Articles de revues : « Film cooled nozzle »

1

Stark, Ralf, Chloé Génin, Christian Mader, Dietmar Maier, Dirk Schneider et Michael Wohlhüter. « Design of a film cooled dual-bell nozzle ». Acta Astronautica 158 (mai 2019) : 342–50. http://dx.doi.org/10.1016/j.actaastro.2018.05.056.

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Pereselkov, A., et O. Kruglyakova. « EXPERIMENTAL STUDY OF ELEMENTARY ACTS OF HYDRODYNAMICS AND HEAT TRANSFER DURING THE INTERACTION BETWEEN WATER DROPS AND FILM AND CASTING ROLLER SURFACE ». Integrated Technologies and Energy Saving, n^o 4 (12 décembre 2022) : 3–12. http://dx.doi.org/10.20998/2078-5364.2022.4.01.

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Experimental studies of the boundary conditions of heat transfer for the thermally stressed state of casting rollers while are spraying with flat-jet nozzles in a thermal preconditioning unit have been carried out. It is shown that the hydrodynamic conditions on the sprinkling surface are formed as a result of both the influx of "primary" dispersed water from the flat jet nozzle, and the "secondary" liquid coming from neighboring areas in the form of reflected drops and films. The heat transfer effecting individual factors that form the hydrodynamic conditions on the sprinkling surface was studied separately. The heat transfer intensity was studied depending on the spraying density, the injection-pressure drop and the temperature of the cooled surface when the "primary" drop flow runs in the heat exchange surface. The local sprinkling density of droplets on the surface under the flat-jet nozzle spray were measured using a sampling tube moved by a coordinator. At the same time, the ingress of “secondary” liquid into it was excluded. The specific heat flux and heat transfer coefficient were determined using a heat meter made of a nichrome tape heated by direct current. In this case, the isothermality of the surface of the measuring section was ensured. Thermocouples measured the temperature of the lower surface of the tape, and then the stationary temperature of the upper surface of the heat meter sprinkled with drops is calculated. As a result of the multivariate analysis of the experimental data, the correlation dependence of the heat transfer coefficient in dependance on the local spraying conditions of the heat meter surface was obtained. Also, studies of the heat transfer during water film flow over the heat meter surface were carried out. A similar situation takes place when water spreads between the adjacent nozzles sprinkling zones of the roller surface. The correlation dependence between the heat transfer coefficient, the water film speed and the cooled surface temperature was obtained. Studies of heat transfer during combined influence of moving water film and a flat-jet nozzle drop flow on the heat exchange surface showed that the heat transfer rate is approximately 80–90 % of the arithmetic sum of the coefficients obtained by separate cooling the heat meter with drops and a water film.

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3

Kukutla, Pol Reddy, et B. V. S. S. S. Prasad. « Numerical Study on the Secondary Air Performance of the Film Holes for the Combined Impingement and Film Cooled First Stage of High Pressure Gas Turbine Nozzle Guide Vane ». International Journal of Turbo & ; Jet-Engines 37, n^o 3 (27 août 2020) : 221–40. http://dx.doi.org/10.1515/tjj-2017-0022.

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AbstractThe present numerical investigation of Leading Edge (LE) Nozzle Guide Vane (NGV) is considered with five rows of impingement holes combined with five rows of film cooled for the secondary coolant flow path analysis. The coolant mass flow rate variations in all the LE rows of the film holes externally subjected to the hot main stream were obtained by making a three-dimensional computational analysis of NGV with a staggered array of film cooled rows. The experiments were carried out for the same NGV using Particle Image Velocimetry technique to determine the effused coolant jet exit velocity at the stagnation row of film holes as mentioned in reference [Kukutla PR, Prasad BVSSS. Secondary flow visualization on stagnation row of a combined impingement and film cooled high pressure gas turbine nozzle guide vane using PIV technique, J Visualization, 2017; DOI: 10.1007/s12650-017-0434-6]. In this paper, results are presented for three different mass flow rates ranges from 0.0037 kg/s to 0.0075 kg/s supplied at the Front Impingement Tube (FIT) plenum. And the mainstream velocity 6 m/s was maintained for all the three coolant mass flow rates. The secondary coolant flow distribution was performed from SH1 to SH5 row of film holes. Each row of a showerhead film hole exit coolant mass flow rate varied in proportion to the amount of coolant mass rates supplied at the FIT cooling channel. The corresponding minimum and maximum values and their film hole locations were altered. The same behaviour was continued for the coolant pressure drop and temperature rise from SH1 to SH5 row of film holes. Owing to the interaction between hot main stream and the coolant that effuses out of the film holes, occasional presence of hot gas ingestion was noticed for certain flow rates. This caused nonlinear distribution in mass flow, pressure drop and temperature rise. The minimum flow rate results estimate oxidation of NGV material near the film cooled hole. And the effect of hot gas ingestion on the ejected film cooled jet which would recommends effective oxidation resistant material which in turn leads to better durability of the NGV surface.

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Wang, Ten-See, et Mike Guidos. « Transient Three-Dimensional Side-Load Analysis of a Film-Cooled Nozzle ». Journal of Propulsion and Power 25, n^o 6 (novembre 2009) : 1272–80. http://dx.doi.org/10.2514/1.41025.

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Yang, R. J. « Assessment of turbulence and chemistry models for film-cooled nozzle flows ». Journal of Thermophysics and Heat Transfer 10, n^o 2 (avril 1996) : 284–89. http://dx.doi.org/10.2514/3.785.

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Sellam, Mohamed, et Amer Chpoun. « Numerical Simulation of Reactive Flows in Overexpanded Supersonic Nozzle with Film Cooling ». International Journal of Aerospace Engineering 2015 (2015) : 1–15. http://dx.doi.org/10.1155/2015/252404.

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Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2gas (GH2) in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are also analyzed in terms of cooling efficiency and the results already give an idea on their magnitude. It was also shown that slightly increasing the film injection pressure can avoid the reignition phenomena by moving the separation shock towards the nozzle exit section.

