@inproceedings{oai:jaxa.repo.nii.ac.jp:00037858,
 author = {黒滝, 卓司 and 中村, 淑子 and 黒崎, 隆二郎 and 片山, 雅之 and 麻生, 茂 and Kurotaki, Takuji and Nakamura, Yoshiko and Kurosaki, Ryujiro and Katayama, Masayuki and Aso, Shigeru},
 book = {航空宇宙技術研究所特別資料, Special Publication of National Aerospace Laboratory},
 month = {Dec},
 note = {極超音速流中の軸対称物体における吹出しによるフィルム冷却効果を数値的に調べ、その結果を実験データと比較した。実験は在来型の衝撃風洞を用いて行った。主流マッハ数4.25で、半球状頭部模型の先端に設けたスロットから冷却用ガス(N2)を供給する。吹き出しは模型表面に接する向きで物体表面上の熱伝達を計測した。数値解析では、陰的有限差分解法を用いて軸対称完全ナビエ･ストークス方程式を解いた。時間積分にはLU-SGSを用い、対流項はRoeタイプの流束分割に基くMUSCL型のTVDスキームを用いた。数値解析結果は表面熱流束が著しく減ることを示し、これは実験でも同様で、数値解析と実験とは定性的に一致する。境界層内の流れを詳細に見ると2層から成り、内層は物体表面に沿う断熱壁の役割を果している。これらの特性は、この種の極超音速流におけるフィルム冷却効果の決定に本質的な関わりをもつと考えられる。, Film cooling effects due to mass addition on axisymmetric body in hypersonic flow field are numerically investigated and results of numerical simulation are compared with experimental data. Experiments are conducted by using a conventional shock tunnel. Free stream Mach number is 4.25 and cooling gas (N2) is supplied through the slot located at the nose of a hemisphere model. The direction of mass addition is tangential from the model surface and heat flux around the body is measured. In numerical analysis, axisymmetric full Navier-Stokes equations are solved by an implicit finite difference method. Lower Upper Symmetric Gauss Seidel (LU-SGS) scheme is used for the time integration and convective terms are evaluated by the MUSCL (Monotonic Upstream Schemes for Conservation Laws)-type Total Variation Diminishing (TVD) scheme based on Roe-type flux splitting. Numerical results show significant decreases of surface heat flux which are also obtained in experiments and agreement of numerical and experimental results are qualitatively good. More detailed investigation of the flow inside the boundary layer indicates that the boundary layer consists of two sublayers and that the inner layer plays a role of adiabatic wall structure on the surface. These characteristics are considered most essential in the determination of film cooling effects in this type of hypersonic flow., 資料番号: AA0004174021, レポート番号: NAL SP-27},
 pages = {165--170},
 publisher = {航空宇宙技術研究所, National Aerospace Laboratory (NAL)},
 title = {極超音速流れにおける軸対称物体回りのフィルムクーリング冷却法に関する研究},
 volume = {27},
 year = {1994}
}