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航空宇宙技術研究所 16-18 Jun. 1999 東京 日本, National Aerospace Laboratory 16-18 Jun. 1999 Tokyo Japan, 翼の失速特性と関連する大迎角流れ解析は多くが試みられ、すでに解決済みとの印象がある。しかし、定量的な信頼性については疑わしい。格子解像度、乱流モデル、計算アルゴリズムなどについて、精度のよい予測のために必要な知見を期待して、CFD(数値流体力学)ワークショップを計画し、以下の要項で課題募集を行った。(1)課題の目的は、CFDによる低速での翼型失速角の予測について、現状を把握し改善するための指針を明らかにすることである。(2)3種類の翼型を課題とする。(3)計算条件は、指定された迎角範囲を指定された角度刻みで、指定されたレイノルズ数とマッハ数について計算する。(4)流体力学方程式モデル、乱流モデル、および計算手法は、任意とする。その他、結果の提示、提出データファイル、応募締め切り、および問い合わせ先について示した。, Many analyses of flow at high attack angles, relating to the stall characteristics of airfoils, have been attempted so far, and they give the impression that most of problems are already solved. But whether the problems of reliability were quantitatively solved is questionable. CFD (Computational Fluid Dynamics) workshop is planned to get the necessary knowledge of grid resolution, turbulence model and computational algorithm for the accurate prediction, and begins with an invitation of papers satisfying the following requirements. (1) The aim of the papers is to clarify the guideline to grasp and improve the present status in predictions of the airfoil stall angle at low speed using CFD. (2) The airfoil types are restricted to three. (3) The model of hydrodynamic equations, turbulence model and computational method are arbitrary. Moreover the presentation of results, the data file are to be submitted, the close of the subscription, and the reference is to be shown., 資料番号: AA0001961005, レポート番号: NAL SP-44

24-25 Jul. 1996, 以下の研究課題について論じた。(1)NALとNASDAの宇宙往還機開発計画、(2)H-2AとJ-1ロケット開発、(3)HYFLEX(極超音速飛行実験)、ALFLEX(小型自動着陸実験)、HOPE-X(宇宙往還技術試験機)、(4)飛行経路設計、誘導則、姿勢制御、(5)電気系、熱防護系の開発、(6)CFD(数値流体力学)解析、非定常熱解析、リフレクトメータによるプラズマ電子密度測定。, The following topics are discussed: (1) spaceplane development plan at NAL/NASDA; (2) H-2A and J-1 rocket development; (3) HYFLEX (Hypersonic Flight Experiment), ALFLEX (Automatic Landing Flight Experiment), and HOPE-X (H-2 Orbiting Plane-Experimental) program; (4) trajectory design, reentry guidance law, and flight control system; (5) development of electrical and thermal protection system; and (6) CFD (Computational Fluid Dynamics) analysis, transient heat conduction simulation, and plasma electron density measurement by reflectometer., 資料番号: AA0000584000, レポート番号: NAL SP-32

An experimental study was conducted to investigate transition process on 10-deg cone. There are two goals of this study. One is to validate transition prediction numerical code LSTAB, which is recently developed. Instability wave is detected by use of flush-mounted a pressure transducer and a hot-film sensor, and its frequency is confirmed to agree with the numerically predicted value. The second goal is to clarify the effects of freestream acoustic disturbances on the boundary-layer transition in supersonic flow. Intensity of freestream disturbances was controlled by a strip of sand paper or distributed roughness, which were attached on the floor or sidewall near the throat. High-frequency disturbance increased in freestream when sand paper was attached on the floor, though freestream disturbance was not affected by larger roughness on straight sidewall. The effect of roughness on the wind tunnel wall is pointed out to be different with the position and condition., 資料番号: AA0045947015, レポート番号: NAL SP-56

この技術論文は、航空宇宙技術研究所の大型低速風洞(6.5m×5.5m)における模擬エンジン付き模型の試験装置について述べる。この実験装置の目的は、配管干渉を除去することによって模型に加わる空気力を高精度で計測するものである。模型は、8%STOL実験機'飛鳥'模型を用いた。本論文では、配管干渉除去機能と試験装置全体について説明する。, This paper describes a test system of an eight percent scale model of STOL (Short Take-Off and Landing) aircraft 'ASKA' with turbine powered simulators tested at the low speed wind tunnel (6.5 m x 5.5 m) of National Aerospace Laboratory (NAL). The purpose of the system is to achieve high accurate measurements of aerodynamic forces by eliminating piping interferences. The paper also describes performance of the piping interference elimination function and overall system test results., 資料番号: AA0028666000, レポート番号: NAL TR-1421

スペースブレーンは低軌道に達するまでに空気中に約15分以上飛行する。この間、特にスクラムジェットを使用する極超音速飛行時に、スペースプレーンは強い空力加熱を受ける。そこで今回、簡単なモデルを使って加速上昇中の単段スペースプレーン(SSTO)熱防護システムの特性について検討した。タイルによる熱防護材と水素による強制冷却を検討対象とした。飛行経路に沿ったシミュレーションを行い、タイル内温度分布あるいは水素冷却パネルの冷却要求について検討を行った。その結果、以下のことが明らかになった。タイルを用いた熱防護システムの場合、タイル裏面の最高温度は軌道到達後に現れた。従ってタイル設計には、軌道到達後のタイルの状態も考慮する必要がある。今回の検討ではタイル裏面の温度が内部断熱材許容上限温度480Kを越えない場合のタイルの最小厚さは、機体上面で50mm、下面で60mmであった。タイル表面温度を使用限界温度1,500K以下にするためには、飛行動圧を100kPa以下にする必要があった。タイル裏面最高温度は、飛行動圧の低下による飛行時間の増加と共に上昇した。機体の能動冷却に必要な水素質量は燃料水素質量の約20%であった。水素冷却パネル表面にタイルを張り合わせることで、機体の冷却に要する水素質量を減らすことができた。, This paper discusses the Thermal Protection System (TPS) of an airframe of the Single Stage To Orbit (SSTO) with a scramjet engine. There is aerodynamic heating to the airframe during flight through the air, especially during hypersonic flight, and this paper examines ceramic tile insulation and hydrogen active cooling. The transient nature of each system was simulated along the ascent trajectory of the SSTO to the low earth orbit of 100 km. When ceramic tiling was used, a tile thickness of around 60 mm was required to protect the inside of the airframe. The highest temperature on the inside surface of the tile occurred after arrival to the orbit. The temperature on the inside surface increased with the decrease of the flight dynamic pressure. In the active cooling system, around 20 percent of fuel were used for airframe-cooling. Attaching thin ceramic tiling on the surface of the cooling panel reduced the cooling requirement., 資料番号: AA0028665000, レポート番号: NAL TR-1420