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Wang, Ten-See, Jeff Lin et Mike Guidos. « Transient Side-Load Analysis of Out-of-Round Film-Cooled Nozzle Extensions ». Journal of Propulsion and Power 29, n^o 4 (juillet 2013) : 855–66. http://dx.doi.org/10.2514/1.b34812.

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Kozyulin, N. N., M. S. Bobrov et M. Y. Hrebtov. « Adjoint shape optimization of a duct for a wall jet film cooling setup ». Journal of Physics : Conference Series 2119, n^o 1 (1 décembre 2021) : 012018. http://dx.doi.org/10.1088/1742-6596/2119/1/012018.

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Abstract The paper presents the results of optimization of the geometric parameters of the simplified wall jet cooling system using a modified Adjoint Shape optimization method for algebraic systems of equations (Discrete Adjoint Optimization). The modification consists in using a linearized discrete system of equations with the replacement of derivatives by their finite-volume approximations. The jet flowed through a duct and out from a nozzle. The duct was inclined at an angle of 35 degrees to the cooled wall. The mean velocity ratio between the jet and the main flow was set to 2. The total heat flux on the cooled wall was taken as a cost function. The problem was considered in a two-dimensional stationary turbulent formulation (RANS). As a result of optimization, the shape of the duct changed significantly, affecting the flow inside it. The optimization led to the disappearance of the recirculation zone and reattaching of the jet to the cooled wall. As a result of the optimization performed, the heat flux at the wall increased by 20%.

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Day, C. R. B., M. L. G. Oldfield et G. D. Lock. « The Influence of Film Cooling on the Efficiency of an Annular Nozzle Guide Vane Cascade ». Journal of Turbomachinery 121, n^o 1 (1 janvier 1999) : 145–51. http://dx.doi.org/10.1115/1.2841223.

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This paper examines the effect of aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios under ambient conditions. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility that correctly models engine Reynolds and Mach numbers, as well as the inlet turbulence intensity. The use of foreign gas coolant poses specific challenges not present in an air-cooled cascade, and this paper addresses two. First, a novel method for the determination of mass flow from pneumatic probe data in a heterogeneous gas environment is presented that eliminates the need to measure concentration in order to determine loss. Second, the authors argue on the grounds of dimensionless similarity that momentum flux ratio is to be preferred to blowing rate for the correct parameterization of film cooling studies with varying coolant densities. Experimental results are presented as area traverse maps, from which values for loss have been calculated. It is shown that air and foreign gas at the same momentum flux ratio give very similar results, and that the main difference between cooled and uncooled configurations is an increase in wake width. Interestingly, it is shown that an increase in the momentum flux ratio above the design value with foreign gas coolant reduces the overall loss compared with the design value. The data have been obtained both for purposes of design and for CFD code Validation.

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Reddy Kukutla, Pol, et BVSSS Prasad. « Network analysis of a coolant flow performance for the combined impingement and film cooled first-stage of high pressure gas turbine nozzle guide vane ». Proceedings of the Institution of Mechanical Engineers, Part G : Journal of Aerospace Engineering 233, n^o 6 (16 avril 2018) : 1977–89. http://dx.doi.org/10.1177/0954410018767290.

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The present paper describes a system-level thermo-fluid network analysis for the secondary air system analysis of a typically film-cooled nozzle guide vane with multiple actions of jet impingement. The one-dimensional simulation was done with the help of the commercially available Flownex 2015 software. The system-level thermo-fluid network results were validated with both the computational fluid dynamics results and experimentally available literature. The entire nozzle guide vane geometry was first mapped to a thermo-fluid network model and the pressure conditions at different nodes. The discharge and heat transfer coefficients obtained from the Ansys FLUENT were specified as inputs to the thermo-fluid network model. The results show that the one-dimensional simulation of the coolant mass flow rates and jet Nusselt number values are in good agreement with the three-dimensional computational fluid dynamics results.

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Sarkar, S., K. Das et D. Basu. « Film cooling on a turbine guide vane : A numerical analysis with a multigrid technique ». Proceedings of the Institution of Mechanical Engineers, Part A : Journal of Power and Energy 215, n^o 1 (1 février 2001) : 39–53. http://dx.doi.org/10.1243/0957650011536552.

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The flow and heat transfer due to film cooling over a turbine nozzle guide vane, which was also cooled by internal convection, were numerically analysed under engine conditions. The time-dependent, two-dimensional, mass-averaged, Navier-Stokes (N-S) equations are solved in the physical plane based on the four-stage Runge-Kutta algorithm in the finite volume formulation. Local time stepping, variable coefficient implicit residual smoothing and a full multigrid technique have been implemented to accelerate the steady state calculations. Turbulence was simulated by the algebraic Baldwin-Lomax (B-L) model. The computed heat transfer distributions with film cooling in conjunction was successful in describing the coolant behavior over the curved suction and pressure surfaces of a turbine blade for varying blowing and temperature ratios.

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Metzger, D. E., et R. S. Bunker. « Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions : Part II—Impingement Cooling With Film Coolant Extraction ». Journal of Turbomachinery 112, n^o 3 (1 juillet 1990) : 459–66. http://dx.doi.org/10.1115/1.2927681.

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An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions that allow extraction, or bleed-off, of a portion of the internal cooling flow to provide leading edge film cooling along the blade external surface. The internal impingement air is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and generally exiting, minus the bleed-off flow, in the opposite or chordwise direction. The film coolant flow extraction takes place through two lines of holes, one each on the blade suction side and the blade pressure side, both fairly close to the airfoil leading edge. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surfaces, and are observed during a test transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice is covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise-average Nusselt numbers and spanwise Nusselt number gradients as functions of chordwise position. The results without film coolant extraction, presented in the Part I companion paper, are used as a basis of comparison to determine the additional effects of the film cooling bleed. Results indicate that heat transfer is primarily dependent on jet Reynolds number with smaller influences from the flow extraction rate. The results also suggest that changes in the spanwise alignment of the impingement nozzles relative to the position of the film cooling holes can cause significant variations in leading edge metal temperatures.

