| Project/Area Number |
08455465
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| Research Category |
Grant-in-Aid for Scientific Research (B)
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| Allocation Type | Single-year Grants |
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| Section | 一般 |
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| Research Field |
Aerospace engineering
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| Research Institution | OSAKA PREFECTURE UNIVERSITY |
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Principal Investigator |
NISHIOKA Michio OSAKA PREFECTURE UNIVERSITY,COLLEGE OF ENGINEERING,PROFESSOR, 工学部, 教授 (60081444)
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| Co-Investigator(Kenkyū-buntansha) |
SAKAUE Shoji OSAKA PREFECTURE UNIVERSITY,COLLEGE OF ENGINEERING,RESEARCH ASSOCIATE, 工学部, 助手 (70244655)
ITOH Nobutake NATIONAL AEROSPACE LABORATORY, 空力性能部, 室長
村上 洋一 大阪府立大学, 工学部, 助教授 (90192773)
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| Project Period (FY) |
1996 – 1998
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| Project Status |
Completed (Fiscal Year 1998)
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| Budget Amount *help |
¥7,400,000 (Direct Cost: ¥7,400,000)
Fiscal Year 1998: ¥1,200,000 (Direct Cost: ¥1,200,000)
Fiscal Year 1997: ¥1,800,000 (Direct Cost: ¥1,800,000)
Fiscal Year 1996: ¥4,400,000 (Direct Cost: ¥4,400,000)
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| Keywords | 3D boundary layer / Supersonic wing boundary layer / Attachment-line boundary layer / Laminar flow control / Boundary layer instability / Boundary layer receptivity / Crossflow instability / Subcritical transition |
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| Research Abstract |
It is a technological challenge posed by the next generation supersonic transport to reduce the skin friction drag and to realize a high L/D ratio above 10 at supersonic cruising. For this purpose it has been proposed to apply laminar flow control (LFC) for the wing boundary layer. To achieve such LFC we have to develop much fundamental knowledge on the receptivity and instability of supersonic boundary layer growing in the leading edge region. This is exactly the motivation of our numerical and experimental studies. Through investigating the supersonic receptivity and instability problems we obtained the following results. Numerically the receptivity of the supersonic boundary layer (at a free-stream Mach number 2.2) is found to be very week, with the excited amplitude of TS wave being as low as 1/10 compared with the corresponding incompressible case. By disturbing the boundary layer near the attachment-line of a swept wing (with swept angle 60 to 70 degrees) in Mach 2.5 free-stream it is numerically shown that streamwise vortices of cross-flow instability type appear with the highest growth. For the case of a Mach 2.5 flat-plate boundary layer disturbed by strong periodic jets from wall orifice we learned numerically that hairpin vortices develop like in low-speed boundary layers, suggesting the occurrence of the non-linear subcritical transition. We made careful experiments to clarify how the acceleration can suppress the subcritical transition and found that the critical Reynolds number (based on momentum thickness) for the subcritical transition increases from 120 for the flatplate flow to 300 for a low-speed boundary layer with a non-dimensional favorable pressure gradient, -4.5x10^<-3>, which being scaled with the dynamic pressure and momentum thickness. It is also stressed that we found the instability due to streamline curvature to exist in the leading edge region of a swept wing in addition to the cross flow instability.
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