航空宇宙技術研究所 7-9 Jun. 2000 東京 日本, National Aerospace Laboratory 7-9 Jun. 2000 Tokyo Japan, この論文では、代数型乱流のナビエ-ストークス・コードの2次元翼型失速予測への応用を示した。3種類の翼型について計算し、数値結果と実験データを比較したところ、低仰角において良い一致を示した。完全な乱流条件では失速角と最大揚力の過大評価が見られたが、この過大評価は遷移点を適切に設定することで克服できた。この論文では、剥離型の遷移点と失速角の効果について調べ、代数型乱流モデルが翼型失速をどれほど予測できるかを示した。, This paper presents the applications of Navier-Stokes code with an algebraic turbulence to 2D (two dimensional) airfoil stall prediction. Three kinds of airfoil have been calculated and the comparison between numerical results and experimental data show good agreement in the region of low angle of attack. Overestimation of stall angle and maximum lift has observed in case of full turbulence condition, however, the overestimation can be overcome by setting transition points appropriately. In this paper the effect of transition point on separation type and stall angle is investigated and the results show the ability of algebraic turbulence model to predict airfoil stall., 資料番号: AA0028635026, レポート番号: NAL SP-46

航空宇宙技術研究所 7-9 Jun. 2000 東京 日本, National Aerospace Laboratory 7-9 Jun. 2000 Tokyo Japan, 本論文では、航空機エンジン、船舶、発電所のターボ機械の空力設計技術者のために効率的な技術環境を実現する翼列設計用のCFD(計算流体力学)プラットフォームについて解説した。プラットフォームをネットワークで接続したPC、ワークステーション(WS)、スーパコンピュータから構成した。このプラットフォームでは設計者が計算グリッドを生成したり、流体計算をしたり、計算結果を解析、視覚化する際に、その操作が容易になる様、CFDプログラムをこれらのコンピュータで運用するためのグラフィックユーザインタフェースを提供した。また、翼列の幾何学的形状のようなデータや、計算した流体のデータの様に空力性能を比較したり、翼列の設計を評価するのに必要なデータをデータベースに保存したり、参照したりする機能も保持した。本稿では、2次元の翼列のために開発したプラットフォームのプロトタイプを紹介した。多段ターボ機械のための流体解法コードも並列計算ライブラリとともに開発し、このプラットフォームに取り込んだ。, The present paper shows a CFD (Computational Fluid Dynamics) platform for blade rows, which realize efficient engineering environment for aerodynamic designers of turbomachine for aircraft engines, ships and electric power plants. The platform consists of PCs, WSs (Workstations) and supercomputers on network. It provides graphic user interfaces on these computers to control the CFD programs, so that the designers can operate very easily their tasks such as generating the computational grids, calculating the flow, and analyzing/visualizing the computed results. It also has the functions to save/restore in the database such data as geometric profiles of blade rows, computed data of the flow, which data are necessary to compare the aerodynamic performance and estimate the design of blade row. In this paper, a prototype platform is demonstrated, which has been developed by the authors for two-dimensional blade rows. A flow solver for a multi-stage turbomachine is also developed with parallel computing library and incorporated in this platform., 資料番号: AA0028635042, レポート番号: NAL SP-46

航空宇宙技術研究所 16-18 Jun. 1999 東京 日本, National Aerospace Laboratory 16-18 Jun. 1999 Tokyo Japan, 本論文では大規模設計および制御最適化問題を解くための新しい並列最適化法およびその応用について報告した。一般に精度の良い最適解を得るためには大規模の問題を解かなければならないが、これは計算機に大きな負荷と時間をかける。従って新しい方法では問題を並行して解けるいくつかの小規模問題に分割する。本論文では最初に問題を分割する方法とこれを解くための基本的なアルゴリズムについて述べた。この研究の重要な点は小規模問題間の接続拘束条件をどのように取り扱い、それぞれの小規模問題の目的関数をどのように定義するかである。最後に、本報告では並列最適化法を宇宙航空機の機体形状と飛行軌道の最適化に適用した。その最適形状と上昇軌道により提案した並列最適化法の有効性を確認した。, This paper presents a new parallel optimization method to solve large-scale design and control optimization problems and its applications. Generally, to obtain an accurate optimal solution, a large-scale problem has to be solved, and it takes much computing load and time. Therefore, the new method divides the problem into several sub-problems, which can be solved in parallel. Firstly, this paper describes the way to decompose the problem and the fundamental algorithm to solve it. The important point of this study is how to deal with the conjunctive constraints among the sub-problems and define objective functions in each sub-problem. Finally, in this paper, the parallel optimization method is applied to a shape and flight trajectory optimization problem of spaceplane, and its optimal shape and ascent trajectory are estimated to demonstrate the effectiveness of the proposed parallel optimization method., 資料番号: AA0001961034, レポート番号: NAL SP-44

航空宇宙技術研究所 16-18 Jun. 1999 東京 日本, National Aerospace Laboratory 16-18 Jun. 1999 Tokyo Japan, 本稿は、「CFD(数値流体力学)は万能か?」というワークショップの紹介であり、ワークショップパネルでの議論を活性化させるためのものである。70年代の初期からCFDは大きく進歩し、今では3次元の問題すらも、大きすぎなければパソコンを使って解くことができる。また同時に、方法論の進歩も著しい。しかしながら、いくつかの障害がまだ残されている。まだ、解の信頼性、シミュレーションの全過程における労力を最小化する戦略、計算時間の削減などに関する議論は重要である。本稿ではこれらの障害を克服するための考えを議論した。例として筆者の経験に基づいて、5つの考えを列挙した。流体力学を理解することの重要性、新しい単純化された方程式群の開発、重合デカルト格子手法、新しい型の後処理法およびCFDデータの蓄積を示し、議論した。, This paper is an introduction of workshop 'What can we do with CFD (Computational Fluid Dynamics)?' to stimulate the discussion in the workshop panel. CFD has made great progress since early 70's, and now even three-dimensional problems, if not too large, can be solved using PC's. At the same time, the progress of the methodologies was remarkable. However, there are still some obstacles remained. The discussions on the reliability of the solutions, of the strategy to minimize the man-time for the whole process of the simulations, of the reduction of computer time, and else are still important. In the present paper, ideas to overcome these obstacles are discussed. As examples, five ideas are itemized from the author's experience. Importance of understanding fluid physics, development of a new set of simplified equations, Cartesian-overset grid approach, a new type of post-processing and accumulations of CFD data are presented and discussed., 資料番号: AA0001961004, レポート番号: NAL SP-44

本稿では、1996年2月12日に打ち上げられた極超音速飛行実験機(以下HYFLEX)のマッハ2以下の飛行について述べる。飛行実験では機体の回収には失敗したものの、テレメトリデータ、レーダトラッキングデータが良好に取得された。本稿ではこれら取得されたデータに基づき、マッハ2以下の飛行について再現を試みた。得られた結果から、HYFLEXがスピン状態で降下していく様子が概ね把握できた。, This paper describes the actual flight trajectory under Mach 2 of the Hypersonic Flight Experiment vehicle (HYFLEX). Although the recovery of the vehicle failed, radar tracking and telemetry data were obtained. In this paper, these data are analyzed and the flight trajectory under Mach 2 is reconstructed. It was found that the vehicle flew in a random spin and not according to the preflight plan. The reason for the failed maneuver was proved by postflight simulation., 資料番号: AA0001567000, レポート番号: NAL TM-726

