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1
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.
2
Nikparto, Ali, and Meinhard T. Schobeiri. "Combined numerical and experimental investigations of heat transfer of a highly loaded low-pressure turbine blade under periodic inlet flow condition." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 232, no. 7 (February 14, 2018): 769–84. http://dx.doi.org/10.1177/0957650918758158.
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This paper experimentally and numerically investigates heat transfer characteristics of a low-pressure turbine blade under steady/unsteady flow conditions. Generally, the low-pressure turbine blades are not exposed to excessive temperatures that require detailed heat transfer predictions. In aircraft engines, they operate at low Re-numbers causing the inception of large separation bubbles on their suction surface. As documented in previous papers, the results of detailed aerodynamic simulations have shown significant discrepancies with experiments. It was the objective of the current investigation to determine the discrepancies between the experimental and numerical heat transfer results. It is shown that small errors in aero-calculation results in large deviations of heat transfer results. The characteristics of the blades mentioned above, make low-pressure turbine blades suitable candidates for evaluating the predictive capability of any numerical method. Documenting the scope of these discrepancies defines the framework of the current paper. The periodic flow inside the gas turbine engine was simulated using the cascade facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. In this study, the wakes that originate from stator blades were simulated by moving rods. The instrumented blade was covered with a liquid crystal sheet and it was used to measure heat transfer coefficient. Reynolds-averaged Navier–Stokes equations were used for numerical investigation purposes. Measurements and simulations were conducted at three different Reynolds numbers (110,000, 150,000, and 250,000). Furthermore, for unsteady flow condition, reduced frequencies of the incoming wakes were varied. The current paper includes a comprehensive heat transfer assessment of the predictive capability of Reynolds-averaged Navier–Stokes based tools. The effect of the separation bubbles on heat transfer is thoroughly discussed in this paper. Comparisons of the experimental and numerical results detail the differences and identify the sources of error that leads to in accurate calculations in terms of predicting heat transfer calculation results.
3
Rodi, W., and G. Scheuerer. "Calculation of Heat Transfer to Convection-Cooled Gas Turbine Blades." Journal of Engineering for Gas Turbines and Power 107, no. 3 (July 1, 1985): 620–27. http://dx.doi.org/10.1115/1.3239781.
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A mathematical model is presented for calculating the external heat transfer coefficients around gas turbine blades. The model is based on a finite-difference procedure for solving the boundary-layer equations which describe the flow and temperature field around the blades. The effects of turbulence are simulated by a low-Reynolds number version of the k-ε turbulence model. This allows calculation of laminar and transitional zones and also the onset of transition. Applications of the calculation method are presented to turbine-blade situations which have recently been investigated experimentally. Predicted and measured heat transfer coefficients are compared and good agreement with the data is observed. This is true especially for the pressure-surface boundary layer which is of a rather complex nature because it remains in a transitional state over the full blade length. The influence of various flow phenomena like laminar-turbulent transition and of the boundary conditions (pressure gradient, free-stream turbulence) on the predicted heat transfer rates is discussed.
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Ling, J., Y. Cao, and W. S. Chang. "Analyses of Radially Rotating High-Temperature Heat Pipes for Turbomachinery Applications." Journal of Engineering for Gas Turbines and Power 121, no. 2 (April 1, 1999): 306–12. http://dx.doi.org/10.1115/1.2817121.
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A set of closed-form solutions for the liquid film distributions in the condenser section of a radially rotating miniature heat pipe and for the vapor temperature drop along the heat pipe length are derived. The heat transfer limitations of the heat pipe are analyzed under turbine blade cooling conditions. Analytical results indicate that the condenser heat transfer limitation normally encountered by low-temperature heat pipes no longer exists for the high-temperature rotating heat pipes that are employed for turbine blade cooling. It is found that the heat pipe diameter, radially rotating speed, and operating temperature are very important to the performance of the heat pipe. Heat transfer limitations may be encountered for an increased heat input and rotating speed, or a decreased hydraulic diameter. Based on the extensive analytical evaluations, it is concluded that the radially rotating miniature heat pipe studied in this paper is feasible for turbine blade cooling applications.
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Sadowski, Tomasz, and Daniel Pietras. "Heat Transfer Process in Jet Turbine Blade with Functionally Graded Thermal Barrier Coating." Solid State Phenomena 254 (August 2016): 170–75. http://dx.doi.org/10.4028/www.scientific.net/ssp.254.170.
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In the jet engine the temperature of exhaust gases should be as high as possible, from the point of view of its efficiency. The value of this temperature is limited by toughness of the turbine blades material. Superalloy Inconel 625, which is commonly used in aerospace industry, indicates 13% less yield point in 800OC than in 25OC. The temperature of exhaust gases can reach 1500OC. The blade material has to be protected due to this fact. The one possibility of turbine blade protection is using of thermal barriers coatings (TBC). The coating has a very low thermal conductivity and therefore it protects against the thermal shock failure of the substrate material. The TBC can be manufactured as: 1) monocrystalline, 2) layered structures (e.g. [1-3]) or 3) as a functionally graded material (e.g. [4-7]). The differences between the properties of blade material and TBC can lead to significant increase of the high shear stresses in the substrate-TBC interface.In this paper numerical analyses of cooled turbine blade with various kinds of functionally graded thermal coatings were performed. The main aim was to find the optimal material properties distribution of the functionally graded TBC to avoid damage initiation and growth between TBC and substrate. In the calculations the effect of temperature on material properties both mechanical and thermal was taken into consideration.
6
Camci, C. "An Experimental and Numerical Investigation of Near Cooling Hole Heat Fluxes on a Film-Cooled Turbine Blade." Journal of Turbomachinery 111, no. 1 (January 1, 1989): 63–70. http://dx.doi.org/10.1115/1.3262238.
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Discrete hole film cooling on highly curved surfaces of a gas turbine blade produces very significant wall temperature gradients and wall heat flux variations near downstream and upstream of rows of circular cooling holes. In this study a set of well-defined external heat transfer coefficient distributions in the presence of discrete hole film cooling is presented. Heat transfer coefficients are measured on the suction side of an HP rotor blade profile in a short-duration facility under well-simulated gas turbine flow conditions. The main emphasis of the study is to evaluate the internal heat flux distributions in a detailed way near the cooling holes by using a computational technique. The method uses the measured external heat transfer coefficients as boundary conditions in addition to available internal heat transfer correlations for the internal passages. The study shows the details of the near hole temperature gradients and heat fluxes. The convective heat transfer inside the circular film cooling holes is shown to be very significant even with their relatively small diameter and lengths compared to the chord length. The study also indicates a nonnegligible wall temperature reduction at near upstream of discrete cooling holes. This is explained with the elliptic nature of the internal conduction field of the blade and relatively low coolant temperature levels at the exit of a film cooling hole compared to the mean blade temperature.
7
Krishnamoorthy, V., B. R. Pai, and S. P. Sukhatme. "Influence of Upstream Flow Conditions on the Heat Transfer to Nozzle Guide Vanes." Journal of Turbomachinery 110, no. 3 (July 1, 1988): 412–16. http://dx.doi.org/10.1115/1.3262212.
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The influence of a combustor located just upstream of a nozzle guide vane cascade on the heat flux distribution to the nozzle guide vane was experimentally investigated. The surface temperature distribution around the convectively cooled vane of the cascade was obtained by locating the cascade, firstly in a low-turbulence uniform hot gas stream, secondly in a high-turbulence, uniform hot gas stream, and thirdly in a high-turbulence, nonuniform hot gas stream present just downstream of the combustor exit. The results indicate that the increased blade surface temperatures observed for the cascade placed just downstream of the combustor can be accounted for by the prevailing turbulence level measured at cascade inlet in cold-flow conditions and the average gas temperature at the cascade inlet.
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Xie, Gongnan, and Bengt Sundén. "Comparisons of Heat Transfer Enhancement of an Internal Blade Tip with Metal or Insulating Pins." Advances in Applied Mathematics and Mechanics 3, no. 3 (June 2011): 297–309. http://dx.doi.org/10.4208/aamm.10-10s2-03.
