Download or read book Film Cooling And Aerodynamic Loss In A Gas Turbine Cascade ANNUAL REPORT written by Sadasuke Ito and published by . This book was released on 1976 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Film Cooling and Aerodynamic Loss in a Gas Turbine Cascade written by Sadasuke Ito and published by . This book was released on 1991 with total page 392 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Aerodynamic Loss Models and Multi dimensional Prediction of Discrete Hole Film Cooling on Gas Turbine Blades written by İlker Topçuoğlu and published by . This book was released on 2013 with total page 145 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Film Cooling and End Wall Heat Transfer in Small Turbine Blade Passages written by and published by . This book was released on 1978 with total page 56 pages. Available in PDF, EPUB and Kindle. Book excerpt: Two topics have been studied related to the cooling of the end wall of a turbine passage. The first concerns the development of a method for measuring the adiabatic wall effectiveness and heat transfer coefficient of a film cooling system for protecting a surface from high heating derived from a hot compressible flow. The second concerns the measurement of the heat transfer rate distribution to a turbine cascade end wall in order to choose an appropriate film cooling system. These are related to providing the background to the final phase of the study in which the effectiveness of a film cooling system to cool a turbine end wall will be made combined with the measurement of the aerodynamic losses incurred by such a system. (Author).
Download or read book Computational Film Cooling Methods for Gas Turbine Airfoils written by and published by . This book was released on 2002 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt: A previously documented CFD methodology is further generalized and applied to realistic turbine airfoil film cooling test cases. First, a series of fundamental test cases are examined in order to document the ability of the robust and practical CFD methodology to deal with the effects of the individually isolated key physics mechanisms on film cooling. These include: (1) favorable and adverse pressure gradients; (2) convex curvature; (3) horseshoe vortex; (4) profile losses with and without coolant jets; (5) laminar-to-turbulent boundary layer transition; and (6) discontinuities. Second, in addition to the fundamental flow test cases, two turbine airfoil cascade, one low subsonic and another transonic, configurations were modeled. The subsonic cascade case was designed to study the aerodynamics losses with and without film cooling jets, as well as, the adiabatic effectiveness for a range of parameters. The transonic turbine cascade represents a truly modern design at realistic engine conditions. A total of 18 test configurations, corresponding to compound-round, axial-shaped, and compound-shaped film holes, are simulated in order to document in detail the current state of the readily available robust and practical CFD technology for use by the gas turbine design community.
Download or read book Scientific and Technical Aerospace Reports written by and published by . This book was released on 1995 with total page 488 pages. Available in PDF, EPUB and Kindle. Book excerpt: Lists citations with abstracts for aerospace related reports obtained from world wide sources and announces documents that have recently been entered into the NASA Scientific and Technical Information Database.
Download or read book Impact of film cooling jets on turbine aerodynamic losses ASME 99 GT 421 written by Dibbon K. Walters and published by . This book was released on 1999 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt: Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Indianapolis, Indiana, Jun 7-Jun 10, 1999.