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

本稿では、1996年2月12日に打ち上げられた極超音速飛行実験機(HYFLEX)の航法・誘導・制御系解析に用いた飛行シミュレーションプログラムと飛行シミュレーション解析結果について報告する。従来、航空宇宙技術研究所(NAL)では、再突入機、あるいはスペース・プレーンの飛行シミュレーションを行う際の運動方程式は、地球を球と仮定して導出されていた。また、HYFLEXの基本設計段階においても、地球の取り扱いを球とした仮定の基で導かれた運動方程式で飛行シミュレーションを行っていた。しかし、HYFLEXの回収に対する信頼性、成功率を上げること、および飛行安全性確保の観点からより精度の高い飛行シミュレーション解析の要求が生じた。そこで本稿では、地球をより精度の高い回転楕円体モデルと仮定して取り扱う。運動方程式導出に際しては、地球モデルの変更に伴い局所水平座標系の定義が従来のものと一致しなくなるため、新たに地心局所座標系と対気速度座標系を定義する。さらに、地球モデルの変更に伴い機体に作用する外力である重力ベクトルも変化する。このため、重力モデルも逆自乗場のモデルから、重力ポテンシャルによるモデルへと変更した。また、計算効率、計算精度を上げるため、新たに2つの状態量を導入した。本稿ではまず最初に、上記に伴い新たに開発された飛行シミュレーションプログラムについて述べる。そして、開発した飛行シミュレーションプログラムにより行った、打ち上げ前の最終的な飛行シミュレーション解析結果についても報告する。得られた飛行シミュレーション解析結果により、設計されたHYFLEXの航法・誘導・制御系が所望の性能を有していることが確認された。また、着水点、および着水点分散域に関するより精度の高いデータを得た。, This paper describes a flight simulation program, used to analyze the navigation, guidance and control system and for flight analysis of the Hypersonic Flight Experiment vehicle (HYFLEX). Up to this time, the earth had been considered as a sphere at NAL. The flight simulation was carried out using the model at the predesign phase of the HYFLEX. However, more precise flight analysis had to be carried out to predict the splash down point as planned to recover the HYFLEX vehicle. In this paper, two coordinate systems were newly introduced to deal with the earth as a ellipsoid of revolution and new equations of motion were derived. Also, the gravity model was changed to the model mentioned in reference. To maintain precision, two state vectors were newly defined. The designed navigation, guidance and control system of the HYFLEX is proved to satisfy the mission requirements through the results of flight analysis. Further, more accurate data of the nominal impact point and dispersion of the impact point are obtained., 資料番号: AA0000924000, レポート番号: NAL TR-1317

24-25 Jul. 1996, 極超音速飛行実験においては、5分間程度の持続的な極超音速飛行時間が確保されるため、往還機に重要なM＞10領域の空力、空力加熱特性データが質量とも充実したレベルで得られることが期待された。機上計測項目の選定においては、これら極超音速領域の実在気体効果を含む空力、空力加熱特性データ、熱構造の特性評価用データ、および主目的のひとつである揚力飛行体の基本技術評価に関わるデータ(姿勢・軌道制御に関するデータなど)などを抽出し、バス系の計測項目をあわせて約250チャンネルの計測計画をたてた。各計測項目を選定した基準、設計のための事前解析・試験、各項目の計測系仕様について述べる。具体的議論は主に実験計測系に限った。また、計測の目的は宇宙往還機設計などへ利用できる予測ツール・手法の確立に資することである。そのため、地上での試験やCFD(数値流体力学)による予測とフライトデータとを比較するが、その際に飛行時の大気状態を可能な限り正確に把握する必要がある。今回は大気状態量の特定のため、NOAA衛星のリモートセンシングデータを使用し推定精度を上げることを試みた。これを含むデータ解析計画の概略を述べた。, In the HYFLEX (Hypersonic Flight Experiment), a consecutive hypersonic flight duration of some five minutes is secured. Therefore, it is expected to get the hypersonic aerodynamic/aerothermodynamic performance data good enough in both qualitative and quantitative sense for the region in which M is greater than ten, where best performance is crucial for the reusable space vehicles. Selection of the measurement items for the onboard measurement system was made among both aerodynamics and aerothermodynamics including the real gas effects, thermo-structural evaluation data, basic data for evaluation of lifting body technology (guidance, navigation and control) etc. consisting of some 250 channels in total. Decision to select measurement items, the analyses and tests for design and confirmation, specifications of each items are outlined and discussed in this paper. But mostly measurement mission items are discussed in this paper. Major objective of HYFLEX onboard measurement is to contribute to establishing reliable prediction/design methods for the winged space vehicles. To do this, the comparison between the flight data and the predictions made in advance based on ground test and analyses is indispensable. For the meaningful comparison, the atmospheric condition along flight path must be known accurately enough. In HYFLEX, the utilization of the remotely sensed atmospheric data by NOAA meteorological satellite was planned to raise the inference accuracy. Data analysis program is briefly presented., 資料番号: AA0000584014, レポート番号: NAL SP-32

航空宇宙技術研究所 7-9 Jun. 1995 東京 日本, National Aerospace Laboratory 7-9 Jun. 1995 Tokyo Japan, 物体表面スプラインネットにより物体形状の表現を行った。スプライン曲線間の点でスプライン内挿関数を用いて形状を内挿し、物体の全体形状を連続的に定義する。ここでは、離散点として与えられるオフセットデータを基にスプラインネットを生成する手法を示した。本手法は格子形成を拡張したもので、IGM(陰的形状法)と呼ぶ。与えられた平滑化、クラスタ化、直交性の要求をバランスよく満たしながらスプラインネットを生成する。前半では平滑化演算子として用いた4階差分演算子がFergusonスプラインと類似の特性を持つことを示した。後半では例として船体表面スプラインネット生成の実行手順を示した。生成のスプラインネットは与えられた要求を満たし、オフセットのスプライン各点を通っていることを確認した。, Body geometries are generally expressed by a spline net covering the body surface. At points between the spline curves, the geometry is interpolated using spline interpolation functions, thus the body geometry is defined continuously all over the surface. In this paper, a method is shown to generate a spline net based on the offset data given at discrete points. The method is an extension of a grid generation method called the implicit geometrical method. The smoothing, clustering, and orthogonality requirements are imposed, and the spline net is generated to satisfy those requirements with suitable balance. In the first half of the paper, it is shown that the fourth-differencing operator used as the smoothing operator has characteristics quite similar to cubic Ferguson splines. In the second half, a spline net generated on a ship hull form is shown as an example. The generated spline net satisfies not only the imposed requirements, but also the constraint that the net passes through the offset spline points., 資料番号: AA0000294024, レポート番号: NAL SP-30