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Holgate, Nicholas E., Peter T. Ireland et Eduardo Romero. « An experimental-numerical method for transient infrared measurement of film cooling effectiveness and heat transfer coefficient in a single test ». Aeronautical Journal 123, n^o 1270 (5 août 2019) : 1982–98. http://dx.doi.org/10.1017/aer.2019.26.

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ABSTRACTAn experimental technique for assessing film cooling performance is proposed which can determine both film effectiveness and heat transfer coefficient distributions from a single infrared experiment. First, the film effectiveness is determined in the experiment’s steady-state phase on a series of film-cooled nozzle guide vane leading edge geometries made of a low thermal conductivity foam. Then, the effectiveness is used to calculate the distribution of the transient phase driving gas temperatures, which is applied to a finite element conduction model. Heat transfer coefficients are guessed and iteratively refined until the surface temperature histories predicted by the finite element model match those which were experimentally observed. Unlike conventional methods based on one-dimensional analytical heat transfer solutions, this approach does not require assumptions about the material thickness underlying the test surface or the uniformity with depth of its initial temperature distribution. This relieves certain experimental constraints and reduces uncertainty in results.

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Day, C. R. B., M. L. G. Oldfield et G. D. Lock. « Aerodynamic performance of an annular cascade of film cooled nozzle guide vanes under engine representative conditions ». Experiments in Fluids 29, n^o 2 (7 août 2000) : 117–29. http://dx.doi.org/10.1007/s003489900062.

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15

Liu, Zhi Gang, Xiang Jun Fang, Si Yong Liu, Ping Wang et Zhao Yin. « Research of Aerodynamic Performance of HP-Turbine with Coolant Injections for Variable Cycle Engine ». Applied Mechanics and Materials 110-116 (octobre 2011) : 1047–53. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.1047.

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A highly loaded high-pressure turbine with a supersonic nozzle guide vane and a transonic rotor for a Variable Cycle Engine (VCE) has been investigated. Film cooling strategies were designed for the whole stage, during which the positions, injection orientations and arrangements of cooling holes were confirmed. Three-dimensional steady numerical simulations have been performed in the two operation modes of low and high bypass ratio with different thermodynamic cycle parameters according to the VCE and the coolant injections have been simulated by means of additional source term method. The influences of coolant injections in the fully cooled turbine stage on aerodynamic performance and flow characteristics have been analyzed. The results indicate that, the supersonic nozzle guide vane, over-expansion degree of main flows, fluctuations of static pressure and intensity of corner vortex are lessened or alleviated. In the transonic rotor, expansion and doing work capabilities in the mixed fluid are strengthened. Proper coolants injections are beneficial to the flow characteristics in the blade passage.

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Barigozzi, Giovanna, Giuseppe Franchini et Antonio Perdichizzi. « End-Wall Film Cooling Through Fan-Shaped Holes With Different Area Ratios ». Journal of Turbomachinery 129, n^o 2 (21 juillet 2006) : 212–20. http://dx.doi.org/10.1115/1.2464140.

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The present paper reports on the aerothermal performance of a nozzle vane cascade, with film-cooled end walls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two end-wall geometries with different area ratios have been compared. Tests have been carried out at low speed (M=0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through three-dimensional (3D) aerodynamic measurements, by means of a miniaturized five-hole probe. Adiabatic effectiveness distributions have been determined by using the wide-band thermochromic liquid crystals technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performances of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

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Butorina, Antonina V., Sergei B. Nesterov et Nikolay A. Andreev. « Experimental study of cooling spray for physiotherapeutic treatment ». Russian Journal of Physiotherapy, Balneology and Rehabilitation 19, n^o 1 (23 octobre 2020) : 40–43. http://dx.doi.org/10.17816/1681-3456-2020-19-1-6.

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For cooling damaged areas of biotissue in order to achieve a rapid analgesic effect, a procedure for applying a thin film to the skin, boiling at a temperature of T0 = 273238 K of a gas mixture, is widely used. This temperature is achieved when using a cooling spray. Experimental data on the temperature distribution on the cooled surface when throttling a propane/butane/R123 mixture from a nozzle with a diameter of 0.5 mm to the temperature level Т0 = 240; 263; 270 К, used for physiotherapy purposes, are presented. A comparison of the efficiency of cooling the skin using a cryopresponder and using ice was also made. It is shown that cooling with a cooling spray is more efficient.

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Lushchik, V. G., V. I. Sizov, L. E. Sternin et A. E. Yakubenko. « Specific impulse losses due to friction and dispersion in a gas-film cooled liquid rocket engine nozzle ». Fluid Dynamics 28, n^o 4 (juillet 1993) : 495–503. http://dx.doi.org/10.1007/bf01342684.

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19

Bunker, R. S., et D. E. Metzger. « Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions : Part I—Impingement Cooling Without Film Coolant Extraction ». Journal of Turbomachinery 112, n^o 3 (1 juillet 1990) : 451–58. http://dx.doi.org/10.1115/1.2927680.

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An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions. Cooling is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and exiting the leading edge region in the opposite or chordwise direction. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surface and are observed during the transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice are covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise average Nusselt number and spanwise Nusselt number gradient as functions of chordwise position. Results indicate general increases in heat transfer with approximately the 0.6 power of jet Reynolds number, increases in heat transfer with decreasing leading edge sharpness as well as with decreasing nozzle-to-apex distance, and increases in spanwise average heat transfer with decreasing jet pitch-to-diameter ratio. The latter increases are accompanied by increases in the spanwise gradient of the heat transfer coefficient. Comparison with available prior results of much coarser spatial resolution shows good agreement and establishes confidence in the use of the results for design purposes and as baseline results for comparison with subsequent experiments involving film cooling bleed.

