Numerical Study of Film Cooling Scheme on a Blunt-Nosed Body in Hypersonic Flow

2011 ◽  
Vol 3 (4) ◽  
Author(s):  
Sung In Kim ◽  
Ibrahim Hassan
Keyword(s):  
Film Cooling ◽  
Thermal Protection ◽  
Numerical Study ◽  
Aerodynamic Heating ◽  
Convective Heating ◽  
Computational Domain ◽  
Heating Rates ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes ◽  
Cooling Technique

In hypersonic flight, the prediction of aerodynamic heating and the construction of a proper thermal protection system (TPS) are significantly important. In this study, the method of a film cooling technique, which is already the state of the art in cooling of gas turbine engines, is proposed for a fully reusable and active TPS. Effectiveness of the film cooling scheme to reduce convective heating rates for a blunt-nosed spacecraft flying at Mach number 6.56 and 40 deg angle of attack is investigated numerically. The inflow boundary conditions used the standard values at an altitude of 30 km. The computational domain consists of infinite rows of film cooling holes on the bottom of a blunt-nosed slab. Laminar and several turbulent calculations have been performed and compared. The influence of blowing ratios on the film cooling effectiveness is investigated. The results exhibit that the film cooling technique could be an effective method for an active cooling of blunt-nosed bodies in hypersonic flows.

Numerical Study of Film Cooling Scheme on a Blunt-Nosed Body in Hypersonic Flow

2010 ◽  
Author(s):  
Sung In Kim ◽  
Ibrahim Hassan
Keyword(s):  
Film Cooling ◽  
Thermal Protection ◽  
Numerical Study ◽  
Aerodynamic Heating ◽  
Convective Heating ◽  
Computational Domain ◽  
Heating Rates ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Cooling Technique

In hypersonic flights, the prediction of aerodynamic heating and the construction of a proper thermal protection system (TPS) are significantly important. In this study, the method of a film cooling technique, which is already the state of the art in cooling gas turbine engine, is proposed for a fully reusable and active TPS. Effectiveness of the film cooling scheme to reduce convective heating rates for a blunt nosed spacecraft flying at Mach number 6.56 and 40 degree angle of attack is investigated numerically. The inflow boundary conditions used the standard values at an altitude of 30 km. Computational domain consists of infinite rows of film cooling holes on the bottom of a blunt-nosed slab. Laminar and several turbulent calculations have been performed and compared each other. The influence of blowing ratios on the film cooling effectiveness is investigated. The results exhibit that the film cooling technique could be an effective method for an active cooling of blunt-nosed bodies in hypersonic flows.

A Numerical Study on Increasing Film Cooling Effectiveness Through the Use of Sister Holes

2008 ◽  
Author(s):  
Marc J. Ely ◽  
B. A. Jubran
Keyword(s):  
Film Cooling ◽  
Thermal Protection ◽  
Numerical Study ◽  
Computational Domain ◽  
Film Cooling Effectiveness ◽  
Turbine Blade Cooling ◽  
Blade Cooling ◽  
Centre Line ◽  
Flow Mixing ◽  
Primary Focus

Film cooling has been the primary focus of turbine blade cooling research for the past half century. However, as engines become more powerful, more effective non-traditional means of cooling become necessary. The current study branches out into a new scheme for film cooling; sister holes. The geometry of the current work makes use of three cylindrical holes inclined at 35° to the horizontal: one primary injectant hole bound by two sister holes. Numerical simulations were run with blowing ratios varying from M = 0.2 to M = 1.5, using the realizable k-ε turbulence model with near-wall modeling. The results were analyzed for both adiabatic thermal effectiveness as well as vortex production due to flow mixing. In general, sister holes offer significant advantages in thermal protection over their single hole counterparts both laterally and along the centre-line, particularly in regions close to the hole. Simulations showed that the laterally averaged adiabatic thermal effectiveness increased by a factor of 1.35 for M = 0.2 up to a factor of 1.62 for M = 1.5. Similarly, the centre-line effectiveness increased by a factor of 1.22 at M = 0.2 up to a factor of 1.68 at M = 1.5. These benefits are heavily weighted by the near-hole region; however, increases are evident throughout the computational domain. This sister hole technique offers significant advantages with minimal penalties, making it a valuable candidate for future blade cooling applications.