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AbstractCooling methods are needed for turbine blade tips to ensure a long durability and safe operation. A common way to cool a tip is to use serpentine passages with 180-deg turn under the blade tip-cap taking advantage of the three-dimensional turning effect and impingement like flow. Improved internal convective cooling is therefore required to increase the blade tip lifetime. In the present study, augmented heat transfer of an internal blade tip with pin-fin arrays has been investigated numerically using a conjugate heat transfer method. The computational domain includes the fluid region and the solid pins as well as the tip regions. Turbulent convective heat transfer between the fluid and pins, and heat conduction within pins and tip are simultaneously computed. The main objective of the present study is to observe the effect of the pin material on heat transfer enhancement of the pin-finned tips. It is found that due to the combination of turning, impingement and pin-fin crossflow, the heat transfer coefficient of a pin-finned tip is a factor of 2.9 higher than that of a smooth tip at the cost of an increased pressure drop by less than 10%. The usage of metal pins can reduce the tip temperature effectively and thereby remove the heat load from the tip. Also, it is found that the tip heat transfer is enhanced even by using insulating pins having low thermal conductivity at low Reynolds numbers. The comparisons of overall performances are also included.
9
Dunn, Michael G. "Convective Heat Transfer and Aerodynamics in Axial Flow Turbines." Journal of Turbomachinery 123, no. 4 (February 1, 2001): 637–86. http://dx.doi.org/10.1115/1.1397776.
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The primary focus of this paper is convective heat transfer in axial flow turbines. Research activity involving heat transfer generally separates into two related areas: predictions and measurements. The problems associated with predicting heat transfer are coupled with turbine aerodynamics because proper prediction of vane and blade surface-pressure distribution is essential for predicting the corresponding heat transfer distribution. The experimental community has advanced to the point where time-averaged and time-resolved three-dimensional heat transfer data for the vanes and blades are obtained routinely by those operating full-stage rotating turbines. However, there are relatively few CFD codes capable of generating three-dimensional predictions of the heat transfer distribution, and where these codes have been applied the results suggest that additional work is required. This paper outlines the progression of work done by the heat transfer community over the last several decades as both the measurements and the predictions have improved to current levels. To frame the problem properly, the paper reviews the influence of turbine aerodynamics on heat transfer predictions. This includes a discussion of time-resolved surface-pressure measurements with predictions and the data involved in forcing function measurements. The ability of existing two-dimensional and three-dimensional Navier–Stokes codes to predict the proper trends of the time-averaged and unsteady pressure field for full-stage rotating turbines is demonstrated. Most of the codes do a reasonably good job of predicting the surface-pressure data at vane and blade midspan, but not as well near the hub or the tip region for the blade. In addition, the ability of the codes to predict surface-pressure distribution is significantly better than the corresponding heat transfer distributions. Heat transfer codes are validated against measurements of one type or another. Sometimes the measurements are performed using full rotating rigs, and other times a much simpler geometry is used. In either case, it is important to review the measurement techniques currently used. Heat transfer predictions for engine turbines are very difficult because the boundary conditions are not well known. The conditions at the exit of the combustor are generally not well known and a section of this paper discusses that problem. The majority of the discussion is devoted to external heat transfer with and without cooling, turbulence effects, and internal cooling. As the design community increases the thrust-to-weight ratio and the turbine inlet temperature, there remain many turbine-related heat transfer issues. Included are film cooling modeling, definition of combustor exit conditions, understanding of blade tip distress, definition of hot streak migration, component fatigue, loss mechanisms in the low turbine, and many others. Several suggestions are given herein for research and development areas for which there is potentially high payoff to the industry with relatively small risk.
10
Tafti, Danesh K., Long He, and K. Nagendra. "Large eddy simulation for predicting turbulent heat transfer in gas turbines." Philosophical Transactions of the Royal Society A: Mathematical, Physical and Engineering Sciences 372, no. 2022 (August 13, 2014): 20130322. http://dx.doi.org/10.1098/rsta.2013.0322.
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Blade cooling technology will play a critical role in the next generation of propulsion and power generation gas turbines. Accurate prediction of blade metal temperature can avoid the use of excessive compressed bypass air and allow higher turbine inlet temperature, increasing fuel efficiency and decreasing emissions. Large eddy simulation (LES) has been established to predict heat transfer coefficients with good accuracy under various non-canonical flows, but is still limited to relatively simple geometries and low Reynolds numbers. It is envisioned that the projected increase in computational power combined with a drop in price-to-performance ratio will make system-level simulations using LES in complex blade geometries at engine conditions accessible to the design process in the coming one to two decades. In making this possible, two key challenges are addressed in this paper: working with complex intricate blade geometries and simulating high-Reynolds-number ( Re ) flows. It is proposed to use the immersed boundary method (IBM) combined with LES wall functions. A ribbed duct at Re =20 000 is simulated using the IBM, and a two-pass ribbed duct is simulated at Re =100 000 with and without rotation (rotation number Ro =0.2) using LES with wall functions. The results validate that the IBM is a viable alternative to body-conforming grids and that LES with wall functions reproduces experimental results at a much lower computational cost.
11
Yang, Liang, Li Hua Chai, Lai Qi Zhang, and Jun Pin Lin. "Numerical Simulation and Process Optimization of Investment Casting of the Blades for High Nb Containing TiAl Alloy." Materials Science Forum 747-748 (February 2013): 105–10. http://dx.doi.org/10.4028/www.scientific.net/msf.747-748.105.
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Low pressure turbine blades (LPT) made by investment casting from intermetallic titanium aluminide alloys for aero-engine applications have about 50% weight saving compared with that from nickel-based counterparts. Investment casting process of the low pressure turbine blades for high Nb containing TiAl alloy was simulated by Procast. The height of the blade is about 125mm and the thinnest part of it is about 6mm. Compositions of the cast and mould are Ti-45.5Al-8Nb (at %) and Zircon sand, respectively. The simulation result showed that there were porosities appearing in the centre of blades, which may be due to the formation of isolated liquid. In this work, the simulation, analysis and comparison of different casting ways were carried out. The result showed that compared with top and bottom casting, blades made by side casting have less porosity defects. And then the casting temperature, casting velocity, mould preheating temperature and interface heat transfer coefficient were optimized based on orthogonal design. The result also indicated that the influence of process parameters to porosity defects of blades can be ranked from strong to weak as follow: casting temperature>shell mould preheating temperature>casting velocity>interface heat transfer coefficient. When the casting temperature was 1700, the mould preheating temperature was 500, the casting velocity was 0.5 m·s-1, and the interface heat transfer coefficient was 500 W·m-2·K-1, the volume of porosity defects was the smallest.
12
Liu, Jia Zeng, Jian Min Gao, Tie Yu Gao, and Jiao Jun Shi. "Heat Transfer in Narrow Rectangular Channels with Rib Turbulators." Advanced Materials Research 354-355 (October 2011): 1245–51. http://dx.doi.org/10.4028/www.scientific.net/amr.354-355.1245.
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An experimental study of heat transfer characteristics of narrow rectangular channels with rib turbulators for Re in the range of 10000-60000 was performed. To simulate the actual geometry and heat transfer structure of blade/vane internal cooling passage, each of the test channels was welded by four stainless steel plates. Because of the three dimensional heat conduction in the walls and heat conduction between the ribbed and smooth walls, the measured temperature distribution along the axial direction of the test channel is a smooth continuous curve, and when the Re is low, the average Nu of the ribbed and smooth walls are nearly the same. For each aspect ratio channels, the average Nu for the channel of α=45° is about 15 to 25 percent higher than that of α=60°. In addition, we have developed the semi-empirical correlations, covering the range of Re, to predict heat transfer coefficient of the channels. The correlations can be used in the design of turbine blade/vane cooling channels.
13
Giel, Paul W., Robert J. Boyle, and Ronald S. Bunker. "Measurements and Predictions of Heat Transfer on Rotor Blades in a Transonic Turbine Cascade." Journal of Turbomachinery 126, no. 1 (January 1, 2004): 110–21. http://dx.doi.org/10.1115/1.1643383.