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Harasgama, S. P., et C. D. Burton. « Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade : Part 2—Analysis and Correlation of Results ». Journal of Turbomachinery 114, n^o 4 (1 octobre 1992) : 741–46. http://dx.doi.org/10.1115/1.2928027.

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Results have been presented on the heat transfer characteristics of the film cooled endwall (platform) of a turbine nozzle guide vane in an annular cascade at engine representative conditions in a companion paper by Harasgama and Burton (1992). The present paper reports on the analysis of these measurements. The experimental results are well represented by the superposition theory of film cooling. It is shown that high cooling effectiveness can be achieved when the data are corrected for axial pressure gradients. The data are correlated against both the slot-wall jet parameter and the discrete hole injection function for flat-plate, zero pressure gradient cases. The pressure gradient correction brings the present data to within ± 11 percent of the discrete hole correlation. Preliminary predictions of heat transfer reduction have been carried out using the STANCOOL program. These indicate that the code can predict the magnitude of heat transfer reduction correctly, although the absolute values are not in good agreement. This is attributed to the three-dimensional nature of the flow at the endwall.

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Barigozzi, Giovanna, Giuseppe Benzoni, Giuseppe Franchini et Antonio Perdichizzi. « Fan-Shaped Hole Effects on the Aero-Thermal Performance of a Film-Cooled Endwall ». Journal of Turbomachinery 128, n^o 1 (1 février 2005) : 43–52. http://dx.doi.org/10.1115/1.2098788.

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The present paper investigates the effects of a fan-shaped hole endwall cooling geometry on the aero-thermal performance of a nozzle vane cascade. Two endwall cooling geometries with four rows of holes were tested, for different mass flow rate ratios: the first configuration is made of cylindrical holes, whereas the second one features conical expanded exits and a reduced number of holes. The experimental analysis is mainly focused on the variations of secondary flow phenomena related to different injection rates, as they have a strong relationship with the film cooling effectiveness. Secondary flow assessment was performed through downstream 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. The results show that at high injection rates, the passage vortex and the 3D effects tend to become weaker, leading to a strong reduction of the endwall cross flow and to a more uniform flow in spanwise direction. This is of course obtained at the expense of a significant increase of losses. The thermal behavior was then investigated through the analysis of adiabatic effectiveness distributions on the two endwall configurations. The wide-banded thermochromic liquid crystals (TLC) technique was used to determine the adiabatic wall temperature. Using the measured distributions of film-cooling adiabatic effectiveness, the interaction between the secondary flow vortices and the cooling jets can be followed in good detail all over the endwall surface. Fan-shaped holes have been shown to perform better than cylindrical ones: at low injection rates, the cooling performance is increased only in the front part of the vane passage. A larger improvement of cooling coverage all over the endwall is attained with a larger mass flow rate, about 1.5% of core flow, without a substantial increase of the aerodynamic losses.

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Kukutla, Pol, et B. Prasad. « Coupled flow network model and CFD analysis for a combined impingement and film cooled gas turbine nozzle guide vane ». Modelling, Measurement and Control B 86, n^o 1 (30 mars 2017) : 250–70. http://dx.doi.org/10.18280/mmc_b.860118.

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Jenkins, Sean C., et David G. Bogard. « Scaling of Guide Vane Coolant Profiles and the Reduction of a Simulated Hot Streak ». Journal of Turbomachinery 129, n^o 3 (8 août 2006) : 619–27. http://dx.doi.org/10.1115/1.2447803.

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The turbine section of a gas turbine engine is subjected to a nonuniform temperature distribution in the gas flow from the combustor. Regions of elevated temperatures, known as “hot streaks,” subject the turbine airfoil to high heat loads. In this study, the reduction of hot streaks by coolant from a film cooled nozzle guide vane was experimentally evaluated. Experiments were conducted with an approach mainstream turbulence level of 20% to simulate actual turbine conditions. The coolant distributions downstream of the vane were measured for varying blowing ratios and varying coolant density, and scaling methods were found for variations in both parameters. For this study, the hot streak peak was positioned to impact the vane at the stagnation line. Measurements of the hot streak strength with coolant blowing showed as much as a 55% decrease in peak temperature compared with no coolant.

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Vogel, G. « Visualization of a narrow-band transient liquid crystal signal on a film-cooled contoured platform of a nozzle guide vane for film cooling performances measurements ». Journal of Visualization 6, n^o 2 (juin 2003) : 89. http://dx.doi.org/10.1007/bf03181607.

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Lee, Je Jun, Young Shin Lee, Jae Hoon Kim, Seong Woo Byun, Song Heo Koo et Soon Il Moon. « Thermal Strength Evaluation of the Super Alloy Structure with Various Thermal Insulation Performances by FEM and Stress-Rupture Experiment ». Key Engineering Materials 353-358 (septembre 2007) : 1064–67. http://dx.doi.org/10.4028/www.scientific.net/kem.353-358.1064.

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The combustor chamber, diffuser and nozzle are the main components of the ramjet engine. In this study, the thermal strength of the combustion chamber of the ramjet engine was evaluated. The combustion chamber consists of an Inconel alloy 718 liner and a 17-4Ph stainless steel housing. The liner is rapidly heated to a high temperature. The heated liner is cooled with a film cooling method that forms a cold boundary layer to separate the hot gas from the surface of the liner. The thermo-structural analysis is evaluated the thermal strength of super alloy structure with various thermal insulation performances by finite element method with code MSC/Nastran. The result of the analysis is compared with accelerated stress rupture test. The experiment is performed to get safety design and estimate actually life-time for combustor chamber under high temperature. In general, the work in this paper is helpful to further improve the understanding and evaluation of thermal strength of the super alloy structure with various thermal insulation performances.

