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The SST (Shear-Stress Transport) turbulence model has been implemented in UPACS, a CFD platform developed in JAXA. Four test cases with increasing level of difficulty were used to validate the present implementation. Three basic test cases included flat plate, backward-facing step, and transonic airfoil RAE2822. For more practical application, the flow simulation around the DLR-F6 transport aircraft was performed. Validations were carried out by comparing the present numerical results with those of other numerical methods and experimental method. Numerical results obtained from the computations above show that the present implementation is valid. We believe that this effort has extended UPACS's applicability to more complicated flow fields., 資料番号: AA0049212001, レポート番号: JAXA-SP-05-017

Development of a next-generation airplane cruising near sonic speed received considerable attention in 2002. It requires a better understanding of corresponding aerodynamic characteristics and CFD validity. As a part of the collaboration between NAL and MHI on transonic wind tunnel tests and CFD computations using NEXST-1 supersonic transport model, the effect of wall interference in NAL 2 m x 2 m transonic wind tunnel at near sonic speed condition was investigated with CFD based on Reynolds-averaged Navier-Stokes equations. The results showed considerable drag reduction (dip) at Mach number just below the drag divergence Mach number due to the interaction of wall and sting support., 資料番号: AA0048062003, レポート番号: JAXA-SP-04-009

Development of a next-generation airplane cruising near sonic speed received considerable attention in 2002. It requires a better understanding of corresponding aerodynamic characteristics and CFD validity. As a part of the collaboration between NAL and MHI on transonic wind tunnel tests and CFD computations using NEXST-1 supersonic transport model, the effect of wall interference in NAL 2 m x 2 m transonic wind tunnel at near sonic speed condition was investigated with CFD based on Reynolds-averaged Navier-Stokes equations. The results showed considerable drag reduction (dip) at Mach number just below the drag divergence Mach number due to the interaction of wall and sting support., 資料番号: AA0047427042, レポート番号: JAXA-SP-03-002