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Detailed heat transfer measurements and predictions are given for a power generation turbine rotor with 127 deg of nominal turning and an axial chord of 130 mm. Data were obtained for a set of four exit Reynolds numbers comprised of the facility maximum point of 2.50×106, as well as conditions which represent 50%, 25%, and 15% of this maximum condition. Three ideal exit pressure ratios were examined including the design point of 1.443, as well as conditions which represent −25% and +20% of the design value. Three inlet flow angles were examined including the design point and ±5deg off-design angles. Measurements were made in a linear cascade with highly three-dimensional blade passage flows that resulted from the high flow turning and thick inlet boundary layers. Inlet turbulence was generated with a blown square bar grid. The purpose of the work is the extension of three-dimensional predictive modeling capability for airfoil external heat transfer to engine specific conditions including blade shape, Reynolds numbers, and Mach numbers. Data were obtained by a steady-state technique using a thin-foil heater wrapped around a low thermal conductivity blade. Surface temperatures were measured using calibrated liquid crystals. The results show the effects of strong secondary vortical flows, laminar-to-turbulent transition, and also show good detail in the stagnation region.
14
Tafti, Danesh, Cody Dowd, and Xiaoming Tan. "High Reynold Number LES of a Rotating Two-Pass Ribbed Duct." Aerospace 5, no. 4 (November 23, 2018): 124. http://dx.doi.org/10.3390/aerospace5040124.
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Cooling of gas turbine blades is critical to long term durability. Accurate prediction of blade metal temperature is a key component in the design of the cooling system. In this design space, spatial distribution of heat transfer coefficients plays a significant role. Large-Eddy Simulation (LES) has been shown to be a robust method for predicting heat transfer. Because of the high computational cost of LES as Reynolds number (Re) increases, most investigations have been performed at low Re of O(104). In this paper, a two-pass duct with a 180° turn is simulated at Re = 100,000 for a stationary and a rotating duct at Ro = 0.2 and Bo = 0.5. The predicted mean and turbulent statistics compare well with experiments in the highly turbulent flow. Rotation-induced secondary flows have a large effect on heat transfer in the first pass. In the second pass, high turbulence intensities exiting the bend dominate heat transfer. Turbulent intensities are highest with the inclusion of centrifugal buoyancy and increase heat transfer. Centrifugal buoyancy increases the duct averaged heat transfer by 10% over a stationary duct while also reducing friction by 10% due to centrifugal pumping.
15
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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Ren, Jing, Xueying Li, and Hongde Jiang. "Conjugate Heat Transfer Characteristics in a Highly Thermally Loaded Film Cooling Configuration with TBC in Syngas." Aerospace 6, no. 2 (February 13, 2019): 16. http://dx.doi.org/10.3390/aerospace6020016.
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Future power equipment tends to take hydrogen or middle/low heat-value syngas as fuel for low emission. The heat transfer of a film-cooled turbine blade shall be influenced more by radiation. Its characteristic of conjugate heat transfer is studied experimentally and numerically in the paper by considering radiation heat transfer, multicomposition gas, and thermal barrier coating (TBC). The Weighted Sum of Gray Gases Spectral Model and the Discrete Transfer Model are utilized to solve the radiative heat transfer in the multicomposition field, while validated against the experimental data for the studied cases. It is shown that the plate temperature increases significantly when considering the radiation and the temperature gradient of the film-cooled plate becomes less significant. It is also shown that increasing percentage of steam in gas composition results in increased temperature on the film-cooled plate. The normalized temperature of the film-cooled plate decreases about 0.02, as the total percentage of steam in hot gas increases 7%. As for the TBC effect, it can smooth out the temperature distribution and insulate the heat to a greater extent when the radiative heat transfer becomes significant.
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Sadowski, Tomasz, and Przemysław Golewski. "The Analysis of Heat Transfer and Thermal Stresses in Thermal Barrier Coatings under Exploitation." Defect and Diffusion Forum 326-328 (April 2012): 530–35. http://dx.doi.org/10.4028/www.scientific.net/ddf.326-328.530.
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Effectiveness of internal combustion turbines in aero-engines is limited by comparatively low temperature of exhaust gas at the entry to turbine of the engine. A thermal efficiency and other capacities of turbine strongly depend on the ratio of the highest to the lowest temperature of a working medium. Continuous endeavour to increase the thermal resistance of engine elements requires, apart from laboratory investigations, also numerical studies in 3D of different aero-engine parts. In the present work, the effectiveness of the protection of turbine blades by thermal barrier coating and internal cooling under thermal shock cooling was analysed numerically using the ABAQUS code. The phenomenon of heating the blade from temperature of combustion gases was studied. This investigation was preceded by the CFD analysis in the ANSYS Fluent program which allows for calculation of the temperature of combustion gases. The analysis was conducted for different levels of the shock temperature, different thickness of applied TBC, produced from different kinds of materials.
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Xu, Liang, Qingyun Shen, Qicheng Ruan, Lei Xi, Jianmin Gao, and Yunlong Li. "Optimization Design of Lattice Structures in Internal Cooling Channel of Turbine Blade." Applied Sciences 11, no. 13 (June 23, 2021): 5838. http://dx.doi.org/10.3390/app11135838.
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Recently, the inlet temperatures in gas turbine units have been drastically increased, which extremely affects the lifespan of gas turbine blades. Traditional cooling structures greatly improve the high temperature resistance of the blade; however, these structures scarcely concern both heat transfer and mechanical performances. Lattice structure (LS) can realize these requirements because of its characteristics of light weight, high strength, and porosity. Although the topology of LS is complex, it can be manufactured with the 3D metal printing technology. In this study, an integral optimization method of lattice cooling structure, used at the trailing edge of turbine blades, concerned with heat transfer and mechanical performance, was presented. Firstly, functions between the first-order natural frequency (freq1), elasticity modulus (E), relative density (ρ¯), and Nusselt number (Nu), and the geometric variables of pyramid type LS (PLS) and X-type LS (XLS) were established, and the reliability of these functions was verified. Then, a mathematical optimization model was developed based on these functions which contained two selected optimization problems. Finally, relations among objectives were analyzed; influence law of geometric variables to objectives were discussed, and the accuracy of the optimal LS was proved by experiment and numerical simulation. The optimization results suggest that, compared to the initial LS, Nu increases by 24.1% and ρ¯ decreases by 31% in the optimal LS of the first selected problem, and the Nu increases by 28.8% while freq1 and ρ¯ are almost unchanged in the optimal LS of the second selected problem compared to the initial LS. This study may provide a guidance for functions integration design of lattice cooling structures used at turbine blades based on 3D printing.
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Karabay, Hasan, Robert Pilbrow, Michael Wilson, and J. Michael Owen. "Performance of Pre-Swirl Rotating-Disc Systems." Journal of Engineering for Gas Turbines and Power 122, no. 3 (January 3, 2000): 442–50. http://dx.doi.org/10.1115/1.1285838.
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This paper summarizes and extends recent theoretical, computational, and experimental research into the fluid mechanics, thermodynamics, and heat transfer characteristics of the so-called cover-plate pre-swirl system. Experiments were carried out in a purpose-built rotating-disc rig, and the Reynolds-averaged Navier-Stokes equations were solved using two-dimensional (axisymmetric) and three-dimensional computational codes, both of which incorporated low-Reynolds-number k-ε turbulence models. The free-vortex flow, which occurs inside the rotating cavity between the disc and cover-plate, is controlled principally by the pre-swirl ratio, βp: this is the ratio of the tangential velocity of the air leaving the nozzles to that of the rotating disc. Computed values of the tangential velocity are in good agreement with measurements, and computed distributions of pressure are in close agreement with those predicted by a one-dimensional theoretical model. It is shown theoretically and computationally that there is a critical pre-swirl ratio, βp,crit, for which the frictional moment on the rotating discs is zero, and there is an optimal pre-swirl ratio, βp,opt, where the average Nusselt number is a minimum. Computations show that, for βp<βp,opt, the temperature of the blade-cooling air decreases as βp increases; for βp>βp,opt, whether the temperature of the cooling air increases or decreases as βp increases depends on the flow conditions and on the temperature difference between the disc and the air. Owing to the three-dimensional flow and heat transfer near the blade-cooling holes, and to unquantifiable uncertainties in the experimental measurements, there were significant differences between the computed and measured temperatures of the blade-cooling air. In the main, the three-dimensional computations produced smaller differences than the two-dimensional computations. [S0742-4795(00)01902-5]
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Krishnaswamy, Karthik, and Srikanth Salyan. "Effect of Discrete Ribs on Heat Transfer and Friction Inside Narrow Rectangular Cross Section Cooling Passage of Gas Turbine Blade." International Journal of Engineering and Advanced Technology 10, no. 6 (August 30, 2021): 192–209. http://dx.doi.org/10.35940/ijeat.f3074.0810621.