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Knost, D. G., et K. A. Thole. « Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First-Stage Vane ». Journal of Turbomachinery 127, n^o 2 (1 avril 2005) : 297–305. http://dx.doi.org/10.1115/1.1811099.

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In gas turbine development, the direction has been toward higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first-stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool, including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

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Horlock, J. H., D. T. Watson et T. V. Jones. « Limitations on Gas Turbine Performance Imposed by Large Turbine Cooling Flows ». Journal of Engineering for Gas Turbines and Power 123, n^o 3 (1 février 2001) : 487–94. http://dx.doi.org/10.1115/1.1373398.

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Calculations of the performance of modern gas turbines usually include allowance for cooling air flow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion. There are two essential effects of turbine blade cooling: (i) the reduction of the gas stagnation temperature at exit from the combustion chamber (entry to the first nozzle row) to a lower stagnation temperature at entry to the first rotor and (ii) a pressure loss resulting from mixing the cooling air with the mainstream. Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled blading with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used. These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine. for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature (below that which would be set by stoichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.

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Kukutla, Pol Reddy, et B. V. S. S. S. Prasad. « Secondary flow visualization on stagnation row of a combined impingement and film cooled high-pressure gas turbine nozzle guide vane using PIV technique ». Journal of Visualization 20, n^o 4 (10 mai 2017) : 817–32. http://dx.doi.org/10.1007/s12650-017-0434-6.

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Guo, S. M., C. C. Lai, T. V. Jones, M. L. G. Oldfield, G. D. Lock et A. J. Rawlinson. « Influence of Surface Roughness on Heat Transfer and Effectiveness for a Fully Film Cooled Nozzle Guide Vane Measured by Wide Band Liquid Crystals and Direct Heat Flux Gages ». Journal of Turbomachinery 122, n^o 4 (1 février 2000) : 709–16. http://dx.doi.org/10.1115/1.1312798.

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The influence of surface roughness on heat transfer coefficient and cooling effectiveness for a fully film cooled three-dimensional nozzle guide vane (NGV) has been measured in a transonic annular cascade using wide band liquid crystal and direct heat flux gages (DHFGs). The liquid crystal methods were used for rough surface measurements and the DHFGs were used for the smooth surfaces. The measurements have been made at engine representative Mach and Reynolds numbers and inlet free-stream turbulence intensity. The aerodynamic and thermodynamic characteristics of the coolant flow have been modeled to represent engine conditions by using a heavy “foreign gas” (30.2 percent SF6 and 69.8 percent Ar by weight). Two cooling geometries (cylindrical and fan-shaped holes) have been tested. The strategies of obtaining accurate heat transfer data using a variety of transient heat transfer measurement techniques under the extreme conditions of transonic flow and high heat transfer coefficient are presented. The surfaces of interest are coated with wide-band thermochromic liquid crystals, which cover the range of NGV surface temperature variation encountered in the test. The liquid crystal has a natural peak-to-peak roughness height of 25 μm creating a transitionally rough surface on the NGV. The time variation of color is processed to give distributions of both heat transfer coefficient and film cooling effectiveness over the NGV surface. The NGV was first instrumented with the DHFGs and smooth surface tests preformed. Subsequently the surface was coated with liquid crystals for the rough surface tests. The DHFGs were then employed as the means of calibrating the liquid crystal layer. The roughness of 25 μm, which is the typical order of roughness for the in-service turbine blades and vanes, increases the heat transfer coefficient by up to 50 percent over the smooth surface level. The film cooling effectiveness is influenced less by the roughness. [S0889-504X(00)00804-7]

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Rowbury, D. A., M. L. G. Oldfield et G. D. Lock. « Large-Scale Testing to Validate the Influence of External Crossflow on the Discharge Coefficients of Film Cooling Holes ». Journal of Turbomachinery 123, n^o 3 (1 février 2000) : 593–600. http://dx.doi.org/10.1115/1.1375171.

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This paper discusses large-scale, low-speed experiments that explain unexpected flow-interaction phenomena witnessed during annular cascade studies into the influence of external crossflow on film cooling hole discharge coefficients. More specifically, the experiments throw light on the crossover phenomenon, where the presence of the external crossflow can, under certain circumstances, increase the discharge coefficient. This is contrary to most situations, where the external flow results in a decrease in discharge coefficient. The large-scale testing reported helps to explain this phenomenon through an increased understanding of the interaction between the emerging coolant jet and the free-stream flow. The crossover phenomenon came to light during an investigation into the influence of external crossflow on the discharge coefficients of nozzle guide vane film cooling holes. These experiments were performed in the Cold Heat Transfer Tunnel (CHTT), an annular blowdown cascade of film cooled vanes that models the three-dimensional external flow patterns found in modern aero-engines. (Rowbury et al., 1997, 1998). The variation in static pressure around the exit of film cooling holes under different flow conditions was investigated in the large-scale tests. The study centered on three holes whose geometries were based on those found in the leading edge region of the CHTT vanes, as the crossover phenomenon was witnessed for these rows during the initial testing. The experiments were carried out in a low-speed wind tunnel, with the tunnel free-stream flow velocity set to match the free-stream Reynolds number (based on the local radius of curvature) and the “coolant” flow velocity set to replicate the engine coolant-to-free-stream momentum flux ratio. It was found that the apparent enhancement of film cooling hole discharge coefficients with external crossflow was caused by a reduction in the static pressure around the hole exit, associated with the local acceleration of the free-stream around the emerging coolant jet. When these measured static pressures (rather than the free-stream static pressure) were used to calculate the discharge coefficient, the crossover effect was absent. The improved understanding of the crossover phenomenon and coolant-to-free-stream interactions that has been gained will be valuable in aiding the formulation of predictive discharge coefficient schemes.

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Michaud, Mathias, Francesco Ornano, Nafiz Chowdhury et Thomas Povey. « Methodology for High-Accuracy Infrared Calibration in Environments with Through-Wall Heat Flux ». Journal of the Global Power and Propulsion Society 4 (1 avril 2020) : 1–13. http://dx.doi.org/10.33737/jgpps/118091.

