity degree of the deposits and methanol decomposition efficiency could both be promoted when deposited film came to a certain thickness. It is therefore recommended that a satisfactory film thickness for acquiring the sufficient film crystallinity would be first priority to provide the optima photodecomposition efficiency for methanol gas or other volatile organic compounds when commercializing the photocatalytic fan blade developed in this study.
Conference papers on the topic "Low blade temperatures"
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Chappel, D., H. Howe, and L. Vo. "Abradable seal testing - Blade temperatures during low speed rub event." In 37th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2001. http://dx.doi.org/10.2514/6.2001-3479.
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Kenkare, A. S., and T. M. Kilner. "A Low-Cost Undergraduate Test Rig for Heat Transfer in Turbine Blade Cooling." In ASME 1985 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-gt-156.
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Although turbine blade cooling has consistently led to the use of higher turbine inlet temperatures leading to improved cycle efficiencies, very little of this technology has found its way into undergraduate laboratory work. The cost of modern blade heat transfer research rigs virtually rules out the possibility of introducing this topic in undergraduate teaching laboratories of Universities or Polytechnics in the UK operating within tight budgetary constraints. However, the underlying principles of blade cooling heat transfer may be demonstrated quite easily by using inlet temperatures about half those existing in the actual turbine and the paper describes the design and development of a low-cost blade cooling heat transfer rig. Test results obtained on the ‘model’ rig enable an appreciation of the problems encountered in turbine blade cooling to be made and may serve as a basis for the design and development of more complicated blade cooling systems.
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Van Treuren, Kenneth, Tyler Pharris, and Olivia Hirst. "Using Turbulence Intensity and Reynolds Number to Predict Flow Separation on a Highly Loaded, Low-Pressure Gas Turbine Blade at Low Reynolds Numbers." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75976.
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The low-pressure turbine has become more important in the last few decades because of the increased emphasis on higher overall pressure and bypass ratios. The desire is to increase blade loading to reduce blade counts and stages in the low-pressure turbine of a gas turbine engine. Increased turbine inlet temperatures for newer cycles results in higher temperatures in the low-pressure turbine, especially the latter stages, where cooling technologies are not used. These higher temperatures lead to higher work from the turbine and this, combined with the high loadings, can lead to flow separation. Separation is more likely in engines operating at high altitudes and reduced throttle setting. At the high Reynolds numbers found at takeoff, the flow over a low-pressure turbine blade tends to stay attached. At lower blade Reynolds numbers (25,000 to 200,000), found during cruise at high altitudes, the flow on the suction surface of the low-pressure turbine blades is inclined to separate. This paper is a study on the flow characteristics of the L1A turbine blade at three low Reynolds numbers (60,000, 108,000, and 165,000) and 15 turbulence intensities (1.89% to 19.87%) in a steady flow cascade wind tunnel. With this data, it is possible to examine the impact of Reynolds number and turbulence intensity on the location of the initiation of flow separation, the flow separation zone, and the reattachment location. Quantifying the change in separated flow as a result of varying Reynolds numbers and turbulence intensities will help to characterize the low momentum flow environments in which the low-pressure turbine must operate and how this might impact the operation of the engine. Based on the data presented, it is possible to predict the location and size of the separation as a function of both the Reynolds number and upstream freestream turbulence intensity (FSTI). Being able to predict this flow behavior can lead to more effective blade designs using either passive or active flow control to reduce or eliminate flow separation.
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Sidwell, Vince, and David Darmofal. "A Selective Assembly Method to Reduce the Impact of Blade Flow Variability on Turbine Life." In ASME Turbo Expo 2004: Power for Land, Sea, and Air. ASMEDC, 2004. http://dx.doi.org/10.1115/gt2004-53930.
