Thermal protection of rocket nozzle by using film cooling technology - effect of lateral curvature

Guelailia, A.; Khorsi, Azzeddine; Boudjemai, A.; Wang, Jin

doi:10.18280/ijht.360338

Cited by 5 publications

(3 citation statements)

References 15 publications

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“…In order to enhance film cooling effectiveness, new hole shape of film cooling (lidded hole) was proposed in the work of Guelailia et al [15]. The effects of surface radial and longitudinal curvature on the performance of film cooling was investigated by Guelailia et al [16,17]. Kumar et al [18] investigated the influence of cold flow injector of a rocket nozzle divergent section on film cooling effectiveness.…”

Section: Introductionmentioning

confidence: 99%

Effects of Crescent Shaped Block Width and Length on Flow Field and Film Cooling Performance

Guelailia¹,

Khorsi

Ahmed³

et al. 2023

Trends Sci

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Turbine blades and rocket nozzles can be efficiently protected from combustion chamber exhaust hot gases by film cooling technic. A crescent-shaped block placed downstream of the injection hole can significantly improve cooling effectiveness. The main objective of this research work is to investigate the inﬂuence of the crescent-shaped block length and width on film cooling effectiveness. ANSYS CFX was used to conduct analysis of a flat plate configuration with cylindrical holes. Nine configurations of the crescent shaped blocks were considered. For each case, the effect of blowing ratios (0.5 and 1) was investigated. The turbulence model shear stress transport (SST) is used for approximating turbulence. Good agreement was obtained by comparing the analysis results with the experimental data. The result indicated that block width variation has a considerable impact on film cooling performance. However, slight effect of block length on cooling effectiveness was obtained. Comparing all analyzed configurations, the best cooling effectiveness was reached for the model 9 (W = 3d, B = 1.5d). HIGHLIGHTS Rocket nozzles can be efficiently protected from combustion chamber exhaust hot gases by film cooling technic One of the latest techniques to optimize the overall performance of film cooling is adding an obstacle configuration downstream the coolant perforation Film cooling performance is considerably enhanced by placing a crescent shaped block downstream the injection hole The maximum enhancement in overall film cooling effectiveness using the optimal configuration (model 6: W = 2d, B = 1.5d) is about 252 % achieved at a blowing ratio of M = 1 GRAPHICAL ABSTRACT

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Section: Introductionmentioning

confidence: 99%

Effects of Crescent Shaped Block Width and Length on Flow Field and Film Cooling Performance

Guelailia¹,

Khorsi

Ahmed³

et al. 2023

Trends Sci

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show abstract

“…The heat transfer performance of showerhead film cooling as applied to the vane leading edge, and considering representative lean burn combustor swirling outflow was numerically investigated by Zhuang et al 15 Kouchih et al 16 numerically investigated several film cooling configurations including Barshan-dune-shaped shells, forward injection hole and backward injection hole. The influence of wall radial curvature on film cooling efficiency was numerically analyzed by Guelailia et al 17 A numerical study of swirling flow in coolant jets was conducted by ZHU et al 18 whose results demonstrated that film cooling effectiveness could be improved with jet swirl at high blowing ratios, and that swirl strength has significant influence on film cooling performance. The effects of mainstream turbulence intensity and blowing ratio on the endwall film cooling and heat transfer were numerically investigated by Yang et al 19 Results show that by increasing turbulence intensity, the variation of film cooling effectiveness was very limited.…”

Section: Introductionmentioning

confidence: 99%

Effect of lid height and blowing ratio on film cooling effectiveness of a novel lidded hole configuration

Guelailia

Khorsi

Slimane

et al. 2022

Proceedings of the Institution of Mechanical Engineers, Part E:

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Film cooling is one of the promising technologies used for protecting rocket nozzles and turbine blades from combustion chamber hot gases. This paper proposes a novel shape of film cooling injection hole, called lidded hole, that can offer significant enhancement of cooling performance. ANSYS CFX is used to perform 3D numerical simulations of a flat plate with a single row of lidded holes, in which the k–ε model approximates turbulence effects. Four cases are investigated to highlight the influence of the hole's lid height (H/d = 0, 0.25, 0.5, 0.75). The effect of blowing ratios (M = 0.5, 1, 1.5) is also analyzed for each configuration. The numerical results of this study are compared with available experimental data, and, generally, a good agreement is achieved. The results obtained show that the lidded hole configuration reduces the coolant flow separation which improves significantly the film cooling effectiveness. In addition, increasing the blowing ratio leads to an increase in lateral and centerline cooling effectiveness. Comparing all studied cases, the optimum coolant coverage was obtained for the lidded hole configuration with H/d = 0.25 at M = 1.5.

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“…The distinctness between gas turbine blades and rocket nozzles is the wall curvature. Turbine blade walls are curved in the stream-wise direction, and rocket nozzles walls can be curved in both stream-wise and span-wise directions [1]. Gas turbines are used extensively for aircraft propulsion, land-based power generation, and industrial applications.…”

Section: Introductionmentioning

confidence: 99%

Heat Transfer Analysis of Gas Turbine Blade by Varying Number of Cooling Holes and at Suitable Coolant Speeds Using CFD

Basavaraju¹,

Kaleru²,

Chintireddy³

2021

IJHT

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In the present work, heat transfer analysis of a gas turbine blade consisting of 5, 7, and 10 holes, coolant flowing with suitable velocities 40 m/s, 75 m/s, and 110 m/s. From the results of several investigations, the suitable velocity ranges of coolant fluid have been taken. The coolant fluid used in this work is air, as it is suitable for aircraft engines working on open cycle gas turbines, and of course, it is cheaply available, and the blade material used is Inconel 718. Simulations are carried out using Computational Fluid Dynamics (CFD) software, ANSYS FLUENT 14.5. An analysis is being done on how the temperature is varying in blade with different configurations. Temperature distribution in the blade is studied and variation of different parameters like velocity, Nusselt number, heat transfer coefficient is observed. It is found that the blade cooling is maximum in the case of a blade with 10 holes and coolant inlet velocity 110 m/s. The average Nusselt number with the coolant inlet velocity of 40 m/s, 75 m/s, and 110 m/s is around 11, 19, and 21, respectively. The lowest temperature attained by the blade on the coolant inlet surface is 1152 K and the coolant exit surface is 1334 K. These two temperatures are observed when the blade has 10 cooling holes and coolant inlet velocity is 110 m/s.

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Thermal protection of rocket nozzle by using film cooling technology - effect of lateral curvature

Cited by 5 publications

References 15 publications

Effects of Crescent Shaped Block Width and Length on Flow Field and Film Cooling Performance

Effects of Crescent Shaped Block Width and Length on Flow Field and Film Cooling Performance

Effect of lid height and blowing ratio on film cooling effectiveness of a novel lidded hole configuration

Heat Transfer Analysis of Gas Turbine Blade by Varying Number of Cooling Holes and at Suitable Coolant Speeds Using CFD

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