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This paper describes a new method for accurate in situ infrared (IR) calibration in environments with significant through-wall heat flux and surface temperature non-uniformity. In the context of turbine research environments, conventional approaches for in situ IR calibrations rely on thermocouples embedded in the surface or bonded to the surface using an adhesive layer. A review of the literature points to lack of emphasis on the uncertainty in the calibration arising from the effect of the adhesive substrate and paint on the temperature measured by the thermocouple, namely that under diabatic conditions (i.e. with through-wall heat flux) the measured temperature deviates from the true surface temperature. We present a systematic study of the sensitivity of the thermocouple temperature to installation conditions seen in typical laboratory IR calibration arrangements, and under realistic conditions of through-wall heat flux. A new technique is proposed that improves the calibration accuracy by reducing the difference between the thermocouple measurement and the external wall temperature seen by the infrared camera. The new technique has the additional advantage of reducing the uncertainty associated with selecting an appropriate pixel in the IR image, by providing a region with greater temperature uniformity especially in environments with significant underlying lateral surface temperature variation. The new approach is experimentally demonstrated and compared to more conventional measurement techniques on a heavily film-cooled nozzle guide vane assembly operated at highly engine-representative conditions. The proposed technique is demonstrated to significantly improve the measurement accuracy for IR in situ calibrations in environment with through-wall heat flux and surface temperature non-uniformity.

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Badinand, T., et T. H. Fransson. « Radiative Heat Transfer in Film-Cooled Liquid Rocket Engine Nozzles ». Journal of Thermophysics and Heat Transfer 17, n^o 1 (janvier 2003) : 29–34. http://dx.doi.org/10.2514/2.6748.

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Montomoli, F., M. Massini, H. Yang et J. C. Han. « The benefit of high-conductivity materials in film cooled turbine nozzles ». International Journal of Heat and Fluid Flow 34 (avril 2012) : 107–16. http://dx.doi.org/10.1016/j.ijheatfluidflow.2011.12.005.

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Chatterton, Steven, Paolo Pennacchi, Andrea Vania et Phuoc Vinh Dang. « Cooled Pads for Tilting-Pad Journal Bearings ». Lubricants 7, n^o 10 (17 octobre 2019) : 92. http://dx.doi.org/10.3390/lubricants7100092.

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Tilting-pad journal bearings (TPJBs) are widely installed in rotating machines owing to their high stability, but some drawbacks can be noted, such as higher cost with respect to cylindrical journal bearings and thermal issues. High temperatures in the pads correspond to low oil-film thicknesses and large thermal deformations in the pads. Therefore, the restriction of the maximum temperature of the bearing is a key aspect for oil-film bearings. The temperature reduction is generally obtained by adopting higher oil inlet flowrates or suitable oil nozzles. In this paper, the idea of using cooled pads with internal channels in which an external cooling fluid is circulated will be applied to a TPJB for the first time. The three-dimensional TEHD model of the TPJB, equipped with a cooled pad, will be introduced, and the results of the numerical simulations will be discussed. Several analyses have been performed in order to investigate the influence of cooling conditions, such as the type, flowrate, inlet temperature and number of cooled pads. Two types of pad geometry with different cross-sections of the cooling circuit, namely, circular and six-square multi-channel sections, have been compared to the reference bearing with solid pads. Simple experimental tests were performed by means of a test rig equipped with a cooled pad bearing obtained with the additive manufacturing process, thus showing the effectiveness of the solution and the agreement with the predictions.

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Wang, Ten-See, Jeff Lin, Joe Ruf et Mike Guidos. « Transient Three-Dimensional Side-Load Analysis of Out-of-Round Film-Cooled Nozzles ». Journal of Propulsion and Power 27, n^o 4 (juillet 2011) : 899–907. http://dx.doi.org/10.2514/1.b34082.

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Wang, Ten-See, Jeff Lin, Joe Ruf, Mike Guidos et Gary C. Cheng. « Effect of Coolant Flow Distribution on Transient Side-Load of Film Cooled Nozzles ». Journal of Propulsion and Power 28, n^o 5 (septembre 2012) : 1081–90. http://dx.doi.org/10.2514/1.b34397.

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Fawzy, Hamza, Qun Zheng et Yuting Jiang. « Impingement cooling using different arrangements of conical nozzles in a film cooled blade leading edge ». International Communications in Heat and Mass Transfer 112 (mars 2020) : 104506. http://dx.doi.org/10.1016/j.icheatmasstransfer.2020.104506.

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Pereselkov, Alexander, et Olga Kruglyakova. « Experimental Studies of the Heat Exchange Between the Water Film and the Casting Roller in the Thermal Preconditioning Chamber ». NTU "KhPI" Bulletin : Power and heat engineering processes and equipment, n^o 4 (30 décembre 2021) : 42–46. http://dx.doi.org/10.20998/2078-774x.2021.04.06.

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When the casting roller is cooled or heated in the preconditioning chamber the water is supplied to its surface by flat-jet nozzles. The visual inspection of the model of the casting roller showed that a considerable part of it can be covered with the water film spreading from sprinkling zones. It was established that the heat conductivity in the roller body is considered to be a crucial thermal preparation factor in the conjugate heat-exchange problem for the roller of a large diameter at Bio criterion values exceeding 20. Hence, it is sufficient to provide an essential level of the heat transfer that corresponds to the heat transfer coefficient of 2000 W/(m2∙K) to provide appropriate operating conditions for the thermal preparation of the roller. The conditions are also met in sprinkling zones. Due to this fact this scientific paper studies the heat exchange conditions under the water film that spreads between the adjacent sprinkling zones. A range of changes in the flow rate of the spreading water film was determined experimentally. The conditions of heat exchange between the surface of alpha-calorimeter and the water film were analyzed depending on its flow rate and the heat meter surface temperature. A generalized correlation equation was derived. It was established that the heat exchange intensity in sprinkling zones and under the spreading water film meets technological roller treatment conditions in the preconditioning chambers. The obtained research data can be used for the rational arrangement of the collectors and flat-jet nozzles in casting roller preconditioning chambers to reduce the cold and hot water consumption and cut down operating costs.