Download or read book Aerodynamic Loss Reduction in a Vane Cascade with Leading Edge Fillet and Upstream Endwall Film Cooling written by Keenesh Arnachellan and published by . This book was released on 2017 with total page 228 pages. Available in PDF, EPUB and Kindle. Book excerpt: Secondary flow structures account for nearly 50% of aerodynamic losses experienced in the turbine blade passages. The adverse effects of these vortex structures transport the hot mainstream fluid towards the endwall blade surfaces, which enhances thermal stresses and leads to blade failure. The effects of leading-edge fillets and film-cooling with flush slots located upstream near the leading-edge region were investigated experimentally in the study in a large-scale linear vane cascade in which the aerodynamic flow field was considered. The introduction of slot film flow and fillet aimed to reduce the effects of the secondary flow structures from the leading edge through the passage towards the exit in an effort to decrease the pressure losses, improve film-cooling coverage and flow field uniformity for the next blade row. The two-dimensional vane profile was obtained from the hub-side airfoil of the GE-E3 engine nozzle guide vane. The slots were configured for two experimental cases to evaluate the influence of coolant flow rate and momentum; first, the effects of slot film injection from all four slots were observed and then compared with the second case injecting coolant only through the two central slots. Further effects were investigated by combining slot film-cooling with the leading-edge fillets employed on the endwall blade junction. The flow field measurements were quantified with spatial distributions of axial vorticity, total pressure loss, endwall static pressure and flow angle deviations taken across the cascade passage. The measurements were obtained at a Reynolds number of 2.0E+05 based on the cascade inlet velocity and vane chord length. Film-cooling inlet blowing ratios between 1.1 and 2.3 were investigated with the supply of coolant provided by a secondary channel. Film-cooling results were compared with the baseline case without slot film flow and fillet. The results indicated substantial improvement in the passage and exit planes with high inlet blowing ratios. The introduction of high momentum coolant flow from the central slots was seen to create laterally reversed axial vorticity, thereby counteracting the cross-flow tendency in the passage. The effects at the passage exit showed suppressed vortex structures with slot film injection from the two central slots only, with further improvements in the flow angle deviations. The leading-edge slots were seen to contribute positive axial vorticity, which enhanced the passage vortex that was pushed away from the endwall at the exit. When the fillet was introduced, it had favourable effects in reducing the pitchwise pressure gradients along the endwall. Filleted film-cooling then resulted in a faint passage vortex system (50-80% size and 20-50% strength reduction) with a restored endwall boundary layer at high film flow rates. The leading-edge fillet was highly effective at the inlet of the blade passage because it weakened the horseshoe vortex formation. Thus, upstream slot film-cooling has great potential to decrease the aerodynamic losses and is further compounded with the leading-edge fillet.
Download or read book Numerical Study of Film Cooling Influence on Performance of Transonic Vane Cascade written by Ahmad Mahmoud Alameldin and published by . This book was released on 2014 with total page 144 pages. Available in PDF, EPUB and Kindle. Book excerpt: Abstract: Gas turbines are a major contributor to world power generation with applications ranging from electricity production to aircrafts propulsion. Their efficiency is subject to continuous research. A gas turbine's overall efficiency is directly proportional to flow inlet temperature. Various methods are implemented to protect hot gas path components from mainstream flow well above their melting temperature, namely, heat resistant coatings, internal cooling and film cooling. The latter is the subject of this work. A 3-D Computational Fluid Dynamics (CFD) model is solved using ANSYS CFX software and compared to experimental measurements of film cooled transonic vane cascade operating at a Mach number of 0.89; the experimental data used for validation is provided by Heat and Power Technology Department of the Royal Institute of Technology (Kungliga Tekniska Hogskolan, KTH) of Stockholm, Sweden. A new approach was used to model the film cooling holes, omitting the need to model both the coolant plenum and cooling tubes, resulting in 180% reduction in grid size and attributed computational cost interpreted in 300% saving in computation time. The new approach was validated on a basic flow problem (flat plate film cooling) and was found to give good agreement with experimental measurements of velocity and temperature at a blowing ratio (BR) of 1 and 2; the experimental data for the flat plate was provided by NASA's Glenn Research Center. The numerical simulation of the cooled vane cascade was compared to experimental measurements for different cooling configurations and different BRs. a) One row on pressure side at BR = 0.8, 0.96 and 2.5. b) Two rows on suction side (location 1) at BR = 0.8, 1.4 and 2.5. c) Two rows on suction side (location 2) at BR = 0.8. And d) Showerhead cooled vane at BR ranges between 1.98 and 5.84. The coolant was applied at the same temperature as the mainstream, to match experimental conditions. A good agreement with the experimental measurements was obtained for exit flow angle, vorticity downstream of the vane, pressure coefficients and aerodynamic loss. The proposed approach of coolant injection modeling is shown to yield reliable results, within the uncertainty of the measurements in most cases. Along with lower computational cost compared to conventional film cooling modeling approach, the new approach is recommended for further analysis for aero and thermal vane cascade flows.
