@inproceedings{oai:jaxa.repo.nii.ac.jp:00036825,
 author = {坪井, 伸幸 and 海田, 武司 and 野本, 秀喜 and 児玉, 優 and 佐藤, 敬二 and 吉永, 崇 and 野田, 順一 and 関根, 英夫 and 楯, 篤志 and 渡辺, 光則 and Tsuboi, Nobuyuki and Kaeda, Takeshi and Nomoto, Hideki and Kodama, Masaru and Sato, Keiji and Yoshinaga, Takashi and Noda, Junichi and Sekine, Hideo and Tate, Atsushi and Watanabe, Mitsunori},
 book = {航空宇宙技術研究所特別資料, Special Publication of National Aerospace Laboratory},
 month = {May},
 note = {1992年の風洞試験結果の要約は下記の通りである。(1)超音速での完全航空機の基礎空力特性を得た。マッハ数2.0での抗力比7.3までの最大揚力(トリムなし、実サイズ補正)を得た。(2)超音速での中性点は亜音速の場合に比較して約6パーセント後方に移動した。(3)超音速巡行速度でのナセル周辺のオイル流れテストは空気取入口でオイル流出が起きていることを示した。したがって、トリム抗力増大を防ぐため、最適重心位置に対するスタビライザ制御の有効性に関するデータ取得のための風洞試験を1993年に実施することを決めた。1993年の風洞試験結果の要約は下記の通りである。(1)スタビライザ制御の有効性を取得し、最適重心を得た。(2)マッハ数2.0での抗力比7.2までの最大揚力(トリム付き、実サイズ補正)を得た。(3)縦方向安定性補償制御が揚抗比の改善に有効であることを確認した。1993年の試験結果として、トリム付きの抗力に対する揚力を超音速巡行速度でのスタビライザ制御の有効性により得た。抗力増大の原因と考えられるナセル前面のオイル流出現象(1992年のオイル流れテスト結果を参照)を揚抗比の改善のため1994年に調査することを決定した。1994年の風洞試験結果の要約は下記の通りである。(1)ナセル前面のオイル流れパターンの原因はオイル流出ではなく、ナセルからの衝撃波と表面下の翼の境界層との間の複雑な相互作用であると考える。(a)オイル流れパターンの本質的な変化は種々のダイバータ高さで見られず、ナセル内部流のマスフロー比から判断するとオイル流出は元のダイバータ高さでも少量であると結論できる。(b)風洞試験のオイル流れの結果を超音速での平板に取付けたくさびの流れ場構造と比較すると、ナセル前面のオイル流れパターンはナセルからの衝撃波と平板境界層の間の相互作用によるものと結論できる。(2)この風洞試験では種々のナセル方位角での完全航空機の抗力への影響はわずかであった。, Results of wind tunnel testing in 1992 are summarized as follows: (1) the basic aerodynamic characteristics of a complete aircraft at supersonic speed was obtained and the maximum lift to drag ratio of 7.3 (with no trim, corrected to actual size) at M = 2.0 was acquired; (2) the neutral point at supersonic speed moves backward approximately six percent in comparison with that at subsonic speed; and (3) oil flow testing around the nacelle at supersonic cruise speed shows that spillage occurs at the air intake. Therefore, it was determined that in order to prevent any trim drag increment, wind tunnel testing acquire data on the stabilizer control effectiveness for the optimum center of gravity position should be conducted in 1993. Results of wind tunnel testing in 1993 are summarized as follows: (1) the stabilizer control effectiveness was acquired and the optimum center of gravity was obtained; (2) the maximum lift to drag ratio of 7.2 at M = 2.0 (with trim, corrected to actual size) was obtained; and (3) it could be confirmed that the longitudinal stability compensation control was effective for improvement of lift to drag ratio. As a result of the 1993 tests, the lift to drag with trim could be acquired by stabilizer control effectiveness at supersonic cruise speed. Hence, it was determined that the spillage phenomenon in front of the nacelle (refer to the oil flow testing results in 1992) that was considered to be a cause of drag increment, should be investigated in 1994 in order to improve lift to drag ratio. Results of wind tunnel testing in 1994 are summarized as follows: (1) it is considered that the oil flow pattern in front of the nacelle is not caused by the spillage but rather by the complicated interaction between the shock waves from the nacelle and the boundary layer on the wing under surface. (a) No essential change of oil flow pattern can be seen with various diverter heights and it can be concluded that the spillage is a small amount even with the original diverter height judging from the mass flow ratio of the nacelle internal flow. (b) As the oil flow results in the wind tunnel testing are compared with the structure of the flow field for a wedge mounted on a flat plate at supersonic speed, it can be concluded that the oil flow pattern in front of the nacelle is due to the interaction between the shock waves from the nacelle and the flat plate boundary layer; and (2) little effect on the drag of the complete aircraft with various azimuthal angles of nacelle is obtained in this wind tunnel testing., 資料番号: AA0000440002, レポート番号: NAL SP-31},
 pages = {17--37},
 publisher = {航空宇宙技術研究所, National Aerospace Laboratory (NAL)},
 title = {Drag measurement of a complete aircraft model in wind tunnel testing},
 volume = {31},
 year = {1996}
}