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Isakadze, Tamaz, et Givi Gugulashvili. « Experimental Study of Possible use of Flexible Capillary Tubes in Cryomedicine ». Works of Georgian Technical University, n^o 4(526) (26 décembre 2022) : 72–77. http://dx.doi.org/10.36073/1512-0996-2022-4-72-77.

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To cool damaged areas of biological tissue in order to achieve a rapid analgesic effect, the procedure of applying a thin film of a boiling gas mixture to the skin at a temperature of 𝑇଴ = 273 ÷ 238 𝐾 is widely used. This temperature is easily achieved using a cooling aerosol. Experimental data were obtained on the temperature distribution on the cooled surface during throttling from nozzles with a diameter of 0.1; 0.5; 1 mm propane-butane mixture per temperature level 𝑇଴ = 240; 248; 258; 263 𝐾 used for sports medicine. A study was made of flexible capillary tubes with an inner diameter of 120 μm, an outer diameter of 220 μm, and a wall thickness of 50 μm for use in local cryotherapy and cryosurgery.

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Fawzy, Hamza, Qun Zheng, Yuting Jiang, Aqiang Lin et Naseem Ahmad. « Conjugate heat transfer of impingement cooling using conical nozzles with different schemes in a film-cooled blade leading-edge ». Applied Thermal Engineering 177 (août 2020) : 115491. http://dx.doi.org/10.1016/j.applthermaleng.2020.115491.

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Zore, Krishna, Cristhian Aliaga, Shoaib Shah, John Stokes, Laith Zori et Boris Makarov. « Conjugate Heat Transfer Simulations of a Nozzle Flow over a Film-Cooled Plate ». Journal of Thermophysics and Heat Transfer, 18 décembre 2022, 1–20. http://dx.doi.org/10.2514/1.t6595.

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This paper presents Ansys Fluent results for a rectangular exhaust nozzle test case from the Fifth AIAA Propulsion Aerodynamics Workshop. The main objective of the workshop was to assess the aerothermal interaction of a heated air exhaust with a film-cooled plate downstream of a convergent nozzle exit, which is representative of engine exhaust flow applications with film cooling. Two sets of hierarchical computational meshes are employed. The first set of meshes consists of workshop-provided Pointwise® meshes. The second set of meshes is generated using Ansys Fluent meshing. Steady conjugate heat transfer simulations with radiation are performed to account for heat conduction through solid materials as well as the electromagnetic heat produced by the solid surfaces of the plate, plenum, and nozzle. Reynolds-averaged Navier–Stokes simulations using the shear-stress [Formula: see text] turbulence model are conducted for a workshop-defined nozzle condition, set point 42, using three blowing ratios. Computational fluid dynamics results are validated against experimental velocity and temperature measurements recorded above the film-cooled plate.

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42

Abdeh, H., G. Barigozzi, A. Perdichizzi, M. Henze et J. Krueckels. « Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade ». Journal of Turbomachinery 141, n^o 5 (21 janvier 2019). http://dx.doi.org/10.1115/1.4041923.

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In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through four rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20 deg. The aero-thermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20 deg.

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Pujari, Arun Kumar, B. V. S. S. S. Prasad et Nekkanti Sitaram. « Effect of Thermal Conductivity on Nozzle Guide Vane Internal Surface Temperature Distribution ». International Journal of Turbo & ; Jet-Engines, 17 janvier 2018. http://dx.doi.org/10.1515/tjj-2017-0061.

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Abstract Conjugate heat transfer analysis is carried out in a cascade domain for a nozzle guide vane. The nozzle guide vane is internally cooled by jet impingement cooling, and the external surface is cooled by film cooling. A computational study was carried out with three different materials, having conductivity values of 0.0048, 0.2 and 1.1 W/m.K. Distribution of local surface temperature along the leading edge, pressure and suction surface is reported. The leading edge region showed the maximum increase in internal surface temperature as the conductivity increased among the different regions of the vane internal surface. However, the pressure and suction surfaces showed relatively less increase in the surface temperature distribution. In order to validate the computational result, the obtained temperature data were compared with experimentally obtained surface temperature data. The flow phenomena like jet lift-off and self-induced cross-flow affect the local temperature distribution differently in the three materials. For a constant mainstream and coolant flow, the surface temperature gradient is higher for the lower conductivity material, and the gradient decreases as conductivity increases. Hence, a material with higher conductivity is desired in a combined impingement and film cooled nozzle guide vane, to increase the durability of the vane.

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Prenter, Robin, Ali Ameri et Jeffrey P. Bons. « Deposition on a Cooled Nozzle Guide Vane With Nonuniform Inlet Temperatures ». Journal of Turbomachinery 138, n^o 10 (26 avril 2016). http://dx.doi.org/10.1115/1.4032924.

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External deposition on a slot film cooled nozzle guide vane, subjected to nonuniform inlet temperatures, was investigated experimentally and computationally. Experiments were conducted using a four-vane cascade, operating at temperatures up to 1353 K and inlet Mach number of approximately 0.1. Surveys of temperature at the inlet and exit planes were acquired to characterize the form and migration of the hot streak. Film cooling was achieved on one of the vanes using a single spanwise slot. Deposition was produced by injecting sub-bituminous ash particles with a median diameter of 6.48 μm upstream of the vane passage. Several deposition tests were conducted, including a baseline case, a hot streak-only case, and a hot streak and film cooled case. Results indicate that capture efficiency is strongly related to both the inlet temperature profiles and film cooling. Deposit distribution patterns are also affected by changes in vane surface temperatures. A computational model was developed to simulate the external and internal flow, conjugate heat transfer, and deposition. Temperature profiles measured experimentally at the inlet were applied as thermal boundary conditions to the simulation. For deposition modeling, an Eulerian–Lagrangian particle tracking model was utilized to track the ash particles through the flow. An experimentally tuned version of the critical viscosity sticking model was implemented, with predicted deposition rates matching experimental results well. Comparing overall deposition rates to results from previous studies indicates that the combined effect of nonuniform inlet temperatures and film cooling cannot be accurately simulated by simple superposition of the two independent effects; thus, inclusion of both conditions in experiments is necessary for realistic simulation of external deposition.