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The performance of a gas turbine during the service life can be enhanced by cooling the turbine blades efficiently. The objective of this study is to achieve high thermohydraulic performance (THP) inside a cooling passage of a turbine blade having aspect ratio (AR) 1:5 by using discrete W and V-shaped ribs. Hydraulic diameter (Dh) of the cooling passage is 50 mm. Ribs are positioned facing downstream with angle-of-attack (α) of 30° and 45° for discrete W-ribs and discerte V-ribs respectively. The rib profiles with rib height to hydraulic diameter ratio (e/Dh) or blockage ratio 0.06 and pitch (P) 36 mm are tested for Reynolds number (Re) range 30000-75000. Analysis reveals that, area averaged Nusselt numbers of the rib profiles are comparable, with maximum difference of 6% at Re 30000, which is within the limits of uncertainty. Variation of local heat transfer coefficients along the stream exhibited a saw tooth profile, with discrete W-ribs exhibiting higher variations. Along spanwise direction, discrete V-ribs showed larger variations. Maximum variation in local heat transfer coefficients is estimated to be 25%. For experimented Re range, friction loss for discrete W-ribs is higher than discrete-V ribs. Rib profiles exhibited superior heat transfer capabilities. The best Nu/Nuo achieved for discrete Vribs is 3.4 and discrete W-ribs is 3.6. In view of superior heat transfer capabilities, ribs can be deployed in cooling passages near the leading edge, where the temperatures are very high. The best THPo achieved is 3.2 for discrete V-ribs and 3 for discrete W-ribs at Re 30000. The ribs can also enhance the power-toweight ratio as they can produce high thermohydraulic performances for low blockage ratios.
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Chambers, Andrew C., David R. H. Gillespie, Peter T. Ireland, and Geoffrey M. Dailey. "The Effect of Initial Cross Flow on the Cooling Performance of a Narrow Impingement Channel." Journal of Heat Transfer 127, no. 4 (March 30, 2005): 358–65. http://dx.doi.org/10.1115/1.1800493.
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Impingement channels are often used in turbine blade cooling configurations. This paper examines the heat transfer performance of a typical integrally cast impingement channel. Detailed heat transfer coefficient distributions on all heat transfer surfaces were obtained in a series of low temperature experiments carried out in a large-scale model of a turbine cooling system using liquid crystal techniques. All experiments were performed on a model of a 19-hole, low aspect ratio impingement channel. The effect of flow introduced at the inlet to the channel on the impingement heat transfer within the channel was investigated. A novel test technique has been applied to determine the effect of the initial cross flow on jet penetration. The experiments were performed at an engine representative Reynolds number of 20,000 and examined the effect of additional initial cross flow up to 10 percent of the total mass flow. It was shown that initial cross flow strongly influenced the heat transfer performance with just 10 percent initial cross flow able to reduce the mean target plate jet effectiveness by 57 percent.
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Johnson, B. V., J. H. Wagner, G. D. Steuber, and F. C. Yeh. "Heat Transfer in Rotating Serpentine Passages With Trips Skewed to the Flow." Journal of Turbomachinery 116, no. 1 (January 1, 1994): 113–23. http://dx.doi.org/10.1115/1.2928265.
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Experiments were conducted to determine the effects of buoyancy and Coriolis forces on heat transfer in turbine blade internal coolant passages. The experiments were conducted with a large-scale, multipass, heat transfer model with both radially inward and outward flow. Trip strips, skewed at 45 deg to the flow direction, were machined on the leading and trailing surfaces of the radial coolant passages. An analysis of the governing flow equations showed that four parameters influence the heat transfer in rotating passages: coolant-to-wall temperature ratio, rotation number, Reynolds number, and radius-to-passage hydraulic diameter ratio. The first three of these four parameters were varied over ranges that are typical of advanced gas turbine engine operating conditions. Results were correlated and compared to previous results from similar stationary and rotating models with smooth walls and with trip strips normal to the flow direction. The heat transfer coefficients on surfaces, where the heat transfer decreased with rotation and buoyancy, decreased to as low as 40 percent of the value without rotation. However, the maximum values of the heat transfer coefficients with high rotation were only slightly above the highest levels previously obtained with the smooth wall model. It was concluded that (1) both Coriolis and buoyancy effects must be considered in turbine blade cooling designs with trip strips, (2) the effects of rotation are markedly different depending upon the flow direction, and (3) the heat transfer with skewed trip strips is less sensitive to buoyancy than the heat transfer in models with either smooth walls or normal trips. Therefore, skewed trip strips rather than normal trip strips are recommended and geometry-specific tests will be required for accurate design information.
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Khetib, Yacine, Ahmad Aziz Alahmadi, Ali Alzaed, Ahamd Tahmasebi, Mohsen Sharifpur, and Goshtasp Cheraghian. "Natural Convection and Entropy Generation of MgO/Water Nanofluids in the Enclosure under a Magnetic Field and Radiation Effects." Processes 9, no. 8 (July 24, 2021): 1277. http://dx.doi.org/10.3390/pr9081277.
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The authors of the present paper sought to conduct a numerical study on the convection heat transfer, along with the radiation and entropy generation (EGE) of a nanofluids (NFs) in a two and three-dimensional square enclosure, by using the FVM. The enclosure contained a high-temperature blade in the form of a vertical elliptical quadrant in the lower corner of the enclosure. The right edge of the enclosure was kept at low temperature, while the other edges were insulated. The enclosure was subjected to a magnetic field (MGF) and could be adjusted to different angles. In this research, two laboratory relationships dependent on temperature and volume fraction were used to simulate thermal conductivity and viscosity. The variables of this problem were Ra, Ha, RAP, nanoparticle (NP) volume fraction, blade aspect ratio, enclosure angles, and MGF. Evaluating the effects of these variables on heat transfer rate (HTR), EGE, and Be revealed that increasing the Ra and reducing the Ha could increase the HTR and EGE. On the other hand, adding radiation HTR to the enclosure increased the overall HTR. Moreover, an augmentation of the volume fraction of magnesium oxide NPs led to an increased amount of HTR and EGE. Furthermore, any changes to the MGF and the enclosure angle imposed various effects on the HTR. The results indicated that an augmentation of the size of the blade increased and then decreased the HTR and the generated entropy. Finally, increasing the blade always increased the Be.
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Xie, Yong-Hui, Dong-Ting Ye, and Zhong-Yang Shen. "Numerical study on film cooling and convective heat transfer characteristics in the cutback region of turbine blade trailing edge." Thermal Science 20, suppl. 3 (2016): 643–49. http://dx.doi.org/10.2298/tsci16s3643x.
Full textAbstract:
Gas turbine blade trailing edge is easy to burn out under the exposure of high-temperature gas due to its thin shape. The cooling of this area is an important task in gas turbine blade design. The structure design and analysis of trailing edge is critical because of the complexity of geometry, arrangement of cooling channels, design requirement of strength, and the working condition of high heat flux. In the present paper, a 3-D model of the trailing edge cooling channel is constructed and both structures with and without land are numerically investigated at different blowing ratio. The distributions of film cooling effectiveness and convective heat transfer coefficient on cutback and land surface are analyzed, respectively. According to the results, it is obtained that the distributions of film cooling effectiveness and convective heat transfer coefficient both show the symmetrical characteristics as a result of the periodic structure of the trailing edge. The increase of blowing ratio significantly improves the film cooling effectiveness and convective heat transfer coefficient on the cutback surface, which is beneficial to the cooling of trailing edge. It is also found that the land structure is advantageous for enhancing the streamwise film cooling effectiveness of the trailing edge surface while the film cooling effectiveness on the land surface remains at a low level. Convective heat transfer coefficient exhibits a strong dependency with the blowing ratio, which suggests that film cooling effectiveness and convective heat transfer coefficient must be both considered and analyzed in the design of trailing edge cooling structure.