プリンストン大学A.Jameson教授による招待講演｢航空機用CFDの現状｣を含め合計57件の論文を集めた。取り扱われている課題は、従来の物体回りの流れの数値解析のほか、乱流の直接シミュレーション、ターボジェットおよびラム/スクラムジェットエンジン燃焼器の中の流れ、航空宇宙技術研究所の並列ベクトル計算機システム｢NWT｣の性能評価などを含む。, Altogether 57 papers are collected, among which is the invited lecture by Professor A. Jameson of Princeton University giving an overview of the current status of CFD (Computational Fluid Dynamics) as the aircraft aerodynamic design tool. In addition to the usual computational analysis of external flow, the topics of the papers include the direct numerical simulation of the turbulence, the combustor flows within turbojet as well as ram/scramjet engines and the performance assessment of the NAL parallel vector computer system called, 'NWT'., 資料番号: AA0004174000, レポート番号: NAL SP-27

境界層の層流制御や乱流制御を考察する場合に、壁乱流が自己維持的であるようなレイノルズ数の下限における乱流特性を知る必要がある。しかしながら、これに関連した情報は、これまでのところ極めて限られた不十分なものでしかなく、自己維持的な壁乱流に関する流れの臨界条件およびそれに結びついた亜臨界遷移のための最小レイノルズ数は未だ明らかにされていない。本論文では、平板の前縁近く、または風洞縮流筒のすぐ下流の位置に置かれた一群の円柱(直径3mm,高さ3.8mm)によりひき起された平板上の亜臨界遷移を観測することにより、自己維持的な壁乱流についての最小臨界レイノルズ数を実験的に調べた。円柱によって生じたヘアピン渦は、運動量厚レイノルズ数Rθが110-130という低い値でも亜臨界擾乱を増幅することがわかった。その下流でRθが210を越えクラウザパラメタG=(Uoo/uτ)×(H-1)/H(uτは摩擦速度、Hは形状パラメタ)が7より小さくなるような位置では、平均流は対数分布法則に従うようになる。これらの結果を記述することにより、本論文は円柱の作り出すヘアピン渦による亜臨界擾乱の成長について論じた。, When considering Laminar Control and/or Turbulence Manipulation of boundary layer, it is necessary to know the turbulence characteristics at the lowest Reynolds numbers for the wall turbulence to be self-sustained. However, the related information is so far quite limited and insufficient. In fact, the critical flow condition for self-sustaining wall turbulence and the related minimum Reynolds number for the possible subcritical transition are not clarified yet. In the present paper, through observing the subcritical transition of a flat-plate boundary layer triggered by a group of circular cylinders (of 3 mm dia. and 3.8 mm height) placed near the leading edge or on a wind wall immediately downstream of the contraction, the minimum critical Reynolds number for self-sustaining wall turbulence is examined experimentally. It is found that the cylinder-generated energetic hairpin eddies can trigger the subcritical disturbance growth at such low R(sub theta) (based on momentum thickness) as 110-130. Downstream, the mean flow can follow the log-law distribution beyond R(sub theta) = 210 and if the Clauser parameter G = (U(sub infinity)/u(sub tau)(H-1)/H is below about 7, where u(sub tau) and H denote the friction velocity and the shape factor respectively. With describing these results, this paper discusses the subcritical disturbance growth due to cylinder-generated hairpin eddies., 資料番号: AA0004172003, レポート番号: NAL SP-25

航空宇宙技術研究所 28-29 Sep. 1992. 25-26 Apr.1993 東京 日本, National Aerospace Laboratory 28-29 Sep. 1992. 25-26 Apr.1993 Tokyo Japan, 次の22論文を掲載する。(1)超音速後流における乱流混合の促進、(2)矩形ダクト流遷移の数値シミュレーション、(3)長方形管流の不安定性と乱流2次流の発生、(4)離散近似Navier-Stokes方程式を用いた平行、非平行流の安定性解析、(5)境界層の遷移と非定常剥離一人体模型の音を通じて、(6)円柱近傍後流に関する実験と計算の比較、(7)Rc附近において回転円柱に作用する力、(8)カルマン渦列のアクティブ制御について、(9)後退翼における境界層遷移(2)、(10)3次元境界層の横流れ進行波について、(11)超音速輸送機に係わる環境問題と流体力学、(12)ポアズイユ流の2次元分岐構造、(13)後流における絶対カオスと対流カオス、(14)染料流脈と粒子流跡の同時抽出法による秩序運動の時空間構造の検討、(15)凹曲面に沿う縦渦波長の実験、(16)カルマン渦列の周波数決定機構、(17)楕円形不安定性に対する密度成層効果、(18)ベナール対流における形の形成、(19)カルマン渦列の3次元構造-縦渦Fingersの発生機構、(20)リブレット粗面開水路乱流の平均流速分布、(21)粗度の誘起する不安定性、(22)横流れ不安定場における乱流遷移プロセス。, The following research results were discussed: Seven papers on instbility include stability analysis using discretized Navier-Stokes equation, instability of the flow in a rectangular duct, two dimensional bifurcation structure, convective chaos in a wake, pattern formation in Rayleigh-Benard convection. Five papers on transition of boundary layers include roughness induced instability, vortices along a concave wall, traveling waves along a yawed cylinder, turbulent transition process in instability field. Five papers on structure of vorticity include the near wake behind a circular cylinder, three dimensional structure of Karman vortex streets. Four papers on turbulent flow include turbulent flow on a sawtooth riblet surface, and turbulent mixing in superonic wake., 資料番号: AA0004170000, レポート番号: NAL SP-21

航空宇宙技術研究所 10-12 Jun. 1992 東京 日本, National Aerospace Laboratory 10-12 Jun. 1992 Tokyo Japan, 今日、航空機や宇宙機の全機形状まわりのナビエ・ストークス計算が手の届くものとなり、計算空気力学(CFD)の進歩には目を見張るものがあるが、一方では、その信頼性の確立という点では、まだ解決すべき多くのことが残されているように思われる。ある特定の課題を設定し、いくつかの異なるCFDプログラムを走らせた結果を持ち寄って、風洞試験データとの比較などを通してそれらを評価するのも、信頼性確立に向けた努力の一環と考えられる。今回、2次元翼の空力特性計算、および全機形状まわりのナビエ・ストークス計算という課題を設定して、ワークショップを行った。本文は、3次元部門の課題「ONERA M5全機形態まわりの流れの解析」について、課題の詳細や課題募集要項のオーガナイザによる説明、5件の応募論文、オーガナイザによる3次元部門のまとめ、および資料編として、応募者リスト、応募者実施課題一覧、応募者解法情報一覧、応募者提出図などを掲載する。, In the first part of this paper, description of 3D Problem is given. It consists of Problem A, B and C, concerning flow analysis around ONERA (Office National d'Etudes et de Recherches Aerospatiales) M5 configuration. Problem A is to compute the flow at freestream conditions of Mach number, 0.84, Reynolds number, 1 x 10(exp 6) and angle of attack, alpha = -1 deg, with free transition condition. Problem B is to compute the flow at freestream conditions of Much number, 0.84, Reynolds number, 6 x 10(exp 7) and angle of attack. alpha = -1 deg. with full turbulence condition. Problem C is to compute the aerodynamic characteristics of the model at freestream conditions of Mach number, 0.84, Reynolds number, 2 x 10(exp 6) and angles of attack, alpha = -3, -2, and -1 deg, and to make plots of C(sub L) - alpha, C(sub L) - C(sub D), and C(sub L) - C(sub M) curves. The output formats are specified for all computed results. Together with given problems, this subsidiary includes a collection of five papers which are presented by applicants as the analysis results of three dimensional aerodynamic problems., 資料番号: AA0004169002, レポート番号: NAL SP-20