Download or read book Annual Report written by United States. National Advisory Committee for Aeronautics and published by . This book was released on 1953 with total page 534 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Coup d il impartial written by and published by . This book was released on 1787 with total page 12 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Interaction Between Secondary Flow Film Cooling Jets of a Realistic Annular Airfoil Cascade high Mach Number written by Cuong Quoc Nguyen and published by . This book was released on 2010 with total page 149 pages. Available in PDF, EPUB and Kindle. Book excerpt: Film cooling is investigated on a flat plate both numerically and experimentally. Conical shaped film hole are investigated extensively and contribute to the current literature data, which is extremely rare in the open public domain. Both configuration of the cylindrical film holes, with and without a trench, are investigated in detail. Design of experiment technique was performed to find an optimum combination of both geometrical and fluid parameters to achieve the best film cooling performance. From this part of the study, it shows that film cooling performance can be enhanced up to 250% with the trenched film cooling versus non-trenched case provided the same amount of coolant. Since most of the relevant open literature is about film cooling on flat plate endwall cascade with linear extrusion airfoil, the purpose of the second part of this study is to examine the interaction of the secondary flow inside a 3D cascade and the injected film cooling jets. This is employed on the first stage of the aircraft gas turbine engine to protect the curvilinear (annular) endwall platform. The current study investigates the interaction between injected film jets and the secondary flow both experimentally and numerically at high Mach number (M=0.7). Validation shows good agreement between obtained data with the open literature. In general, it can be concluded that with an appropriate film coolant to mainstream blowing ratio, one can not only achieve the best film cooling effectiveness (FCE or [eta]) on the downstream endwall but also maintain almost the same aerodynamic loss as in the un-cooled baseline case. Film performance acts nonlinearly with respect to blowing ratios as with film cooling on flat plate, in the other hand, with a right blowing ratio, film cooling performance is not affect much by secondary flow. In turn, film cooling jets do not increase pressure loss at the downstream wake area of the blades.
Download or read book Film Cooling of the Gas Turbine Cascade Endwall by Discrete Hole Injection written by Kenneth Charles Marston and published by . This book was released on 1991 with total page 160 pages. Available in PDF, EPUB and Kindle. Book excerpt:
Download or read book Transonic aerodynamic losses due to turbine airfoil suction surface film cooling ASME 99 GT 260 written by D. J. Jackson and published by . This book was released on 1999 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt: Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Indianapolis, Indiana, June 7-June 10, 1999.
Download or read book Film Cooling Heat Transfer and Aerodynamic Measurements in a Three Stage Research Gas Turbine written by Arun Suryanarayanan and published by . This book was released on 2010 with total page pages. Available in PDF, EPUB and Kindle. Book excerpt: The existing 3-stage turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A and M University, is re-designed and newly installed to enable coolant gas injection on the first stage rotor platform to study the effects of rotation on film cooling and heat transfer. Pressure and temperature sensitive paint techniques are used to measure film cooling effectiveness and heat transfer on the rotor platform respectively. Experiments are conducted at three turbine rotational speeds namely, 2400rpm, 2550rpm and 3000rpm. Interstage aerodynamic measurements with miniature five hole probes are also acquired at these speeds. The aerodynamic data characterizes the flow along the first stage rotor exit, second stage stator exit and second stage rotor exit. For each rotor speed, film cooling effectiveness is determined on the first stage rotor platform for upstream stator-rotor gap ejection, downstream discrete hole ejection and a combination of upstream gap and downstream hole ejection. Upstream coolant ejection experiments are conducted for coolant to mainstream mass flow ratios of MFR=0.5%, 1.0%, 1.5% and 2.0% and downstream discrete hole injection tests corresponding to average hole blowing ratios of M = 0.5, 0.75, 1.0, 1.25, 1.5, 1.75 and 2.0 for each turbine speed. To provide a complete picture of hub cooling under rotating conditions, experiments with simultaneous injection of coolant gas through upstream and downstream injection are conducted for an of MFR=1% and Mholes=0.75, 1.0 and 1.25 for the three turbine speeds. Heat transfer coefficients are determined on the rotor platform for similar upstream and downstream coolant injection. Rotation is found to significantly affect the distribution of coolant on the platform. The measured effectiveness magnitudes are lower than that obtained with numerical simulations. Coolant streams from both upstream and downstream injection orient themselves towards the blade suction side. Passage vortex cuts-off the coolant film for the lower MFR for upstream injection. As the MFR increases, the passage vortex effects are diminished. Effectiveness was maximum when Mholes was closer to one as the coolant ejection velocity is approximately equal to the mainstream relative velocity for this blowing ratio. Heat transfer coefficient and film cooling effectiveness increase with increasing rotational speed for upstream rotor stator gap injection while for downstream hole injection the maximum effectiveness and heat transfer coefficients occur at the reference speed of 2550rpm.