Coolant Gas Injection on a Blunt-Nosed Re-Entry Vehicle

2013 ◽  
Author(s):  
Jeswin Joseph ◽  
S. R. Shine
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Thermal Protection System ◽  
Protection System ◽  
Aerodynamic Heating ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Mach Numbers

Very high thermal loads are expected in re-entry vehicles traveling at hypersonic Mach numbers due to severe aerodynamic heating. In the present study, numerical investigations are carried out to analyze the use of film cooling technology for a fully reusable and active thermal protection system of the re-entry vehicle. Simulations are done to examine the fundamental flow phenomenon and the performance of blunt body film cooling in hypersonic flows. Simulations are conducted for a blunt -nosed spacecraft flying at Mach numbers varying from 4 to 8 and 40 deg angle of attack. Film cooling holes are provided on the bottom of the blunt-nosed body. Standard values at an altitude of 30 km are used as in flow boundary conditions. The dependency of blowing ratios, stream-wise injection angle and inlet Mach number on the film cooling effectiveness are investigated. It is observed that the film cooling effectiveness reduces with increase in coolant injection angle. The film cooling performance is found to be decreasing with increase in Mach number. The results could provide useful inputs for optimization of an active thermal protection system of re-entry vehicles.

Numerical Study on Cooling Characteristics of Double-Row Cylindrical Holes on Upstream Endwall

2019 ◽  
Author(s):  
Ding Luo ◽  
Ruishan Lu ◽  
Jiang Lei ◽  
Lesley M. Wright
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Computational Domain ◽  
Double Row ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Horseshoe Vortices ◽  
Cylindrical Holes ◽  
Endwall Film Cooling

Abstract In this paper film-cooling effectiveness of two rows of cylindrical holes located on the endwall upstream of a vane is investigated numerically. Five different cooling schemes, including three schemes of double rows cylindrical holes (β = −45°, 0°, 45°), and two schemes of Double-jet film cooling (DJFC) holes (β = −45°, 45° and β = 45°, −45°), are arranged on the endwall at four blowing ratios (M = 0.5, 1.0, 1.5, 2.0). Both primary effect (on downstream endwall) and secondary effect (on pressure and suction surfaces) of the endwall film cooling are considered. ICEM is used to mesh the computational domain, and simulation is carried out by ANSYS 14.0. The result shows that cooling jets with compound angles can effectively suppress lifting-off and increase film-cooling effectiveness. In addition, at low blowing ratios, it is difficult for jets no matter what directions to cool the neighborhood of the leading edge and the pressure side due to the effect of horseshoe vortices. However, passage vortices have different effects on the cooling jets with different compound angles which will result in different film coverage on both endwall and airfoil.

scholarly journals A Numerical Study of Anti-Vortex Film Cooling Designs at High Blowing Ratio

2008 ◽  
Author(s):  
James D. Heidmann
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Experimental Studies ◽  
Navier Stokes ◽  
Round Hole ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Lift Off ◽  
Density Ratios ◽  
Film Cooling Holes

A concept for mitigating the adverse effects of jet vorticity and lift-off at high blowing ratios for turbine film cooling flows has been developed and studied at NASA Glenn Research Center. This “anti-vortex” film cooling concept proposes the addition of two branched holes from each primary hole in order to produce a vorticity counter to the detrimental kidney vortices from the main jet. These vortices typically entrain hot freestream gas and are associated with jet separation from the turbine blade surface. The anti-vortex design is unique in that it requires only easily machinable round holes, unlike shaped film cooling holes and other advanced concepts. The anti-vortex film cooling hole concept has been modeled computationally for a single row of 30 degree angled holes on a flat surface using the 3D Navier-Stokes solver Glenn-HT. A modification of the anti-vortex concept whereby the branched holes exit adjacent to the main hole has been studied computationally for blowing ratios of 1.0 and 2.0 and at density ratios of 1.0 and 2.0. This modified concept was selected because it has shown the most promise in recent experimental studies. The computational results show that the modified design improves the film cooling effectiveness relative to the round hole baseline and previous anti-vortex cases, in confirmation of the experimental studies.