航空宇宙技術研究所 10-12 Jun. 1992 東京 日本, National Aerospace Laboratory 10-12 Jun. 1992 Tokyo Japan, 今日、航空機や宇宙機の全機形状まわりのナビエ・ストークス計算が手の届くものとなり、計算空気力学(CFD)の進歩には目を見張るものがあるが、一方では、その信頼性の確立という点では、まだ解決すべき多くのことが残されているように思われる。ある特定の課題を設定し、いくつかの異なるCFDプログラムを走らせた結果を持ち寄って、風洞試験データとの比較などを通してそれらを評価するのも、信頼性確立に向けた努力の一環と考えられる。今回、2次元翼の空力特性計算、および全機形状まわりのナビエ・ストークス計算という課題を設定して、ワークショップを行った。本文は、2次元部門の課題「Garabedian-Korn 75-06-12翼型の空気力学的特性解析」について、課題の詳細、すなわち、設計点の解析、設計点近傍の空力特性など、の課題募集要項のオーガナイザによる説明、7件の応募論文、オーガナイザによる2次元部門のまとめ、および、資料編として、応募者氏名解法情報一覧、応募ケース一覧、結果一覧表、応募者提出図などを掲載する。, In the first part of this paper, description of 2D Problem is given. The Problem A is to compute a shock-less flow of Garabedian-Korn 75-06-12 Airfoil at design condition of Mach number, 0.75, C(sub 1), 0.63, Reynolds number, 6 and 20 x 10(exp 6). The Problem B is to obtain the aerodynamic characteristics of Mach number sweep and angle of attack sweep. The reference point is the same condition in Problem A. The output data format is described. Together with given problems, this subsidiary includes a collection of nine papers which are presented by applicants as the analysis results of two dimensional aerodynamic problems., 資料番号: AA0004169001, レポート番号: NAL SP-20

次の2論文を掲載する。(1)FGM開発のためのSP/MSP試験法。この研究は、1987年より科学技術振興調整費による総合課題として、産学官30余の研究機関が参加し、材料設計分科会、構造制御分科会、および特性評価分科会の提携のもとに推進されている。各研究機関が独自に信頼性のあるデータを収集できる目的のために傾斜機能材料特有の評価方法について順次評価を進めてきた。開発途上の傾斜機能材料については、従来のJIS試験法に対応する大型の試験片を合成することが困難である。ここに提案する方法は微小試験片でも有効な機械的特性値を取得し、早期に材料設計グループにフィードバックし、最適な材料設計に資することを目的に開発整備された試験法である。(2)セラミックス/金属複合材料のための破壊特性評価システムの開発に関する研究。これは付録として掲載されている。, This report is composed of two related papers: The first one is specification on MSP (Modified Small Punch) method and the second one is research paper on the evaluation of MSP method using ceramic materials as testing material. The first is titled as SP (Small Punch) and MSP (Modified Small Punch) tests for the evaluation of FGM (Functionally Gradient Materials) and defines testing apparatus, preparation method and dimensions of test pieces, and testing procedures. The second one is titled as research on the development of fracture characteristics evaluation system for ceramics and metal composite materials. In this paper fracture characteristics of ceramics/metal composite materials are evaluated using newly developed MSP method and the results are compared with those obtained by using conventional SP method. Microscopic fracture processes are observed by the acoustic emission method. The testing temperature was varied to high temperature of 1873 K., 資料番号: AA0004166000, レポート番号: NAL SP-17

The Numerical Wind Tunnel (NWT) is under research and development by NAL-Fujitsu joint activity. it is a parallel computer system of distributed memory architecture composed of vector processors, whose goal is to perform CFD simulation about 100 times faster than FACOM VP400. In this paper we show three fundamental functions global data, parallel execution of DO-loop, and data decomposition and allocation, which the language-processor system has to provide in order to realize parallel execution on the NWT. We choose FORTRAN77 as a basic programming language for NWT and add some compiler directives to make effective use of the NWT. In this paper, we also explain the programming model of the NWT and the execution image with these directives., 資料番号: NALSP0016018, レポート番号: NAL SP-16

When aero-engines are developed, several different types of wind tunnel tests are involved. For instance, one type of tunnel is required for fundamental tests of the cascades; a second is required for high altitude engine performance tests. Experimental models for these tests consist of real engine components or of close replicas. Certain effects may not be adequately addressed in wind tunnel tests, for example, the effects of high speed revolution, flow properties in unmeasurable regions, realistic boundary conditions. The twofold object of computational fluid dynamics (CFD) for internal flows is to predict the performance of models before expensive wind tunnel tests are done and to provide information of the internal flow that is inaccessible to experiment measurement. CFD is already useful for improving designs or making new ones. As computers and CFD technology advance, CFD can expand from the role of prediction and design to that of a numerical wind tunnel. This paper shows how features of numerical wind tunnels parallel those of experimental wind tunnels The paper also assesses future trends for the use of numerical wind tunnels., 資料番号: NALSP0016053, レポート番号: NAL SP-16

The parallel processing scheme for the Direct Simulation Monte Carlo methed is presented in this paper. The parallel computer of this paper, a multi CPU computer system, consists of three Transputers made by INMOS of the UK. In unsteady flows of rarefied gas, an ensemble average process is essential to reduce the fluctuation of macroscopic quantities after a number of independent DSMC routines are executed. Therefore, in this paper the ensemble average process is paralleled. The acceleration rate of the parallel processing by the present scheme and system is 2.995, and the efficiency of the process is 0.998. Hence, the calculation time for simple unsteady flows of rarefied gas can reduce to 1/2.995, as compared with the time by a sequential scheme., 資料番号: NALSP0014006, レポート番号: NAL SP-14