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A selective assembly method is proposed that decreases the impact of blade passage manufacturing variability on the life of a row of cooled turbine blades. The method classifies turbine blades into groups based on the effective flow areas of the blade passages, then a row of blades is assembled exclusively from blades of a single group. A simplified classification is considered in which blades are divided into low-flow, nominal-flow, and high-flow groups. For rows assembled from the low-flow class, the blade plenum pressure will tend to rise and the individual blade flows will be closer to the design intent than for a single low-flow blade in a randomly-assembled row. Since the blade metal temperature is strongly dependent on the blade flow, selective assembly can lower the metal temperature of the lowest-flowing blades and increase the life of a turbine row beyond what is possible from a randomly-assembled row. Furthermore, the life of a nominal-flow or high-flow row will be significantly increased (relative to a randomly-assembled row) since the life-limiting low-flow blades would not be included in these higher-flowing rows. The impact of selective assembly is estimated using a model of the first turbine rotor of an existing high-bypass turbofan. The oxidation lives of the nominal-flow and high-flow blade rows are estimated to increase approximately 50% and 100% compared to randomly-assembled rows, while the life of the low-flow rows are the same as the randomly-assembled rows. Alternatively, selective assembly can be used to increase turbine inlet temperature while maintaining the maximum blade metal temperatures at random-assembly levels. For the nominal-flow and high-flow classes, turbine inlet temperature increases are estimated to be equivalent to the turbine inlet temperature increases observed over several years of gas turbine technology development.
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Amano, R. S., Krishna Guntur, and Jose Martinez Lucci. "Computational Study of Gas Turbine Blade Cooling Channel." In 2010 14th International Heat Transfer Conference. ASMEDC, 2010. http://dx.doi.org/10.1115/ihtc14-22920.
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It has been a common practice to use cooling passages in gas turbine blade in order to keep the blade temperatures within the operating range. Insufficiently cooled blades are subject to oxidation, to cause creep rupture, and even to cause melting of the material. To design better cooling passages, better understanding of the flow patterns within the complicated flow channels is essential. The interactions between secondary flows and separation lead to very complex flow patterns. To accurately simulate these flows and heat transfer, both refined turbulence models and higher-order numerical schemes are indispensable for turbine designers to improve the cooling performance. Power output and the efficiency of turbine are completely related to gas firing temperature from chamber. The increment of gas firing temperature is limited by the blade material properties. Advancements in the cooling technology resulted in high firing temperatures with acceptable material temperatures. To better design the cooling channels and to improve the heat transfer, many researchers are studying the flow patterns inside the cooling channels both experimentally and computationally. In this paper, the authors present the performance of three turbulence models using TEACH software code in comparison with the experimental values. To test the performance, a square duct with rectangular ribs oriented at 90° and 45° degree and placed at regular intervals. The channel also has bleed holes. The normalized Nusselt number obtained from simulation are validated with that of experiment. The Reynolds number is set at 10,000 for both the simulation and experiment. The interactions between secondary flows and separation lead to very complex flow patterns. To accurately simulate these flows and heat transfer, both refined turbulence models and higher-order numerical schemes are indispensable for turbine designers to improve the cooling performance. The three-dimensional turbulent flows and heat transfer are numerically studied by using several different turbulence models, such as non-linear low-Reynolds number k-omega and Reynolds Stress (RSM) models. In k-omega model the cubic terms are included to represent the effects of extra strain-rates such as streamline curvature and three-dimensionality on both turbulence normal and shear stresses. The finite volume difference method incorporated with the higher-order bounded interpolation scheme has been employed in the present study. The outcome of this study will help determine the best suitable turbulence model for future studies.
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Martin, Evan L., Lesley M. Wright, and Daniel C. Crites. "Computational Investigation of Jet Impingement on Turbine Blade Leading Edge Cooling With Engine-Like Temperatures." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-68811.
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A numerical investigation of leading edge impingement is completed in this study. Impingement onto a half cylinder, concave surface is used to model the leading edge of a modern gas turbine airfoil. The temperature difference between the impinging jet and the target surface is varied from ΔT = 60°F (33.3°C) (typical of traditional laboratory experiments) to ΔT = 1000°F (555.6°C) (representative of temperature differences encountered in modern engines). Over this range of temperatures, the simulations are validated against experimental data and extended to engine-like conditions. In addition to the varying temperatures, the effect of jet Reynolds number is also investigated (Rejet = 5000–25000). The jet geometry is also varied in this investigation to model the effect of jet-to-jet spacing (s/d = 2–8), the effect of jet–to–target surface distance (ℓ/d = 2–8.5), and the effect of target surface diameter (D/d = 3.6 and 5.5). For all simulations the k-ω, Shear Stress Transport (SST) turbulence model is used to simulate the impingement flows. Over the range of flow conditions and geometry variations, the SST model is proven to be effective in predicting leading edge heat transfer coefficients. With multiple direct comparisons between the numerical simulations and existing experimental data, the simulations predict the surface Nusselt numbers within an average of 11% of the experimental data. Furthermore, the predictions indicate the existing correlations developed in low temperature laboratory experiments are sufficient for calculating stagnation region Nusselt numbers under engine-like temperatures.