Numerical Investigation of Effects of Blockage, Inclination Angle, and Hole-Size on Film Cooling Effectiveness at Concave Surface

2021 ◽  
Vol 143 (2) ◽  
Author(s):  
Fu-qiang Wang ◽  
Jian Pu ◽  
Jian-hua Wang ◽  
Wei-dong Xia
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Numerical Study ◽  
Density Ratio ◽  
Concave Surface ◽  
Blockage Ratio ◽  
Coverage Area ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle

Abstract Film-hole can be often blocked by thermal-barrier coatings (TBCs) spraying, resulting in the variations of aerodynamic and thermal performances of film cooling. In this study, a numerical study of the blockage effect on the film cooling effectiveness of inclined cylindrical-holes was carried out on a concave surface to simulate the airfoil pressure side. Three typical blowing ratios (BRs) of 0.5, 1.0, and 1.5 were chosen at an engine-similar density ratio (DR) of 2.0. Two common inclination angles of 30 deg and 45 deg were designed. The blockage ratios were adjusted from 0 to 20%. The results indicated the blockage could enhance the penetration of film cooling flow to the mainstream. Thus, the averaged effectiveness and coolant coverage area were reduced. Moreover, the pressure loss inside of the hole was increased. With the increase of BR, the decrement of film cooling effectiveness caused by blockage rapidly increased. At BR = 1.5, the decrement could be acquired up to 70% for a blockage ratio of 20%. The decrement of film cooling effectiveness caused by blockage was nearly nonsensitive to the injection angle; however, the larger angle could generate the higher increment of pressure loss caused by blockage. A new design method for the couple scheme of film cooling and TBC was proposed, i.e., increasing the inlet diameter according to the blockage ratio before TBC spraying. In comparison with the original unblocked-hole, the enlarged blocked-hole not only kept the nearly same area-averaged effectiveness but also reduced slightly the pressure loss inside of the hole. Unfortunately, application of enlarged blocked-hole at large BR could lead to a more obvious reduction of effectiveness near hole-exit, in comparison with the original common-hole.

A Parametric Numerical Study of the Effects of Freestream Turbulence Intensity and Length Scale on Anti-Vortex Film Cooling Design at High Blowing Ratio

2013 ◽  
Author(s):  
Timothy W. Repko ◽  
Andrew C. Nix ◽  
James D. Heidmann
Keyword(s):  
Turbulence Intensity ◽  
Film Cooling ◽  
Numerical Study ◽  
Length Scale ◽  
Freestream Turbulence ◽  
Cooling Effectiveness ◽  
Length Scales ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cooling Design

An advanced, high-effectiveness film-cooling design, the anti-vortex hole (AVH) has been investigated by several research groups and shown to mitigate or counter the vorticity generated by conventional holes and increase film effectiveness at high blowing ratios and low freestream turbulence levels. [1, 2] The effects of increased turbulence on the AVH geometry were previously investigated and presented by researchers at West Virginia University (WVU), in collaboration with NASA, in a preliminary CFD study [3] on the film effectiveness and net heat flux reduction (NHFR) at high blowing ratio and elevated freestream turbulence levels for the adjacent AVH. The current paper presents the results of an extended numerical parametric study, which attempts to separate the effects of turbulence intensity and length-scale on film cooling effectiveness of the AVH. In the extended study, higher freestream turbulence intensity and larger scale cases were investigated with turbulence intensities of 5, 10 and 20% and length scales based on cooling hole diameter of Λx/dm = 1, 3 and 6. Increasing turbulence intensity was shown to increase the centerline, span-averaged and area-averaged adiabatic film cooling effectiveness. Larger turbulent length scales were shown to have little to no effect on the centerline, span-averaged and area-averaged adiabatic film-cooling effectiveness at lower turbulence levels, but slightly increased effect at the highest turbulence levels investigated.

