Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing
Euler and Navier-Stokes results are presented for a blunt delta wing at Mach 7.15 and 300 angle of attack. The viscous calculations were done at a Reynolds number based on chord of 5.85 x 106 with freestream and wall temperatures set to 74K and 288K respectively. The inviscid simulations were ca...
| Main Author: | |
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| Format: | Technical Report |
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Aerospace Computational Design Lab, Dept. of Aeronautics & Astronautics, Massachusetts Institute of Technology
2012
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| Online Access: | http://hdl.handle.net/1721.1/70568 |
| _version_ | 1826188604432449536 |
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| author | Lee, Kuok Ming |
| author_facet | Lee, Kuok Ming |
| author_sort | Lee, Kuok Ming |
| collection | MIT |
| description | Euler and Navier-Stokes results are presented for a blunt delta wing at Mach
7.15 and 300 angle of attack. The viscous calculations were done at a Reynolds
number based on chord of 5.85 x 106 with freestream and wall temperatures set to
74K and 288K respectively.
The inviscid simulations were carried out using a finite volume, central difference
code written by Roberts [21] and Goodsell [7]. The Navier-Stokes results were
obtained on the semi-implicit extension of the inviscid code, developed by Loyd
[17].
The inviscid results showed a strong shock on the windward side of the wing
at a stand-off angle of about -5' from the body. As the flow traverses around the
leading edge it accelerates strongly through an expansion fan. On the upper surface
of the wing, separation occurs at about 60% span resulting in a region of reverse
cross stream flow.
The viscous calculations display a similar shock structure. Furthermore the
boundary layer on the windward side is thin and variations in the circumferential
direction are small. The flow on the leeward side of the wing separates in 2 places.
The primary separation occurs just inside of the leading edge, and the secondary
separation region is located further inboard.
The inviscid CL and CD are 0.547 and 0.383 respectively, whereas the viscous
values are 0.547 and 0.386. The viscous component contributes only an insignificant
2.32 x 10- S to the CD of the Navier-Stokes calculations. |
| first_indexed | 2024-09-23T08:02:19Z |
| format | Technical Report |
| id | mit-1721.1/70568 |
| institution | Massachusetts Institute of Technology |
| last_indexed | 2024-09-23T08:02:19Z |
| publishDate | 2012 |
| publisher | Aerospace Computational Design Lab, Dept. of Aeronautics & Astronautics, Massachusetts Institute of Technology |
| record_format | dspace |
| spelling | mit-1721.1/705682019-04-09T16:14:16Z Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing Lee, Kuok Ming Euler and Navier-Stokes results are presented for a blunt delta wing at Mach 7.15 and 300 angle of attack. The viscous calculations were done at a Reynolds number based on chord of 5.85 x 106 with freestream and wall temperatures set to 74K and 288K respectively. The inviscid simulations were carried out using a finite volume, central difference code written by Roberts [21] and Goodsell [7]. The Navier-Stokes results were obtained on the semi-implicit extension of the inviscid code, developed by Loyd [17]. The inviscid results showed a strong shock on the windward side of the wing at a stand-off angle of about -5' from the body. As the flow traverses around the leading edge it accelerates strongly through an expansion fan. On the upper surface of the wing, separation occurs at about 60% span resulting in a region of reverse cross stream flow. The viscous calculations display a similar shock structure. Furthermore the boundary layer on the windward side is thin and variations in the circumferential direction are small. The flow on the leeward side of the wing separates in 2 places. The primary separation occurs just inside of the leading edge, and the secondary separation region is located further inboard. The inviscid CL and CD are 0.547 and 0.383 respectively, whereas the viscous values are 0.547 and 0.386. The viscous component contributes only an insignificant 2.32 x 10- S to the CD of the Navier-Stokes calculations. 2012-05-10T18:12:32Z 2012-05-10T18:12:32Z 1989-10 Technical Report http://hdl.handle.net/1721.1/70568 ACDL Technical Reports;CFDL-TR-89-9 An error occurred on the license name. An error occurred getting the license - uri. application/pdf Aerospace Computational Design Lab, Dept. of Aeronautics & Astronautics, Massachusetts Institute of Technology |
| spellingShingle | Lee, Kuok Ming Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title | Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title_full | Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title_fullStr | Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title_full_unstemmed | Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title_short | Numerical Simulation of Hypersonic Flow Over a Blunt Leading Edge Delta Wing |
| title_sort | numerical simulation of hypersonic flow over a blunt leading edge delta wing |
| url | http://hdl.handle.net/1721.1/70568 |
| work_keys_str_mv | AT leekuokming numericalsimulationofhypersonicflowoverabluntleadingedgedeltawing |