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Teng, Shuye, Dong Kee Sohn, and Je-Chin Han. "Unsteady Wake Effect on Film Temperature and Effectiveness Distributions for a Gas Turbine Blade." Journal of Turbomachinery 122, no. 2 (February 1, 1999): 340–47. http://dx.doi.org/10.1115/1.555457.
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The film effectiveness and coolant jet temperature profile on the suction side of a gas turbine blade were measured using a transient liquid crystal and a cold-wire technique, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3×105. Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.6 to 1.2. Wake Strouhal number was kept at 0 and 0.1. Results show that unsteady wake reduces film cooling effectiveness. Results also show that film injection enhances local heat transfer coefficient while the unsteady wake promotes earlier boundary-layer transition. The development of coolant jet temperature profiles could be used to explain the film cooling performance. [S0889-504X(00)00402-5]
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Sargison, J. E., S. M. Guo, M. L. G. Oldfield, G. D. Lock, and A. J. Rawlinson. "A Converging Slot-Hole Film-Cooling Geometry—Part 1: Low-Speed Flat-Plate Heat Transfer and Loss." Journal of Turbomachinery 124, no. 3 (July 1, 2002): 453–60. http://dx.doi.org/10.1115/1.1459735.
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This paper presents experimental measurements of the performance of a new film-cooling hole geometry—the con¯vergings¯lot-hole¯ or console. This novel, patented geometry has been designed to improve the heat transfer and aerodynamic loss performance of turbine vane and rotor blade cooling systems. The physical principles embodied in the new hole design are described, and a typical example of the console geometry is presented. The cooling performance of a single row of consoles was compared experimentally with that of typical 35-deg cylindrical and fan-shaped holes and a slot, on a large-scale, flat-plate model at engine representative Reynolds numbers in a low-speed tunnel with ambient temperature main flow. The hole throat area per unit width is matched for all four hole geometries. By independently varying the temperature of the heated coolant and the heat flux from an electrically heated, thermally insulated, constant heat flux surface, both the heat transfer coefficient and the adiabatic cooling effectiveness were deduced from digital photographs of the color play of narrow-band thermochromic liquid crystals on the model surface. A comparative measurement of the aerodynamic losses associated with each of the four film-cooling geometries was made by traversing the boundary layer at the downstream end of the flat plate. The promising heat transfer and aerodynamic performance of the console geometry have justified further experiments on an engine representative nozzle guide vane in a transonic annular cascade presented in Part 2 of this paper.
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Bohn, D., and J. Gier. "The Effect of Turbulence on the Heat Transfer in Closed Gas-Filled Rotating Annuli." Journal of Turbomachinery 120, no. 4 (October 1, 1998): 824–30. http://dx.doi.org/10.1115/1.2841795.
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Higher turbine inlet temperatures are a common measure for increasing the thermal efficiency of modern gas turbines. This development leads not only to the need for more efficient turbine blade cooling but also to the requirement for a more profound knowledge of the mechanically and thermally stressed parts of the rotor. For determining thermal stresses from the temperature distribution in the rotor of a gas turbine, one has to encounter the convective transfer in rotor cavities. In the special case of an entirely closed gas-filled rotating annulus, the convective flow is governed by a strong natural convection. Owen and other researchers have found that the presence of turbulence and its inclusion in the modeling of the flow causes significant differences in the flow development in rotating annuli with throughflow, e.g., different vortex structures. However, in closed rotating annuli there is still a lack of knowledge concerning the influence of turbulence. Based on previous work, in this paper the influence of turbulence on the flow structure and on the heat transfer is investigated. The flow is investigated numerically with a three-dimensional Navier–Stokes solver, based on a pressure correction scheme. To account for the turbulence, a low-Reynolds-number k–ε model is employed. The results are compared with experiments performed at the Institute of Steam and Gas Turbines. The computations demonstrate that turbulence has a considerable influence on the overall heat transfer as well as on the local heat transfer distribution. Three-dimensional effects are discussed by comparing the three-dimensional calculation with a two-dimensional calculation of the same configuration and are found to have some impact.
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Fois, N., M. Watson, and MB Marshall. "The influence of material properties on the wear of abradable materials." Proceedings of the Institution of Mechanical Engineers, Part J: Journal of Engineering Tribology 231, no. 2 (August 5, 2016): 240–53. http://dx.doi.org/10.1177/1350650116649528.
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In aero-engines it is possible for the blades of the compressor, turbine or fan to incur into their casings. At these interfaces a lining of composite abradable material is used to limit damage to components and thereby sustain the efficiency and longevity of the engine as a whole. These composite materials must have good abradability and erosion resistance. Previously, the wear mechanisms at the contact between the blade and the coating have been characterised using stroboscopic imaging and force measurement on a scaled test-rig platform. This work is focused on the characterisation of the wear mechanism for two different hardnesses of abradable lining. The established stroboscopic imaging technique and contact force measurements are combined with sectioning of the abradable material in order to analyse the material’s response during the tests. A measure of the thermal properties and the resulting temperature of the linings during the test have also been made to further understand the effect of coating hardness. The wear mechanism, material response, contact force and thermal properties of the coating have been used to characterise the different material behaviour with different hardness. At low incursion rates, with a soft coating, the blade tip becomes worn after an initial adhesive transfer from the coating. Post-test sectioning showed blade material and significant compaction present in the coating. The harder coating produced adhesion on the blade tip with solidification observed in the coating. Thermal diffusivity measurements and modelling indicated that thermally driven wear observed was as a consequence of the increased number of boundaries between the metal and hBN phases present interrupting heat flow, leading to a concentration of surface heat. At higher incursion rates, the wear mechanism is more similar between the coatings and a cutting mechanism dominates producing negligible adhesion and blade wear.
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Yu, Jing, Qing Yan Xu, Bai Cheng Liu, Jia Rong Li, and Hai Long Yuan. "Numerical Simulation of Unidirectional Solidification Process of Turbine Blade Castings." Advanced Materials Research 26-28 (October 2007): 947–52. http://dx.doi.org/10.4028/www.scientific.net/amr.26-28.947.
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A mathematical model for three-dimensional simulation of unidirectional solidification process and microstructure evolution of Ni-based superalloy investment castings was developed based on CA-FD method. The modified ray tracing method was used to solve the complicated heat radiation transfer among the multiple blades and outer space during withdrawal process. Various withdrawal rates were used. During one process high withdrawal rate was used first before the platform approached the baffle. Then the low withdrawal rate was used to reduce the temperature difference of the platform in horizontal section and avoid the defects formed in the corner of the platform. The experimental cooling curves of different positions in the blades and microstructure were compared with the simulation results. Both the results showed that the various withdrawal rates process was effective to reduce the temperature difference of the platform and avoid the formation of stray grains. This process could be helpful to increase the productivity.
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Gad-Briggs, A., A. Haslam, and P. Laskaridis. "Effect of change in role of an aircraft on engine life." Aeronautical Journal 117, no. 1196 (October 2013): 1053–70. http://dx.doi.org/10.1017/s000192400000868x.
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Abstract New aircraft require years of development from concept to realisation and can be prone to delays. Consequently, military operators take existing fleets and operate them in a different role. The objective of this study is to examine the effect of operating a typical low bypass military fast jet engine, originally designed for a European theatre, in a hot and harsh climate. The specific purpose is to determine the effect on the high-pressure turbine blade life and the life- cycle cost of the engine. A mission profile and respective performance conditions were analysed and modelled using an in-house performance tool. The flow conditions were simulated using ANSYS® FLUENT. A conjugated heat transfer solution was adopted to determine the blade metal temperature. The blade was modelled physically in 3D using SIMULIA® ABAQUS FEA software. The stresses were derived and used to calculate the temperature coupled low cycle fatigue and creep life. A deterioration case was also studied to evaluate the effect of sand and dust ingestion. There was a significant life reduction of approximately 50% due to creep. The reduction in life was inversely proportional to the life cycle cost of the engine depending on the operating conditions. The results were compared with similar engines and summarised in the context of airworthiness regulations and component integrity.
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Hänni, Dominic, Rainer Schädler, Reza Abhari, Anestis Kalfas, Gregor Schmid, Ewald Lutum, and Nicolas Lecoq. "Purge flow effects on rotor hub endwall heat transfer with extended endwall contouring into the disk cavity." Journal of the Global Power and Propulsion Society 3 (May 13, 2019): 555–68. http://dx.doi.org/10.33737/jgpps/109838.