A new array data assignment technique, to be used in CFD (Computational Fluid Dynamics)simulations on massively parallel processor is presented. The implenentation was done for an H2P parallel processor. H2P is an MIMD (Multiple-Instructions Multiple-Data stream) type of parallel processor in which every processing element (PE) has a local memory (i.e. a distributed-memory parallel processor) and a vector facility. The main contribution of this paper is in proposing an original data assignment technique which leads to an efficient use of the vector facilities. The proposed data assignment technique uses a combination of two methods (referred to as SWAP and Wave Front in the paper) in order to achieve its purpose, an efficient use of the vector facilities. The original CFD program is modified into a parallelized one, by applying the proposed methods. It is shown that, in running this modified program, H2P (in a 512 PEs configuration) can achieve a 180 times speedup over the original CFD program running on a uniprocessor (using the same type of PE)., 資料番号: NALSP0013008, レポート番号: NAL SP-13

The progress of computational aerodynamics over the past decade has been so tremendous, partly pushed by the rapid advance in computer hardware, that aircraft designers have obtained an alternative tool for aerodynamic analysis comparable to wind tunnel testing. The Navier-Stokes computation, which has been used successfully in research recently, is also likely to become a practical tool for analyzing viscous flows. NSFOIL, the high Reynolds number transonic airfoil analysis code developed at NAL, has been verified by comparing it with the results of high Reynolds number wind tunnel tests at NAL. This paper presents the NSFOIL computation of the transonic airfoil with aileron. following the clean airfoil result presented at the previous symposium in 1984. To show the effectiveness of this new code, the comparison of numerical results with the experiments is also presented in this paper., 資料番号: NALSP0005029, レポート番号: NAL SP-5

A development of a large launch vehicle (H-II) started in 1985. The first stage of this lauch vehicle uses a liquid oxygen (LOX)/hydrogen (LH2) engine (LE-7), which generates 1,180 KN in thrust. National Aerospace Laboratory (NAL) takes charge of the LOX turbopump development till the prototype engine is completed. The LOX turbopump consists of a main pump, a preburner pump, and a single stage impulse turbine. The design rotational speed is 20,000 rpm, and the delivery pressure of the main pump is 20.3 MPa and the preburner pressure is 11.4 MPa, and the flow rate is 229.1 kg/s. The high rotational speed, large low rate, and low suction pressure of the main pump require the combination of a cavitating inducer with three swept back helical blades and a main impeller. In order to design a high pressure and a high efficiency pump, it is necessary to analyze the pump blade passage flows in detail, using a three dimensional flow analysis method. This paper describes the design method and the experiment, which was carried out using liquid nitrogen and liquid oxygen at the turbopump test facility of the NAL Kakuda Branch. The design method in this paper combines a streamline analysis and a blade-to-blade analysis. The predicted pump head coefficients are in reasonable agreement with the experimental results considering the pump efficiency. Because a large amount of input data are usually necessary to calculate three dimensional flows, a method of calculating blade shapes correctly only by giving two blade angles at the impeller inlet and outlet is developed, 資料番号: NALSP0008016, レポート番号: NAL SP-8

Many high speed image processors have been developed to date. In this paper, these processors are classified from the view point of parallel implementation of identical processor elements and pipeline implementation with many types of processor elements. These processors are designed not only for high speed processing capabilities but also for compactness in hardware size. An image processor TIP-1 (Template-controlled Image Processor-1) was proposed and has been developed as progressive architecture for pipeline mechanisms. In this paper TIP-1 architecture is presented and an advanced system, TIP-3, using VLS1 chips, is proposed., 資料番号: NALSP0003021, レポート番号: NAL SP-3

Explicit and implicit time-marching methods for solving the three-diemsional compressible Euler equations developed by the authors are presented in this paper. The main features of the present methods are employment of the Eular equations of con-travariant velocities; that is, the equations of motion in the computational space, and simple and accurate treatment of the solid wall boundary conditions. In the present paper all the equations required in the computation are shown and compared with the existing methods., 資料番号: NALSP0007032, レポート番号: NAL SP-7

In order to examine the flowfield around ATP by numerical calculations, the finite difference methods are applied. In the first part of this paper, the partial differential equation for the disturbance velocity potential is solved by the line relaxation technique. In order to do the calculations, boundary fitted meshes are generated. The transforma-tions between physical and computational space are performed by chain rule. The SR-3 is used for numerical calculations. In the second part of the paper, the Euler equations are solved by non-iterative implicit ADI schemes in AF alogrithm. Boundary fitted meshes are also used for this case. The results show qualitative agreement with the results obtained by Bober et al. Finally, the aerodynamic performances, such as power co-efficients and efficiencies, are obtained. Potential calculations show that these quantities are relative]y larger than the experimental values obtained by NASA; however, Euler solutions show that the values come closer to the experimental values., 資料番号: NALSP0007008, レポート番号: NAL SP-7

To use a vector processor efficiently, the most important point is to increase the vector processing ratio of a problem. But, through the process of coding the problem, it is usual to decrease the parallelism of the problem. And the lost parallelism cannot be restored to the former level. The reason for this fact is that the parallelism of the problem depends largely on the algorithm and programming language used. This paper describes the process of coding and the cause of the decreasing parallelism. The selection of the programming language is the key point in the effective use of vector processors. Various kinds of language systems have been developed and are used for vector processors. Most of these language systems are based on standard FORTRAN, since FORTRAN is the main programming language used in the scientific and engineering fields. These language systems can be classified as follows: (1) Procedure oriented programming language i) Standard FORTRAN - Automatic vectorizing compiler method - Building block method ii) Extended FORTRAN - Vector (Array) processing language - Architecture oriented language (2) Problem oriented programming language In this paper, we present the characteristics of these language systems with examples, and discuss the merits and demerits of each., 資料番号: NALSP0001004, レポート番号: NAL SP-1

This paper treats the Parabolized Navier-Stokes equations. We previously solved the Euler equations by finite volume method. The results showed good agreement with experimental results. Then, this was extended to include viscous effect by solving the Navier-Stokes equations, using Beam-Warming scheme. These results also showed good agreement with the Euler code. we also could obtain thermal effect on the Shuttle-like body. In this paper, we employed the PNS equations in order to reduce CPU time. Actually, the calculation time was reduced to some thousands as much as the thin layer Navier-Stokes equations. The two cases were calculated. One is hemisphere cylinder, and another is hemisphere-eliptic cylinder at an attack angle. The first case was compared with the time-dependent thin-layer NS equations. The comparison between them was good. The second case showed clear effect of attack angle such as the turn of flow from the lower surface to the upper surface., 資料番号: NALSP0008006, レポート番号: NAL SP-8