Numerical Study on Flow and Heat Transfer of a Cooled First Stage Vane of a Gas Turbine

2016 ◽  
Author(s):  
Lv Ye ◽  
Zhao Liu ◽  
Xiangyu Wang ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Coolant Flow ◽  
The Other ◽  
Edge Region ◽  
Flow Rates ◽  
Flow And Heat Transfer ◽  
Film Cooling Holes

This paper presents a numerical simulation of composite cooling on a first stage vane of a gas turbine, in which gas by fixed composition mixture is adopted. To investigate the flow and heat transfer characteristics, two internal chambers which contain multiple arrays of impingement holes are arranged in the vane, several arrays of pin-fins are arranged in the trailing edge region, and a few arrays of film cooling holes are arranged on the vane surfaces to form the cooling film. The coolant enters through the shroud inlet, and then divided into two parts. One part is transferred into the chamber in the leading edge region, and then after impinging on the target surfaces, it proceeds further to go through the film cooling holes distributed on the vane surface, while the other part enters into the second chamber immediately and then exits to the mainstream in two ways to effectively cool the other sections of the vane. In this study, five different coolant flow rates and six different inlet pressure ratios were investigated. All the cases were performed with the same domain grids and same boundary conditions. It can be concluded that for the internal surfaces, the heat transfer coefficient changes gradually with the coolant flow rate and the inlet total pressure ratio, while for the external surfaces, the average cooling effectiveness increases with the increase of coolant mass flow rates while decreases with the increase of the inlet stagnation pressure ratios within the study range.

Enhancement of Full Coverage Film Cooling Effectiveness with Mixed Injection Holes

2018 ◽  
Vol 0 (0) ◽  
Author(s):  
Mukesh Prakash Mishra ◽  
A K Sahani ◽  
Sunil Chandel ◽  
R K Mishra
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Velocity Ratio ◽  
Perforated Plate ◽  
Combustion Chambers ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Full Coverage ◽  
Upstream Direction ◽  
The Difference

Abstract In the present work numerical study of full coverage film cooling on an adiabatic flat plate is carried out. Cooling performance of three configurations of cylindrical holes is studied with downstream injection, upstream injection and mixed injection. In mixed injection configuration one column of holes inject in downstream direction and the holes in the adjacent column inject in the upstream direction. Numerical simulations are carried out at different velocity ratios and circumferentially averaged value of adiabatic film cooling effectiveness is estimated. Simulation results indicate that the mixed injection configuration has better and more uniform cooling, throughout the perforated plate, than with downstream injection. The difference is greater with increase in the velocity ratio. Configuration with upstream injection gives better cooling than mixed injection at front few rows of cooling holes but it shows poorer performance with downstream injection in the downstream rows of cooling holes. The obtained results from this study can be an invaluable input for highly loaded combustion chambers.

Numerical Study on the Influence of Trench Width on Film Cooling Characteristics of Double-Wave Trench

2017 ◽  
Author(s):  
Bo-lun Zhang ◽  
Li Zhang ◽  
Hui-ren Zhu ◽  
Jian-sheng Wei ◽  
Zhong-yi Fu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Trench Width

Film cooling performance of the double-wave trench was numerically studied to improve the film cooling characteristics. Double-wave trench was formed by changing the leading edge and trailing edge of transverse trench into cosine wave. The film cooling characteristics of transverse trench and double-wave trench were numerically studied using Reynolds Averaged Navier Stokes (RANS) simulations with realizable k-ε turbulence model and enhanced wall treatment. The film cooling effectiveness and heat transfer coefficient of double-wave trench at different trench width (W = 0.8D, 1.4D, 2.1D) conditions are investigated, and the distribution of temperature field and flow field were analyzed. The results show that double-wave trench effectively improves the film cooling effectiveness and the uniformity of jet at the downstream wall of the trench. The span-wise averaged film cooling effectiveness of the double-wave trench model increases 20–63% comparing with that of the transverse trench at high blowing ratio. The anti-counter-rotating vortices which can press the film on near-wall are formed at the downstream wall of the double-wave trench. With the double-wave trench width decreasing, the film cooling effectiveness gradually reduces at the hole center-line region of the downstream trench. With the increase of the blowing ratio, the span-wise averaged heat transfer coefficient increases. The span-wise averaged heat transfer coefficient of the double-wave trench with 0.8D and 2.1D trench width is higher than that of the double-wave trench with 1.4D trench width at the high blowing ratio conditions.

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