Full textAbstract:
Efficiency improvements for gas turbines are strongly coupled with increasing turbine inlet temperatures. This imposes new challenges for designers for efficient and adequate cooling of turbine components. Modern gas turbines inject bleed air from the compressor into the stator/rotor rim seal cavity to prevent hot gas ingestion from the main flow, while cooling the rotor disk. The purge flow interacts with the main flow field and static pressure field imposed by the turbine blades. This complex interaction causes non-uniform and jet-like penetration of the purge flow into the main flow field, hence affecting the endwall heat transfer on the rotor. To improve the understanding of purge flow effects on rotor hub endwall heat transfer, an unshrouded, high-pressure representative turbine design with 3D blading and extended endwall contouring of the rotor into the cavity seal was tested. The measurements were conducted in the 1.5-stage axial turbine facility LISA at ETH Zurich, where a state-of-the-art measurement setup with a high-speed infrared camera and thermally managed rotor insert was used to perform high-resolution heat transfer measurements on the rotor. Three different purge flow rates were investigated with regard to hub endwall heat transfer. Additionally, steady-state computational fluid dynamics simulations were performed to complement the experiments. It was found that the local heat transfer rate changes up to ±20% depending on the purge flow rate. The main part of the purged air is ejected at the endwall trough location and swept towards the rotor suction side, which is caused by the interaction of main flow and the cavity extended endwall design. The presence of low momentum purge flow locally reduces the heat transfer rate. Changes in adiabatic wall temperature and heat transfer (depending on purge rate) are observed from the platform start up to the cross passage migration of the secondary flow structures.
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Lebedev, V. V., O. V. Lebedev, and A. E. Remizov. "Formation of film cooling on the turbine blade back and pressure side in the case of using V-shaped dimples." VESTNIK of Samara University. Aerospace and Mechanical Engineering 18, no. 4 (January 21, 2020): 96–105. http://dx.doi.org/10.18287/2541-7533-2019-18-4-96-105.
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Alongside the development of methods of intensifying convective heat transfer inside the blade, development of methods of local improvement of the efficiency of film cooling of the blade’s surface is still of immediate interest. The film is formed on the blade surface in conditions of high-camber shape and low initial velocity of the gas flow in the vicinity of the leading edge with its subsequent abrupt acceleration. The paper presents some data on the peculiarities of film formation on the back and pressure side of the blade in the vicinity of the leading edge. Experimental temperature distribution over the adiabatic wall was obtained with the use of a FLIR-E 64501 thermal imager. It was found that the conditions for the film formation on the blade back are more favorable than those on the pressure side. It manifests itself in the fact that optimal blowing parameters on the blade back are considerably lower than those on the pressure side. The use of V-shaped dimples located on the wall immediately behind the holes for blowing was suggested as a measure for local improvement of film cooling efficiency. The efficiencies of film cooling in the formation of a curtain, without the use and with the use of V-shaped dimples behind the holes for blowing were compared. Local improvement of efficiency and uniformity of film cooling distribution with the use of V-shaped dimples behind the holes for blowing was observed.
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Sargison, J. E., S. M. Guo, M. L. G. Oldfield, G. D. Lock, and A. J. Rawlinson. "A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss." Journal of Turbomachinery 124, no. 3 (July 1, 2002): 461–71. http://dx.doi.org/10.1115/1.1459736.
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This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a con¯vergings¯lot-hole¯ or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.
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Tu, Baofeng, Xinyu Zhang, and Jun Hu. "Experimental and Numerical Investigation on Effects of the Steam Ingestion on the Aerodynamic Stability of an Axial Compressor." Entropy 22, no. 12 (December 15, 2020): 1416. http://dx.doi.org/10.3390/e22121416.
Full textAbstract:
In order to investigate the influence of steam ingestion on the aerodynamic stability of a two-stage low-speed axial-flow compressor, multiphase flow numerical simulation and experiment were carried out. The total pressure ratio and stall margin of the compressor was decreased under steam ingestion. When the compressor worked at 40% and 53% of the nominal speed, the stall margin decreased, respectively, by 1.5% and 6.3%. The ingested steam reduced the inlet Mach number and increased the thickness of the boundary layer on the suction surface of the blade. The low-speed region around the trailing edge of the blade was increased, and the flow separation region of the boundary layer on the suction surface of the blade was expanded; thus, the compressor was more likely to enter the stall state. The higher the rotational speed, the more significant the negative influence of steam ingestion on the compressor stall margin. The entropy and temperature of air were increased by steam. The heat transfer between steam and air was continuous in compressor passages. The entropy of the air in the later stage was higher than that in the first stage; consequently, the flow loss in the second stage was more serious. Under the combined action of steam ingestion and counter-rotating bulk swirl distortion, the compressor stability margin loss was more obvious. When the rotor speed was 40% and 53% of the nominal speed, the stall margin decreased by 6.3% and 12.64%, respectively.
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Mahato, N., S. M. Banerjee, R. N. Jana, and S. Das. "MoS2-SiO2/EG hybrid nanofluid transport in a rotating channel under the influence of a strong magnetic dipole (Hall effect)." Multidiscipline Modeling in Materials and Structures 16, no. 6 (June 27, 2020): 1595–616. http://dx.doi.org/10.1108/mmms-12-2019-0232.
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PurposeThe article focuses on the magnetohydrodynamic (MHD) convective flow of MoS2-SiO2 /ethylene glycol (EG) hybrid nanofluid. The effectiveness of Hall current, periodically heating wall and shape factor of nanoparticles on the magnetized flow of hybrid nanocomposite molybdenum disulfide- silicon dioxide (MoS2-SiO2) suspended in ethylene glycol (EG) in a vertical rotating channel under the influence of strong magnetic dipole (Hall effect) and thermal radiation is assessed. One of the channel walls has an oscillatory temperature gradient. Four different shapes (i.e. brick, cylinder, platelet and blade) of nanoparticles disseminated in base fluid (EG) are considered for simulation of the flow.Design/methodology/approachThe analytical solution of governing equations has been presented. Influences of emerging physical parameters on the velocity and temperature profiles, the shear stresses and the rate of heat transfer are pointed out and discussed via graphs and tables.FindingsThe analysis revealed that Hall parameter has suppressing behavior on the velocity profiles within the rotating channel. The impact of nanoparticle shape factor advances the temperature characteristics significantly in the rotating channel. Brick-shape nanoparticles put up relatively low-temperature distribution in the rotating channel. The Hall parameter reduces the amplitudes of the shear stresses at the channel wall. However, the radiation parameter enhances the amplitude of the rate of heat transfer at the channel wall.Social implicationsThe important technical advantage of hybrid composition of nanoparticles as a drug carrier is its stability, high thermal conductivity, high load carrying capacity, etc. The proposed model may be beneficial in biomedical engineering, automobile parts, mineral and cleaning oils manufacturing, rubber and plastic industries.Originality/valueTo the best of our knowledge, there is little or no report on the aspects of assessment of the effectiveness of Hall current and nanoparticle shape factor on an MHD flow and heat transfer of an electrically conducting MoS2-SiO2/EG ethylene glycol-based hybrid nanofluid confined in a vertical channel with periodically varying wall temperature subject to a rotating frame. The present work furnishes a robust benchmark for the dynamics of nanofluids.
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Zhang, Chun-Yi, Jing-Shan Wei, Ze Wang, Zhe-Shan Yuan, Cheng-Wei Fei, and Cheng Lu. "Creep-Based Reliability Evaluation of Turbine Blade-Tip Clearance with Novel Neural Network Regression." Materials 12, no. 21 (October 29, 2019): 3552. http://dx.doi.org/10.3390/ma12213552.