Download or read book Sweeping Jet Film Cooling written by Mohammad Arif Hossain and published by . This book was released on 2020 with total page 242 pages. Available in PDF, EPUB and Kindle. Book excerpt: Gas turbine is an integrated part of modern aviation and power generation industry. The thermal efficiency of a gas turbine strongly depends on the turbine inlet temperature (TIT), and the turbine designers are continuously pushing the TIT to a higher value. Due to the increased freedom in additive manufacturing, the complex internal and external geometries of the turbine blade can be leveraged to utilize innovative cooling designs to address some of the shortcomings of current cooling technologies. The sweeping jet film cooling has shown some promise to be an effective method of cooling where the coolant can be brought very close to the blade surface due to its sweeping nature. A series of experiments were performed using a row of fluidic oscillators on a flat plate. Adiabatic cooling effectiveness, convective heat transfer coefficient, thermal field, and discharge coefficient were measured over a range of blowing ratios and freestream turbulence. Results were compared with a conventional shaped hole (777-hole), and the sweeping jet hole shows improved cooling performance in the lateral direction. Numerical simulation also confirmed that the sweeping jet creates two alternating vortices that do not have mutual interaction in time. When the jet sweeps to one side of the hole exit, it acts as a vortex generator as it interacts with the mainstream ow. This prevents the formation of the counter-rotating vortex pair (CRVP) and allows the coolant to spread in the lateral direction. The results obtained from the low speed at plate tests were utilized to design the sweeping jet film cooling hole for more representative turbine vane geometry. Experiments were performed in a low-speed linear cascade facility. Results showed that the sweeping jet hole has higher cooling effectiveness in the near hole region compared to the shaped hole at high blowing ratios. Next, a detailed experimental investigation of sweeping jet film cooling on the suction surface of a near engine scale transonic nozzle guide vane at an engine relevant Mach number (Ma = 0.8) and Reynolds number (Re = 1x10e6) to determine the effect of compressibility. The heat transfer measurements were conducted with a transient IR method, and the convective heat transfer coefficient (HTC) and adiabatic film cooling effectiveness were estimated using a dual linear regression technique (DLRT). Aerodynamic loss measurements were also performed at an exit plane downstream of the vane cascade. Finally, a comprehensive design integration of sweeping jet film hole was carried out in a Direct Metal Laser Sintering (DMLS) enabled engine scale nozzle guide vane and experimental investigation of overall cooling effectiveness at engine relevant temperature conditions were assessed. The systematic evolution of a sweeping jet film cooling hole design from a large scale flat plate to an engine scale nozzle guide vane has been presented.
Download or read book A systematic experimental study on the aerodynamics of leading edge film cooling on a large scale high pressure turbine cascade ASME 98 GT 434 written by Sabine Ardey and published by . This book was released on 1998 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt: Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Stockholm, Sweden, Jun 2-5, 1998.
