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Research Papers: Forced Convection

Thermal Response of Supersonic Airflow to a Fin Protrusion Situated on a Curved Surface

[+] Author and Article Information
Majid Molki

Department of Mechanical Engineering, Southern Illinois University Edwardsville, Edwardsville, IL 62025mmolki@siue.edu

David W. Fasig

Department of Mechanical Engineering, Southern Illinois University Edwardsville, Edwardsville, IL 62025

J. Heat Transfer 131(11), 111704 (Aug 26, 2009) (5 pages) doi:10.1115/1.3155003 History: Received January 22, 2009; Revised May 07, 2009; Published August 26, 2009

Aerodynamic heating of an airfoil with a short fin attached to its surface is computationally investigated. This research is motivated by the fact that the gap fillers inserted between the insulation tiles of the space shuttle thermal protection system may sometimes get loose and extend beyond the surface and cause an uneven aerodynamic heating of the surface. It is often difficult for engineers to determine whether the protruded gap filler would cause an undesirable effect in the boundary layer including early turbulence transition or shockwaves that could cause an unsafe increase in surface temperature. In this investigation, the supersonic flow over NACA 0012 airfoil on which a short fin is attached is studied using a computational approach. The method is validated by the experimental data available in published literature. The results indicate a significant increase in the surface temperature in the vicinity of the fin. This elevated temperature extends downstream beyond the location of the fin and covers a large portion of the airfoil downstream of the fin. The fin induces an oblique shockwave followed by an expansion wave.

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Copyright © 2009 by American Society of Mechanical Engineers
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References

Figures

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Figure 2

Grid refinement study and validation of the computational results for NACA 0012 airfoil. The experimental data are taken from Fig. 1 of Ref. 9.

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Figure 3

Contours of Mach number for supersonic flow over NACA 0012 airfoil at M=2.5. Flow is from left to right. The fin is located at x/c=0.15, 0.30, and 0.80.

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Figure 4

Contours of pressure for supersonic flow over NACA 0012 airfoil at M=2.5. Flow is from left to right. The fin is located at x/c=0.15, 0.30, and 0.80.

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Figure 1

The NACA 0012 airfoil with a fin protrusion

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Figure 9

Distribution of the temperature recovery factor on the upper and lower surfaces of the airfoil as a function of Prandtl number; NACA 0012; M=2.5. The range of experimental data is from Table 2 of Ref. 3.

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Figure 5

Distribution of pressure coefficient on the upper and lower surfaces of the airfoil; NACA 0012; M=2.5

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Figure 6

Distribution of skin friction coefficient on the upper and lower surfaces of the airfoil; NACA 0012; M=2.5

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Figure 7

Contours of static temperature for supersonic flow over NACA 0012 airfoil at M=2.5. Flow is from left to right. The fin is located at x/c=0.15, 0.30, and 0.80.

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Figure 8

Distribution of the temperature recovery factor on the upper and lower surfaces of the airfoil; NACA 0012; M=2.5. The experimental data are from Fig. 9 of Ref. 3.

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