Full textAbstract:
To reveal the effect of high-temperature creep on the blade-tip radial running clearance of aeroengine high-pressure turbines, a distributed collaborative generalized regression extremum neural network is proposed by absorbing the heuristic thoughts of distributed collaborative response surface method and the generalized extremum neural network, in order to improve the reliability analysis of blade-tip clearance with creep behavior in terms of modeling precision and simulation efficiency. In this method, the generalized extremum neural network was used to handle the transients by simplifying the response process as one extremum and to address the strong nonlinearity by means of its nonlinear mapping ability. The distributed collaborative response surface method was applied to handle multi-object multi-discipline analysis, by decomposing one “big” model with hyperparameters and high nonlinearity into a series of “small” sub-models with few parameters and low nonlinearity. Based on the developed method, the blade-tip clearance reliability analysis of an aeroengine high-pressure turbine was performed subject to the creep behaviors of structural materials, by considering the randomness of influencing parameters such as gas temperature, rotational speed, material parameters, convective heat transfer coefficient, and so forth. It was found that the reliability degree of the clearance is 0.9909 when the allowable value is 2.2 mm, and the creep deformation of the clearance presents a normal distribution with a mean of 1.9829 mm and a standard deviation of 0.07539 mm. Based on a comparison of the methods, it is demonstrated that the proposed method requires a computing time of 1.201 s and has a computational accuracy of 99.929% over 104 simulations, which are improvements of 70.5% and 1.23%, respectively, relative to the distributed collaborative response surface method. Meanwhile, the high efficiency and high precision of the presented approach become more obvious with the increasing simulations. The efforts of this study provide a promising approach to improve the dynamic reliability analysis of complex structures.
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Dai, Shijie, Miao Gong, Liwen Wang, and Tao Wang. "Numerical Analysis and Experimental Verification of Optimum Heat Input in Additive Manufacturing of Aero Ultrathin Blade." Mathematical Problems in Engineering 2021 (January 30, 2021): 1–23. http://dx.doi.org/10.1155/2021/1648075.
Full textAbstract:
Heat input is a crucial factor affecting the quality in blade additive manufacturing repairing. First, a moving heat source model was established, and through numerical analysis and experimental comparison, the optimal geometric parameters of the heat source model for ultrathin blade repair were obtained. Second, a heat transfer model is established based on the optimal heat source model. By analyzing the thermophysical properties of different alloys, the heat input range of the blade was calculated. By heat transfer calculation under different heat inputs, the heat transfer model of blade repair was optimized. Then, a mathematical model of additive height is established. The optimized heat transfer model is used to solve the temperature distribution of the additive section with time under different wire feeding speeds through numerical analysis, which further reduced the heat input range. Third, the experiments are carried out based on the results of numerical analysis. The evolution law of the microstructure and heat input rate of the additive manufacturing zone was revealed, and the optimal heat input parameters were obtained. Under the optimal parameters, the segregation zone disappeared; hence, the test data were close to the base metal, and the additive manufacturing zone achieved better quality. The results and methods are of great guiding significance for the optimization design in additive manufacturing repair of the aero blades. The study also contributes to carrying out a series of research on heat transfer of ultrathin nickel-based alloy welding.
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Li, Yan, Ce Sun, Yu Jiang, and Fang Feng. "Scaling Method of the Rotating Blade of a Wind Turbine for a Rime Ice Wind Tunnel Test." Energies 12, no. 4 (February 15, 2019): 627. http://dx.doi.org/10.3390/en12040627.
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In order to research the law of rime ice accretion on different scaling blades surface, a new rime ice scaling method was proposed in this research. According to previous research, there are three kinds of ice types on blade surfaces: rime ice, glaze ice and mixed ice. Under the condition of rime ice, both the freezing fraction and the coefficient of heat transfer between super-cold water droplets and blade are 100%. The heat transfer model of rime ice is simpler than that of glaze ice and mixed ice. In this research, the scaling parameters including flow field, water droplets, temperature, pressure and rotating parameters were defined. The Weber number (We) based on water film thickness as an important parameter was applied in this study. The rotating parameters including rotating speed and radius had been added into the icing scaling method. To verify the effectiveness of the new rime ice scaling method, icing wind tunnel tests were carried out. The NACA0018 airfoil was used for the test blade. Two kinds of scale chord blades were selected, the chord of full-scale blade was 200 mm and of subscale blade was 100 mm. The test temperature was −15 °C. The ice accretion on different scale blades surface were captured by high-speed camera and the icing shapes of different scaling blades were obtained. To quantitatively analyze the similar degree of icing shapes on different scale blades, an evaluation method which included similar degree (Sim) was established based on the typical characteristic parameters proposed by previous research. The results show that the icing shapes of subscale blades are similar to that of full-scale blades. The similar degree is between 75.22% and 93.01%. The icing wind tunnel test indicates that the new rime ice scaling method is an effective method to study the rime ice of large scale rotating blades. This study can be used as a reference for research on anti-icing and de-icing technologies for large-scale HAWTs (Horizontal Axis Wind Turbines).
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Freakley, P. K., and S. R. Patel. "Internal Mixing: A Practical Investigation of the Flow and Temperature Profiles during a Mixing Cycle." Rubber Chemistry and Technology 58, no. 4 (September 1, 1985): 751–73. http://dx.doi.org/10.5254/1.3536091.
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Abstract From the results of mixing trials with a highly instrumented BR Banbury and biconical rotor rheometry of mixed batches, a detailed analysis of flow and mixing characteristics in the region of a rotor wing has been undertaken. An ‘angled spreader blade’ analogy of the rotor wing is proposed as being a viable basis for mathematical modelling. A one-dimensional flow analysis is used, in which power-law flow behavior and isothermal conditions are assumed. Dispersive mixing, which depends on the stress levels generated during mixing, is shown to occur throughout the entire mass of material swept in front of the rotor wing and not simply at the rotor tip. In addition, the stress levels depend more strongly on batch temperature than on rotor speed. High rotor speeds tend to lead to reduced stress levels as a result of the associated rapid rise in batch temperature, although choosing an appropriate fill factor can minimize temperature rise by promoting efficient heat transfer to the cooling water. During each rotor revolution, the rotor wing collects a mass of material from the reservoir between the rotors. This mass of material is then progressively reduced by leakage flow under the rotor tip and flow around the end of the wing, until the revolution is completed by the return of a residue to the reservoir. The flow around the end of the rotor is shown to be consistently greater than the leakage flow, although the ratio can be influenced by both fill factor and rotor speed. At high rotor speeds and low fill factors, it appears that material is retained in the regions of the side frames of the mixer and may give batch inhomogeneity through poor distribution mixing.
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Qu, Mao Hua, Su Ya Sun, and Ping Xi. "Conjugate Heat Transfer Analysis and Design Optimization of Internally Cooling Turbine Blade." Applied Mechanics and Materials 148-149 (December 2011): 862–67. http://dx.doi.org/10.4028/www.scientific.net/amm.148-149.862.
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To improve the cooling efficiency of turbine blade, a multidisciplinary design optimization (MDO) system involving aerodynamics, heat transfer and structures has been developed. In this system, a MDO procedure for a turbine blade with complicate internal structure is performed. The structural size of rib turbulators, partitions and trailing edge cooling slots, which serve as design variables, is used for parametric modeling of three dimensional turbine blade. Conjugate heat transfer analysis is employed to get the temperature of the blade. The temperature in the blade body obtained from former coupled analysis is specified as boundary conditions for structural analysis. Meanwhile, a combined algorithm of multi-island genetic algorithm (MIGA) and sequential quadratic programming (SQP) is applied for optimization in specified space. While the flow rate of cooling air remains unchanged, the maximum and average temperatures of the blade decrease under the condition of meeting the strength requirement. The result shows that the cooling efficiency of turbine blade is improved, and the system exhibits higher stability, feasibility and efficiency for engineering applications.
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Wilcoxon, R. K., and A. Moutsoglou. "Second Law Analysis in Assessing Constant Power Input Systems." Journal of Heat Transfer 113, no. 2 (May 1, 1991): 321–28. http://dx.doi.org/10.1115/1.2910564.
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A criterion for comparing the relative performance of various heat transfer augmentation methods used in constant power input systems is introduced. The analysis is based on the principle of minimizing the rate of total entropy generation. The heat transfer load (HTL), a parameter determined by the operating requirements of the heat dissipating process that indicates the difficulty of the heat transfer duty to be performed, is defined in the present study. By comparing the irreversibility distribution ratio (φ) of various configurations at a given heat transfer load, the most exergy efficient system can be selected. The data for three different types of fin configurations used in two constant power input applications (electronic equipment and internal turbine blade cooling) are utilized in demonstrating the technique. The results indicate which specific fin geometry of the particular configuration type analyzed will transfer the dissipated heat at the specified base surface temperature while requiring the least pumping power. Although the φ versus HTL criterion is applied to only fins in this study, the method can be extended to many other applications such as jet impingement cooling or mass transfer.