An effective finite-difference scheme for solving full compressible Navier-Stokes equations was initated by Beam and Warming. The purpose of this paper is to develop that technique and apply it to the calculation of a typical subsonic or transonic, inviscid or viscous steady flow. We almost identically follow the Beam-Warming technique that may be summarized into the following threee characteristics: a) Delta-form approximate factorization algorithm, b) Implicit three-level scheme. c) The cross derivative viscous terms were explicitly replaced by the values of previous time. First, the strong conservation-law form of the Navier-Stokes equations is written in Cartesian coordinates, and then transformed into a general grid system. Following the Beam-Warming difference scheme, we generally adopt the parameter combination delta, An effective finite-difference scheme for solving full compressible Navier-Stokes equations was initated by Beam and Warming. The purpose of this paper is to develop that technique and apply it to the calculation of a typical subsonic or transonic, inviscid or viscous steady flow. We almost identically follow the Beam-Warming technique that may be summarized into the following threee characteristics: a) Delta-form approximate factorization algorithm, b) Implicit three-level scheme. c) The cross derivative viscous terms were explicitly replaced by the values of previous time. First, the strong conservation-law form of the Navier-Stokes equations is written in Cartesian coordinates, and then transformed into a general grid system. Following the Beam-Warming difference scheme, we generally adopt the parameter combination delta = l/2 and theta = 1. After introducing spatial factorization, second-order dissipative terms are added to the left-hand side, while fourth-order terms are added to the right-hand side; the fourth-order smoothing terms are changed to second-order terms at points adjacent to the boundaries. The numerical computations were carried out only for a typical NACA 0012 airfoil. The grid system used is a C-type, where 51 points are distributed uniformly over the airfoil surface with identical arc length. The lines of constant xi consist of two parts: One part of those emanating from the airfoil sulface points is the solution of the Laplace equation and the others are parabolas. The distribution of points on wake cut and in the eta-direction is exponential. The calculations start from uniform free-stream variables throughout the flow field. The boundary conditions in the far field are free-stream. The following four cases are calculated:a) M infinity =0.63 and alpha=0, causing an entirely subsonic flow. b) M infinity =0.63 and alpha=2 degrees, the highest velocity of upper surface being close to sonic c) M infinity =0.75 and alpha=0, when supersonic regions appear but there are no shock waves. The flow is supercritical. d) M infinity =0.75 and alpha=2 degrees, generating a shock wave over the upper flow field. The flow is transonic in the case of inviscid calculations. Viscous calculations, on the other hand, show no shock waves, while flow separation on the upper surface caused by the angle of attack is conspicuous., 資料番号: NALSP0003027, レポート番号: NAL SP-3

A method for generation of stractural 3-D body surface grids is described in this paper. The present grid generation procedure is based on the use of the interporation function of MASTER DIMENSION SYSTEM, which is used to define the wind tunnel test model shape, and which generates body surface grid automatically satisfying several geometrical requirements. The advantage of this method is that we can improve the efficiency of grid generation work and make grids quickly corresponding to the change of wind tunnel test model shape. The typical example of grid around a wing-body configuration generated by this method and the result of Navier-Stokes analysis are also presented in this paper., 資料番号: NALSP0010029, レポート番号: NAL SP-10

In this paper, numerical simulations of aerodynamic response to changes of the deflection angle of a longitudinal control surface of NAL experimental supersonic airplanes are discussed. To treat the change of the deflection angle of the control surface efficiently, an unstructured dynamic mesh method with surface grid movement is used. Computations of a rocket powered experimental supersonic airplane and a jet powered experimental supersonic airplane are performed at the supersonic cruising Mach numbers (M(sub infinity) = 2.0 and 1.7) and a transonic Mach number (M(sub infinity) = 0.95). The computational results are compared and are discussed., 資料番号: AA0045948052, レポート番号: NAL SP-57

This paper describes the concept and design of virtual computing environment project ITBL. The goal of ITBL project is the mutual utilization of computing-resources among multiple organizations and the convenient environment for users to access resources or to develop user application. For this purpose, ITBL project is not based on only GRID technology. To realize the concept, several infrastructure models are proposed. Three independence technologies ITBL-VPN, PitSaw and ITBL portal were developed. At present these developments are almost finished and ITBL experimental system works will be finished shortly. A further direction of this study will be to evaluate these implementations., 資料番号: AA0045948010, レポート番号: NAL SP-57

The unsteady flow phenomenon in turbomachinery changes drastically with changing operating condition. It is important that the various flow phenomena are predicted accurately from unsteady simulations of complicated flow field in blade rows. The period and scale of unsteady phenomena should be known for simulation in advance. In this study, the visualization system, which can visualize the unsteady flow phenomena during computation, has been applied to analyzing turbine flow field. The system which was conducted for multi-block parallel computing visualized effectively the unsteady flow behavior inside a turbine stage such as Karman vortex shedding, nozzle-rotor interaction, periodic separation occurrence on the rotor blade, etc. In this paper, the advantage of the visualization approach is clarified., 資料番号: AA0045948040, レポート番号: NAL SP-57

According to the recent rapid progress of high performance computing and numerical algorithm, a CFD-based aerodynamic design methodology is becoming an important tool with wind tunnel testing. The 3rd Research Center, TRDI-JDA has implemented the development and performance tests a CFD-based design system CASPER (Computational Aerodynamics System for Performance Evaluation and Research), with major domestic aircraft manufactures until March 2001. In this paper, mainly shown is part of aerodynamic and aeroelastic computed results obtained in performance tests. Actually, for the code validation, various types of aircraft and airfoil configurations are computed using the CASPER and compared well to wind tunnel testing data and other computed results. In addition, as an example of CFD multidisciplinary simulation providing the CFD data obtained in its aerodynamic characteristic module, the aeroelastic simulation can be normally performed. The CASPER can be expected to be powerful tool for conceptual and basic design stage of advanced aircraft at the next generation., 資料番号: AA0045948021, レポート番号: NAL SP-57

As a speed of computational server has been growing, a huge scale computation, e.g. combustion chemical reaction flow, can be done. While I/O processing has become the bottleneck of numerical simulation. So in this paper, the design of storage system that does not have the I/O bottleneck is discussed, even when a huge scale numerical simulation is done. Studied is the storage system design of next generation, supercomputer system at National Aerospace Laboratory of Japan (Numerical Simulator 3). Using survey of calculation data size for actual computational fluid dynamics research, performance of speed and capacity were designed. Refereeing the simplified IEEE OSSI reference model, acceptable condition of operation was studied. As a result of these studies, high speed mass storage system was designed, which based upon computational server's local file system and which cooperates with disk and tape parallel I/O, Hierarchical Storage Management, GSN, and NFS., 資料番号: AA0045948004, レポート番号: NAL SP-57

In this paper, computations are carried out in transonic and high Reynolds number flows around ONERA Model M5 configuration using the multiblock method together with Baldwin-Lomax model. The present computed results are quantitatively compared with wind-tunnel experiments in terms of wing surface pressure distributions, C(sub L)-alpha, C(sub L)-C(sub D) and C(sub M)-C(sub L) curves. The present numerical solutions agree with experiments well, except at Re = 60.0 x 10(exp 6)., 資料番号: AA0045948012, レポート番号: NAL SP-57