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Ragab, Kasem Eid, et Lamyaa El-Gabry. « Heat Transfer Analysis of the Surface of a Nozzle Guide Vane in a Transonic Annular Cascade ». Journal of Thermal Science and Engineering Applications 11, n^o 1 (24 octobre 2018). http://dx.doi.org/10.1115/1.4041266.

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Abstract In the current study, a numerical analysis was performed for the heat transfer over the surface of nozzle guide vanes (NGVs) using three-dimensional computational fluid dynamics (CFD) models. The investigation has taken place in two stages: the baseline nonfilm-cooled NGV and the film-cooled NGV. A finite volume based commercial code was used to build and analyze the CFD models. The investigated annular cascade has no heat transfer measurements available; hence in order to validate the CFD models against experimental data, two standalone studies were carried out on the NASA C3X vanes, one on the nonfilm-cooled C3X vane and the other on the film-cooled C3X vane. Different modeling parameters were investigated including turbulence models in order to obtain good agreement with the C3X experimental data; the same parameters were used afterward to model the industrial NGVs.

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46

Bacci, Tommaso, Riccardo Becchi, Alessio Picchi et Bruno Facchini. « Adiabatic Effectiveness on High-Pressure Turbine Nozzle Guide Vanes Under Realistic Swirling Conditions ». Journal of Turbomachinery 141, n^o 1 (5 novembre 2018). http://dx.doi.org/10.1115/1.4041559.

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In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

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Alqefl, Mahmood H., Kedar P. Nawathe, Pingting Chen, Rui Zhu, Yong W. Kim et Terrence W. Simon. « Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 1 : Aerodynamics ». Journal of Turbomachinery 143, n^o 12 (14 juillet 2021). http://dx.doi.org/10.1115/1.4050329.

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Abstract The first-stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics is not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage. The test section involves an engine-representative combustor–turbine interface geometry, combustor coolant flow, and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

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Alqefl, Mahmood H., Kedar P. Nawathe, Pingting Chen, Rui Zhu, Yong W. Kim et Terrence W. Simon. « Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 2 : Thermal Measurements ». Journal of Turbomachinery 143, n^o 12 (14 juillet 2021). http://dx.doi.org/10.1115/1.4051556.

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Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage (NGV). The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing, and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

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Ziefle, Jörg, et Leonhard Kleiser. « Numerical Investigation of a Film-Cooling Flow Structure : Effect of Crossflow Turbulence ». Journal of Turbomachinery 135, n^o 4 (3 juin 2013). http://dx.doi.org/10.1115/1.4023361.

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Numerical simulation results using large-eddy simulation of a flow configuration relevant to the film cooling of turbine blades are presented. The flow configuration and the simulation parameters are chosen according to an experiment from literature, in which a hot turbulent crossflow over a flat plate is cooled by fluid issuing from a large isobaric plenum through a short inclined circular nozzle. Special attention is paid to the flow structure within the jet nozzle and the mixing region, as well as to the effect of the crossflow fluctuations thereon. To this end, the numerical results with the turbulent crossflow are compared to our previous data obtained with a steady mean-turbulent inflow profile. While the flow inside the nozzle is very similar for the two cases, large differences occur in the mixing region, where a much enhanced spreading of the coolant is observed with the turbulent crossflow. Consequently, the good agreement of the film-cooling efficiencies with the experimental data for the turbulent-crossflow case is contrasted by large deviations with the stationary inflow due to the lack of crossflow fluctuations.

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Andrei, Luca, Luca Innocenti, Antonio Andreini, Bruno Facchini et Lorenzo Winchler. « Film Cooling Modeling for Gas Turbine Nozzles and Blades : Validation and Application ». Journal of Turbomachinery 139, n^o 1 (8 septembre 2016). http://dx.doi.org/10.1115/1.4034233.

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The design of modern gas turbines cooling systems cannot be separated from the use of computational fluid dynamics (CFD) and the accurate estimation of the effect of film cooling. Nevertheless, a complete modeling of film cooling holes within the computational domain requires an effort both from the point of view of the mesh creation and from computational time. It is here proposed a new way to model the film cooling (FCM), capable of representing the effect of the coolant at hole exit. This is possible due to the introduction of local source terms near the hole exit in a delimited portion of the domain, avoiding the meshing process of perforations. The goal is to provide a reliable and accurate tool to simulate film-cooled turbine blades and nozzles without having to explicitly mesh the holes. The model was subjected to an intensive validation campaign, composed of two phases. During the first one, FCM results are compared to experimental data and numerical results (obtained with complete cooling holes meshing) on a series of test cases reproducing flat plate cooling configurations for different coolant conditions (in terms of blowing and density ratio). In the second phase, a film-cooled vane test case has been studied, in order to consider a real injection system and flow conditions: FCM predictions are compared to an in-house developed correlative approach and full conjugate heat transfer (CHT) results. Finally, a comparison between FCM predictions and experimental data was performed on an actual nozzle of a GE Oil & Gas heavy-duty gas turbine, in order to prove the feasibility of the procedure. The presented film cooling model (FCM) proved to be a feasible and reliable tool, able to evaluate adiabatic effectiveness, simplifying the design phase avoiding the meshing process of perforations.

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