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Phased-Array Microphone system has been developed at 2 m x 2 m Low-speed Wind Tunnel (LWT2) of JAXA. As a function checkout of the system, NACA0012 airfoil was tested. The results of this test were observed distinctly the noise generated near or on the trailing edge. In this paper, the outline of the results is reported., 資料番号: AA0063156011, レポート番号: JAXA-SP-06-013

The wing and flap have, generally, certain amounts of thickness at the trailing edge. It is known that the effect of the blunt trailing edge on the lift and drag coefficients of an airplane is not negligible. However, most of the flow computations of airplanes assume the sharp trailing edges because of the numerical cost in treating the blunt trailing edges. In this paper, the effect of trailing edge thickness of a 3D wing model is investigated. The DLR-F4 wing-body configuration that was used at the first AIAA CFD Drag Prediction Workshop is employed as a baseline model. In order to generate a high dense mesh around the wing especially at the trailing edge, the mesh refinement technique is effectively utilized. Computations of DLR-F4 with/without blunt trailing edge are performed at transonic cruising Mach number of M(sub infinity) = 0.75. In addition, chord extension models with sharp trailing edge are calculated. The computed results show that the flow structures near the trailing edge are significantly affected by the thickness of the trailing edge. This effect is larger in the wing tip region. The aerodynamic coefficients of the wing with blunt trailing edge are similar to that of the chord extension model with sharp trailing edge., 資料番号: AA0047427040, レポート番号: JAXA-SP-03-002

航空宇宙技術研究所 12-13 Jun. 1997 東京 日本, National Aerospace Laboratory 12-13 Jun. 1997 Tokyo Japan, 本論文で考察する問題については、その解法を関数論の枠組みの中で十分に確立している。その解法は、単位円の外部全体を考察している翼形の外部全体に写像する等角写像が既知であれば容易に適用できるものである。しかし幾つかの場合には、あらかじめ設定した精度を満たすように等角写像を構成することは大変な数値計算の仕事になると予測される。本論文では、クッタの条件、または、ジューコフスキーの仮定を後縁において満たす翼周りの流れの決定に有限要素法を用いることができることを示した。翼形Cに対する主要な制限は以下のようである。翼の後縁Pの近くでは翼上面と翼下面は正の角度をもってPで交わる2直線と一致している。この解法の有効性を示すため、NACA(国家航空諮問委員会)23012翼形の数値結果を本論文の後半で報告した。, As for the problem considered in the paper, the solution method has been well established within the framework of function theory. The method is easily applicable to the case in which there is the conformal mapping which transform the whole exterior domain of the unit disc onto that of the wing profile to be considered. In some cases, however, it seems to be a heavy numerical task to have an approximate conformal mapping with accuracy to be expected. In this work, it is demonstrated that the finite element method can be applied to determine the flow around a wing satisfying the Kutta condition (or the Joukowski assumption) at its trailing edge. Main restriction on the shape of the wing, C, is the following: near the trailing edge point P of the wing, the upper surface and the lower surface coincide with two straight lines intersecting at P with an angle of positive degree. To show the effectiveness of the method, the numerical result for NACA23012 wing is reported in the latter half of this paper., 資料番号: AA0001433055, レポート番号: NAL SP-37

航空宇宙技術研究所 6-7 JUN. 1996 東京 日本, National Aerospace Laboratory 6-7 JUN. 1996 Tokyo Japan, ナビエ・ストークスコードを用いて、ピッチング振動と上下振動が位相差をもって連成モードで振動する翼のまわりの失速現象について数値シミュレーションを行った。異なる2種の羽ばたきパラメータについて、位相差と低振動数の種々の組合わせに対して推進効率と推進力を計算し、計算から得られた各フローパターンを検討して、ダイナミック・ストール現象が推進効率と推進力におよぼす効果を詳細に論じた。ピッチング振動が上下振動よりも90度進んでいる時に最大推進効率を観測した。大振幅振動にも拘らず流れの剥離は後縁近傍の翼面上の小領域に局所化していた。90度以外の位相角については、大規模な前縁剥離が発生して推進効率は低下した。, Numerical simulation of dynamic stall phenomena around an airfoil oscillating in a coupled mode, where the pitching and heaving oscillations have some phase differences, has been performed using the Navier-Stokes code. The propulsive efficiency and the thrust have been calculated for various combinations of the phase difference and the reduced frequency for two different feathering parameters, and the effects of the dynamic stall phenomena on the behaviors of the propulsive efficiency and thrust are also discussed in detail by examining each flow pattern obtained. The highest efficiency has been observed for the case where the pitching oscillation advances 90 deg over the heaving oscillation, for which the flow separation is confined in the small region on the airfoil surface near the trailing edge in spite of the large amplitude oscillations. For phase angles other than 90 deg the efficiency is degraded by the occurrence of the large scale leading edge separation., 資料番号: AA0000685049, レポート番号: NAL SP-34

遷音速域を運用範囲にもつ超高速輸送機においては、設計の初期段階から構造の強度設計だけでなく空力弾性設計も同時に行う必要がある。この空力弾性設計に必要な遷音速域での空力弾性現象、特にフラッタ速度推算用解析ツールは現在検証段階にある。2003年度のJADC-FHI-JAXAの共同研究「超高速輸送機のフラッタ特性に関する研究」においては、遷音速領域におけるフラッタ特性推算手法の確立のために、超音速機の弾性模型によるフラッタ風洞試験をJAXAの0.6m×0.6m遷音速フラッタ風洞設備にて行い、解析ツール検証に必要なデータを取得した。今回フラッタ風洞試験に用いた弾性模型は、旧航空宇宙技術研究所で計画され基本設計を行った小型超音速ジェット実験機(1次形状)の遷音速フラッタ特性取得のために製作した11％スケール弾性模型を改修したものである。フラッタ発生時に模型損傷回避策が有効に働いたので、同一のフラッタ模型にて短期間に6つのフラッタポイントを取得することができた。, 資料番号: AA0048140001, レポート番号: JAXA-SP-04-011

Wake structure behind an NACA0012 with no incidence was examined to scrutinize the frequency selection mechanism of growing disturbances at Re = 7,200 based on the wing chord and the free-stream velocity. Two different kinds of unstable disturbances, which exponentially grow in space, were observed. The minor disturbance is due to convective instability, while the origin of the major one may be attributed to absolute instability. Theoretical comparison is necessary to identify these disturbances., 資料番号: AA0063156004, レポート番号: JAXA-SP-06-013