This paper presents comparison of EFD/CFD results submitted to the 3rd SST-WorkShop held in December 2001. Comparison among CFD results shows mostly good agreement. However, it shows big discrepancy in C(sub D) estimation. Comparison between CFD and EFD shows mostly good agreement except in several aspects. A big discrepancy of C(sub p) distributions can be seen at the leading edge on the upper surface as well as at the outer wing. The comparison also shows a small difference in C(sub L-alpha) and C(sub M-alpha), indicating good agreement in most of C(sub L) and C(sub D) estimation. Sensitivity analysis for free stream Mach number and attack angle, as well as dependency analysis for the limiter functions of numerical scheme have been conducted as a verification of CFD, indicating that the discrepancies between EFD and CFD can not be resolved by these considerations alone., 資料番号: AA0045948015, レポート番号: NAL SP-57

Recently, CFD analyses have been extensively applied to noise reduction technologies of a supersonic transport engine. This paper introduces the CFD activities for noise reduction technologies. For fan noise, the magnitude of unsteady pressure on the 3D stator vane evaluated through a use of unsteady 3D stage analysis. For jet noise, Large Eddy Simulation (LES) is applied to prediction of exhaust jet mixing of a lobed mixer nozzle. It is confirmed that a CFD application takes an important role in noise reduction technologies., 資料番号: AA0045948041, レポート番号: NAL SP-57

In this paper, a generalized interpolation-based lattice Boltzmann method (GILBM) is introduced. The standard lattice Boltzmann equation is transformed into generalized coordinates using metrics. The advection term of the transformed equation is discretized with sufficient spatial accuracy to suppress the numerical dissipation. Additionally, a wall boundary condition suitable for body-fitted grids and a non-reflecting outer boundary condition are introduced. With the present method, the flow around a complex geometry can be solved using structured grids. The three examples of, the cavity flow, the flow around a circular cylinder, and the flow around an airfoil are shown as validation of the method to a wide range of Reynolds number. The present results show good quantitative and qualitative agreement with other computational studies., 資料番号: AA0045948047, レポート番号: NAL SP-57

ITBL (Information Technology-Based Laboratory) project has been started in the fiscal year 2000, aiming at the realization of virtual research laboratories on the high speed network connecting the computer resources such as supercomputers, graphical workstations, databases and so on. In the project, National Aerospace Laboratory of Japan is in charge of developing a collaborative simulation system for aerospace engineering. NAL is constructing three kinds of software as ITBL applications, UPACS (Unified Platform for Aerospace Computational Simulation) - CFD environment for aerospace applications, distributed analysis server, and database for CFD validation. In this paper, the usefulness of ITBL environment for aerospace community and the three applications NAL is developing are explained., 資料番号: AA0045948011, レポート番号: NAL SP-57

Computational analysis of High Speed Flight Demonstrator (HSFD) Phase 2 Vehicle are presented. HSFD Phase 2 is the flight experiment to reduce uncertainties in the HOPE-X transonic aerodynamics. HSFD vehicle is 25 percent scale model of HOPE-X. The phase 2 flight experiment will be scheduled at the Swedish Space Corporation's Esrange site in Sweden in 2003 and conducted in collaboration with Centre National d'Etudes Spatiales (CNES) of France. In this paper, multi-block Navier-Stokes simulations have been performed for the HSFD model and total aerodynamic performance including base regions has been investigated. Especially, study of the base flow effects is focused on for the accurate prediction of total aerodynamic characteristics. Computational results are compared with the experimental data of NAL transonic wind tunnel at M(sub infinity) = 0.8, 1.05 and 1.2., 資料番号: AA0045948026, レポート番号: NAL SP-57

This paper presents software development environment and user interface of Numerical Simulator 3 (NS3). NS3 has been well designed to become a user friendly system, which has standard and open features. A computational engine of NS3 is called Central Numerical Simulation System (CeNSS), which is an SMP cluster type super computer, capable of not only process parallelization but also thread parallelization. XPFortran and MPI are used for process parallelization whereas OpenMP and automatic parallelization are used for thread parallelization. Moreover, it is possible to run parallel programs developed at NWT on CeNSS by mapping the concept of vectorization to that of thread parallelization., 資料番号: AA0045948005, レポート番号: NAL SP-57

This paper shows the results of the elementary wind tunnel test of blade active control using a 1-bladed model rotor at the 2.5 m x 2.5 m multi-purpose wind tunnel in Kawada Industry. In this test, the effects of higher harmonic control (HHC) and active flap control (AFC) on rotor blade-vortex interaction (BVI) noise were investigated. The frequency of the control was 20 Hz (2/rev), and the amplitudes of HHC and AFC were +/- 2.0 deg and +/- 15.0 deg, respectively. The phase of control was changed at intervals of 45 deg. Two types of BVI were tested. The one is the parallel vortex interaction using the vortex generator upstream of a non-lifting rotor. The other is self-generated vortex interaction of a lifting rotor. As a result, the effects of BVI noise reduction were observed in the some cases., 資料番号: AA0045946010, レポート番号: NAL SP-55

Various types of vortex streets are formed behind a circular cylinder oscillated streamwise in a uniform flow. The vortex streets behind the cylinder were classified into seven types including Net-type and P-type vortex streets in the previous paper. These types of the vortex-streets appear in the lock-in range referring to the frequency of the oscillating cylinder and that of vortex shedding. The Net-type vortex streets have a peculiar but a stable 3D vortex structure. While P-type vortex streets have a stable 2D vortex structure. In this report the formation processes of the vortex streets in the lock-in range were studied by mean of flow visualization experiments. By comparing the flow patterns in the near wake to form Net type and P-type vortex streets, the mechanism how to form the two different types vortex streets in a lock-in range is discussed., 資料番号: AA0045947005, レポート番号: NAL SP-56

National Aerospace Laboratory of Japan has been conducting researches on the aviation safety and environmental compatibility technology (ASET) since 1994. In 2001 a new research plan of the ASET was started toward next four years (2001-2004). This paper describes the current outline of the ASET research activity., 資料番号: AA0045946001, レポート番号: NAL SP-55

Blade-Vortex Interaction (BVI) noise is generated by interaction between blade and tip vortex. In this paper, BVI is simulated as an interaction between a blade and an externally generated vortex from a fixed-wing vortex generator. This vortex was modeled by the Scully model. Euler/FW-H method and Euler/Kirchhof method are used for noise prediction methods. Experiments were conducted in NAL and compared with calculation results by above noise prediction methods., 資料番号: AA0032819033, レポート番号: NAL SP-53

UPACS (Unified Platform for Aerospace Computational Simulation) is a common CFD platform program developed at National Aerospace Laboratory. The project to develop UPACS aims not only to overcome the increasing difficulty in recent CFD code programming on parallel computers, but also to accelerate the development of CFD technology by sharing a common base code among research scientists and engineers. This paper reports the extension to a rotary coordinate system of UPACS and computational results of the flow around transonic compressor rotors, NASA rotor 37 and rotor 35, which have the number of published computational and experimental works., 資料番号: AA0032819014, レポート番号: NAL SP-53