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Amano, R. S., Krishna Guntur, Jose Martinez Lucci, and Yu Ashitaka. "Study of Flow Through a Stationary Ribbed Channel for Blade Cooling." In ASME Turbo Expo 2010: Power for Land, Sea, and Air. ASMEDC, 2010. http://dx.doi.org/10.1115/gt2010-23031.
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The firing temperature in gas turbine relates itself directly to the power output and the efficiency of the turbine. The higher the firing (operating) temperatures, higher the wall temperature of blades. However, an increase in the firing temperature is limited by the first stage blade material properties. This is because the higher firing temperature may cause a creep rupture, oxidizing, melting and ultimately failing of blades. Prior to blade cooling, the firing temperature was the same as the blade material temperature. Advancements in cooling technology have resulted in high firing temperatures with acceptable material temperatures. To better design the cooling channels and to improve heat transfer, many researchers are studying the flow patterns inside the cooling channels both experimentally and computationally. In this paper, the authors present the performance of three turbulence models using a Computational Fluid Dynamics code in comparison with the experimental values. To test the performance, a square duct was used with rectangular ribs oriented at 90° and 45° degree and placed at regular intervals. The channel also has bleed holes. The wall Nusselt numbers are compared in both the experimental and the computational results after suitable normalization. The Reynolds number is set to 10,000. The interactions between secondary flows and separation lead to very complex flow patterns. To accurately simulate these flows and heat transfer, both refined turbulence models and higher-order numerical schemes are indispensable for turbine designers to improve the cooling performance. The three-dimensional turbulent flows and heat transfer are numerically studied by using several different turbulence models, such as a non-linear low-Reynolds number k-ω and Reynolds Stress (RSM) models. In the k-ω model the cubic terms are included to represent the effects of extra strain-rates such as streamline curvature and three-dimensionality on both normal and shear turbulence stresses. The finite volume difference method incorporated with the higher-order bounded interpolation scheme has been employed in the present study. The outcome of this study helps to determine the best suitable turbulence model for future studies.
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Byerley, Aaron R., and August J. Rolling. "Exploring the Impact of Elevated Turbine Blade Cooling Effectiveness and Turbine Material Temperatures on Gas Turbine Engine Performance and Cost." In ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-44061.
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Since the 1950’s, the turbine inlet temperatures of gas turbine engines have been steadily increasing as engine designers have sought to increase engine thrust-to-weight and reduce fuel consumption. In turbojets and low-bypass turbofan engines, increasing the turbine inlet temperature boosts specific thrust, which in some cases can support supersonic flight without the use of an afterburner. In high-bypass gas turbine engines, increasing the turbine inlet temperature makes possible higher bypass ratios and overall pressure ratios, both of which reduce specific fuel consumption. Increased turbine inlet temperatures, without sacrificing blade life, have been made possible through advances in blade cooling effectiveness and high-temperature turbine blade materials. Investigating the impact of higher turbine inlet temperatures and the corresponding cooling air flow rates on specific thrust, specific fuel consumption, and engine development cost is the subject of this paper. A physics-based cooling effectiveness correlation is presented for linking turbine inlet temperature to cooling flow fraction. Two cases are considered: 1) a low-bypass, mixed-exhaust, non-afterburning turbofan engine intended to support supercruising at Mach 1.5 and 2) a high-bypass, unmixed-exhaust turbofan engine intended to support highly efficient, long range flight at Mach 0.8. For each of these two cases, both baseline and enhanced cooling effectiveness values as well as both baseline and elevated turbine blade material temperatures are considered. Comparing these cases will ensure that students taking courses in preliminary engine design understand why huge research investments continue to be made in turbine blade cooling and advanced, high-temperature turbine blade material development.