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Han, S., and R. J. Goldstein. "Heat Transfer Study in a Linear Turbine Cascade Using a Thermal Boundary Layer Measurement Technique." Journal of Heat Transfer 129, no. 10 (March 1, 2007): 1384–94. http://dx.doi.org/10.1115/1.2754972.
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An experimental system is designed, constructed, and operated to make local measurements of heat transfer from constant-temperature surfaces in a linear turbine cascade. The system includes a number of embedded heaters and a control system to maintain the turbine blades and end walls in the cascade at a uniform temperature. A five-axis measurement system is used to determine temperature profiles normal to the pressure and suction sides of the blades and to the end wall. Extrapolating these measurements close to the surface, the local heat transfer is calculated using Fourier’s law. The system has been tested in the laboratory, and results are shown for the temperature distributions above the surfaces and for the local variations in the Nusselt number on the different surfaces in the cascade. The system can also be used to study the heat and mass transfer analogy as considerable data are available for mass transfer results with similar geometries.
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Zhang, L., and J. C. Han. "Combined Effect of Free-Stream Turbulence and Unsteady Wake on Heat Transfer Coefficients From a Gas Turbine Blade." Journal of Heat Transfer 117, no. 2 (May 1, 1995): 296–302. http://dx.doi.org/10.1115/1.2822520.
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The combined effect of free-stream turbulence and unsteady wakes on turbine blade surface heat transfer was studied. The experiments used a five-blade linear cascade in a low-speed wind tunnel facility. A turbulence grid and spoked-wheel type wake generator produced the free-stream turbulence and unsteady wakes. The mainstream Reynolds numbers based on the cascade inlet mean velocity and blade chord length were 100,000, 200,000, and 300,000. Results show that the blade time-averaged heat transfer coefficient depends on the mean turbulence intensity, regardless of whether this mean turbulence intensity is from unsteady wake only, turbulence grid only, or a wake and grid combination. The higher mean turbulence promotes earlier boundary layer transition and causes much higher heat transfer coefficients on the suction surface. It also significantly enhances the heat transfer coefficients on the pressure surface. The unsteady wake greatly affects blade heat transfer for low oncoming free-stream turbulence; however, the wake effect diminishes for high oncoming turbulence. The free-stream turbulence also strongly affects blade heat transfer for a low wake passing frequency, but the oncoming turbulence effect diminishes for a high unsteady wake condition.
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Pearce, Robert, Peter Ireland, and Edwin Dane. "Influence of HTC levels on temperature and stress levels in a leading edge impingement system." Journal of the Global Power and Propulsion Society 3 (January 31, 2019): WLAL1F. http://dx.doi.org/10.22261/jgpps.wlal1f.
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Accurate analysis of the performance of a turbine blade cooling system is essential to allow the blade life to be safely predicted. The latter is essential as the business model for an engine can be strongly dependent on the duration between engine shop visits. Some recent heat transfer research has focused on increasing heat transfer levels in order to reduce turbine blade metal temperatures, however for engine designers it is the life of the blade, determined in part by the stress levels within it, that are of main concern. This paper uses heat transfer and stress analysis within the same software environment to examine the influence of the HTC levels in different regions of an engine representative leading edge impingement cooling system on both metal temperature and stress levels. The results of these analyses are then combined to show that, with attention to cooling in different regions of the blade, reductions in stress levels of 6% can be achieved in the most highly stressed regions of the blade with achievable alterations in heat transfer levels.
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Teng, Shuye, Je-Chin Han, and G. M. S. Azad. "Detailed Heat Transfer Coefficient Distributions on a Large-Scale Gas Turbine Blade Tip." Journal of Heat Transfer 123, no. 4 (December 5, 2000): 803–9. http://dx.doi.org/10.1115/1.1373655.
Full textAbstract:
Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in a large-scale, low-speed wind tunnel facility. Tests were made on a five-blade linear cascade. The low-speed wind tunnel is designed to accommodate the 107.49 deg turn of the blade cascade. The mainstream Reynolds number based on cascade exit velocity was 5.3×105. Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. The central blade had a variable tip gap clearance. Measurements were made at three different tip gap clearances of about 1.1 percent, 2.1 percent, and 3 percent of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface using a transient liquid crystal technique. Results show that reduced tip clearance leads to reduced heat transfer coefficient over the blade tip surface. Results also show that reduced tip clearance tends to weaken the unsteady wake effect on blade tip heat transfer.
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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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Rajadas, J. N., A. Chattopadhyay, N. Pagaldipti, and S. Zhang. "Shape optimization of turbine blades with the integration of aerodynamics and heat transfer." Mathematical Problems in Engineering 4, no. 1 (1998): 21–42. http://dx.doi.org/10.1155/s1024123x98000702.
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A multidisciplinary optimization procedure, with the integration of aerodynamic and heat transfer criteria, has been developed for the design of gas turbine blades. Two different optimization formulations have been used. In the first formulation, the maximum temperature in the blade section is chosen as the objective function to be minimized. An upper bound constraint is imposed on the blade average temperature and a lower bound constraint is imposed on the blade tangential force coefficient. In the second formulation, the blade average and maximum temperatures are chosen as objective functions. In both formulations, bounds are imposed on the velocity gradients at several points along the surface of the airfoil to eliminate leading edge velocity spikes which deteriorate aerodynamic performance. Shape optimization is performed using the blade external and coolant path geometric parameters as design variables. Aerodynamic analysis is performed using a panel code. Heat transfer analysis is performed using the finite element method. A gradient based procedure in conjunction with an approximate analysis technique is used for optimization. The results obtained using both optimization techniques are compared with a reference geometry. Both techniques yield significant improvements with the multiobjective formulation resulting in slightly superior design.
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Rubinsky, Boris. "MICROSCALE HEAT TRANSFER IN BIOLOGICAL SYSTEMS AT LOW TEMPERATURES." Experimental Heat Transfer 10, no. 1 (January 1997): 1–29. http://dx.doi.org/10.1080/08916159708946531.
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Krivchikov, A. I., O. A. Korolyuk, and O. O. Romantsova. "Heat transfer in crystalline clathrate hydrates at low temperatures." Low Temperature Physics 33, no. 6 (June 2007): 612–16. http://dx.doi.org/10.1063/1.2755205.
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Rashid, Dr Farhan Lafta Rashid, Dr Haider Nadhom Azziz Azziz, and Dr Emad Qasem Hussein Hussein. "Heat Transfer Enhancement in Air Cooled Gas Turbine Blade Using." Journal of Petroleum Research and Studies 8, no. 3 (May 6, 2021): 52–69. http://dx.doi.org/10.52716/jprs.v8i3.230.
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In this paper, an investigation of using corrugated passages instead of circular crosssection passages was achieved in conditions simulate the case in the gas turbine blade coolingusing ANSYS Fluent version (14.5) with Boundary conditions: inlet coolant air temperature of300 K with different air flow Reynolds numbers (191000, 286000 and 382000). Thesurrounding constant hot air temperatures was (1700 K). The numerical simulations was done bysolving the governing equations (Continuity, Reynolds Averaging Navier-stokes and Energyequation) using (k-ε) model in three dimensions by using the FLUENT version (14.5). Thepresent case was simulated by using corrugated passage of 3 m long, internal diameter of 0.3 m,0.01 m groove height and wall thickness of 0.01 m, was compared with circular cross sectionpipe for the same length, diameter and thickness. The temperature, velocity distributioncontours, cooling air temperature distribution, the inner wall surface temperature, and thermalperformance factor at the two passages centerline are presented in this paper. The coolant airtemperature at the corrugated passage centerline was higher than that for circular one by(12.3%), the temperature distribution for the inner wall surface for the corrugated passage islower than circular one by (4.88 %). The coolant air flow velocity seems to be accelerated anddecelerated through the corrugated passage, so it was shown that the thermal performance factoralong the corrugated passage is larger than 1, this is due to the fact that the corrugated wallscreate turbulent conditions and increasing thermal surface area, and thus increasing heat transfercoefficient than the circular case.