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Thibault, D., P. K. Dubois, B. Picard, A. Landry-Blais, J. S. Plante, and M. Picard. "Experimental Assessment of a Sliding-Blade Inside-Out Ceramic Turbine." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-15137.
Full textAbstract:
Abstract In order to reach 40% efficiency, sub-MW turbines must operate in a recuperated gas Brayton cycle at a turbine inlet temperature (TIT) above 1300°C. Current sub-MW turbines have material-related operating temperature limits. Still to this day, there is no cost-effective rotor design which operates at such high temperatures. This paper introduces a novel, sliding-blade, inside-out ceramic turbine (ICT) wheel configuration, which could enable high-efficiency sub-MW recuperated engines to be achieved with cheap monolithic ceramic blades. The inside-out configuration uses a rotating structural hoop, or shroud, to convert centrifugal forces into compressive blade loading. The sliding-blade architecture uses a hub with angled planes on which ceramic blades slide up and down, allowing to match the radial expansion of the structural shroud. This configuration generates low stress values in both ceramic and metallic components and can achieve high tip speeds. A prototype is designed and its reliability is calculated using CARES software. The result is a design which has a single blade probability of failure (Pf) of 0.1% for 1000 h of steady operation. Analyses also demonstrate that reliability is greatly dependent on friction at ceramic-to-metal interfaces. Low friction could lead to acceptable reliability levels for engine applications. The prototype was successfully tested in a laboratory turbine environment at a tip speed of 350 m/s and a TIT of 1100 °C without any damage. These achievements demonstrate the robustness of the sliding-blade ICT configuration. Further research and development will focus on increasing tip speed and TIT to higher values.
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Cosi, Lorenzo, Jonathon Slepski, Steven DeLessio, Michele Taviani, and Amir Mujezinovic´. "Design, Manufacturing and Testing of a New Family of Steam Turbine Low Pressure Stages." In ASME 2007 Power Conference. ASMEDC, 2007. http://dx.doi.org/10.1115/power2007-22056.
Full textAbstract:
New low pressure (LP), stages for variable speed, mechanical drive and geared power generation steam turbines have been developed. The new blade and nozzle designs can be applied to a wide range of turbine rotational speeds and last stage blade annulus areas, thus forming a family of low pressure stages—High Speed (HS) blades and nozzles. Different family members are exact scales of each other and the tip speeds of the corresponding blades within the family are identical. Thus the aeromechanical and aerodynamic characteristics of the individual stages within the family are identical as well. Last stage blades and nozzles have been developed concurrently with the three upstream stages, creating optimised, reusable low pressure turbine sections. These blades represent a step forward in improving speed, mass flow capability, reliability and aerodynamic efficiency of the low pressure stages for the industrial steam turbines. These four stages are designed as a system using the most modern design tools applied on Power Generation and Aircraft Engines turbo-machineries. The aerodynamic performance of the last three stage of the newly designed group will be verified in a full-scale test facility. The last stage blade construction incorporates a three hooks, axial entry dovetail with improved load carrying capability over other blade attachment methods. The next to the last stage blade also uses a three hooks axial entry dovetail, while the two front stage blades employ internal tangential entry dovetails. The last and next to the last stage blades utilize continuous tip coupling via implementation of integral snubber cover while a Z-lock integral cover is employed for the two upstream stages. Low dynamic strains at all operating conditions (off and on resonance speeds) will be validated via steam turbine testing at realistic steam conditions (steam flows, temperatures and pressures). Low load, high condenser pressure operation will also be verified using a three stage test turbine operated in the actual steam conditions as well. In addition, resonance speed margins of the four stages have been verified through full-scale wheel box tests in the vacuum spin cell, thus allowing the application of these stages to Power Generation applications. Stator blades are produced with a manufacturing technology, which combines full milling and electro-discharge machining. This process allows machining of the blades from an integral disc, and thus improving uniformity of the throat distribution. Accuracy of the throat distribution is also improved when compared to the assembled or welded stator blade technology. This paper will discuss the aerodynamic and aeromechanical design, development and testing program completed for this new low pressure stages family.