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

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, 無人往還機HOPE-X(宇宙往還技術機)の遷音速空力特性を調べるための高速飛行実証プログラムを計画した。この計画で使用される飛行実証機には、機首先端にADS(空気データシステム)のためのピトーブームが取り付けられた。この計画の準備として、ピトーブームの空力的影響を評価するために、この飛行実証機周囲の遷音速流の数値シミュレーションを行った。ここで使用したCFD(計算流体力学)コードは風上TVD(全変動減少)スキームを適用したFVM法(有限体積法)を使って差分化した、レイノルズ数平均化ナビエ-ストークス方程式に基づいている。ここで使用した計算格子は、4個のブロックからなるマルチブロック構造格子で、格子点の総数は約60万である。CFDシミュレーションの結果はピトーブームの空力的影響が非常に小さいことを示している。, A high-speed flight demonstration program has been planned to investigate transonic aerodynamic characteristics of the HOPE-X (H-2 Orbiting Plane-Experimental) unmanned re-entry vehicle. The flight demonstrator used in this program has a Pitot-boom for ADS (Air Data System) on its nose-tip. As a preparation for this program, numerical simulations of transonic flows about this flight demonstrator were executed in order to estimate the aerodynamic effect of the Pitot-boom. The CFD (Computational Fluid Dynamics) code used here is based on Reynolds averaged Navier-Stokes equations discretized using an FVM (Finite Volume Method) with a TVD (Total Variation Diminishing) upwind scheme. The computational grids used here are multi-block structured ones consisting of four blocks. Total number of the grid-points is about 0.6 millions. The CFD results show that the aerodynamic effect of Pitot-boom is very small., 資料番号: AA0028635035, 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, 新しいHOPE(H-2ロケット打上げ型有翼回収機)11モデルの空力特性に関する数値解析を行った。本研究では総計約300例のCFD(計算流体力学)計算を実行した。亜/遷/超音速領域についてマッハ数0.4から3.0、迎角0度から20度、そして横すべり効果がある場合とない場合についてパラメトリックなマルチブロックのナビエ-ストークス(N-S)解析を実施した。遷音速流に関しては、マッハ数0.9と1.1でスティング干渉効果についても調べた。打ち上げ時の低超音速で低迎角の場合、乱流加熱のあることが予言されているが、数値計算の結果をAEDC(アーノルド技術開発センター)-HWT(極超音速風洞)実験と比較した。超音速再突入時の基本的な空力特性、空熱力学特性も、マッハ数0.8から26の範囲で、迎角30度、40度、50度の場合について調べた。この際、ボディフラップおよびエレボンの偏向などの舵面効き特性の検討を行った。実在気体効果は7化学種の非平衡1温度N-Sコードを使って解析した。全体的空力特性に対する化学反応の寄与を分割するため、3種類の異る気体モデル(完全気体、凍結流、実在気体モデル)を使った計算を実施した。これらのパラメトリック計算を通して、新しいHOPE11基本設計モデルの全体的な空力特性、空力加熱特性を評価した。, Numerical study of aerodynamic characteristics of new HOPE (H-2 Orbiting Plane) 11 model is performed. In the present study, totally about 300 CFD (Computational Fluid Dynamics) calculations are carried out. For subsonic, transonic and supersonic speed range, parametric multi-block Navier-Stokes (N-S) analysis has been conducted from M(sub infinity) = 0.4 to 3.0 at angles of attack from 0 to 20 deg, with and without side slip effects. For transonic flows, sting interference effects are also investigated at M(sub infinity) = 0.9 and 1.1. Turbulent heating at low hypersonic speed in launching phase is predicted at low angles of attack and the numerical results are compared with AEDC (Arnolds Engineering Development Center) HWT (Hypersonic Wind Tunnel) experiments. Basic aero- and aerothermodynamic characteristics at hypersonic re-entry phase are also investigated from M(sub infinity) = 0.8 to M(sub infinity) = 26 at angles of attack alpha = 30 deg, 40 deg, and 50 deg, where the efficiency of control surface such as body flaps and elevon deflections is also studied. Real gas effects are analyzed by using a seven species chemically non-equilibrium one temperature N-S code. In order to divide contributions of chemical reactions to total aerodynamic performance, calculations using three different kind of gas models (perfect, frozen and real gas model) have been made. Through these parametric computations, total aero- and aerothermodynamic characteristics of new HOPE 11 basic design model are evaluated., 資料番号: AA0028635036, 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, 宇宙往還技術試験機の空力設計を支援するため、数々のHOPE-X(宇宙往還技術試験機)に関する流れのシミュレーションが行われている。空力特性を詳細に調べるために、エルボンが20度下げの状態におけるHope-Xまわりの遷音速流の予備的な解析を行った。この計算では、TVD(全変動減少)風上差分スキームによるFVM(有限体積法)を使って差分化したレイノルズ平均ナビエ・ストークス方程式を基にしたCFDコードを用いた。計算に用いた計算格子は、42ブロックからなるマルチブロック格子である。全格子点の数は、約370万であった。, A large number of flow simulations about HOPE-X (H-2 Orbiting Plane Experimental) has been carried out to support its aerodynamic design. In order to investigate the aerodynamic characteristics in detail, preliminary computations of transonic flows about HOPE-X with -20-degree steered elevons were executed. The CFD (Computational Fluid Dynamics) code used here is based on Reynolds averaged Navier-Stokes equations discretely divided using FVM (Finite Volume Method) method with TVD (Total Variation Diminishing) upwind scheme. The computational grid used here is multi block structured one consisting of 42 blocks. Total number of the grid-points is about 3.7 millions., 資料番号: AA0001961024, レポート番号: NAL SP-44

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, ボーイング747-200機の翼周りの3次元Navier-Stokesシミュレーションを、航空宇宙技術研究所の並列ベクトル計算機NWT(数値風洞)を用いて実行した。本研究の主目的は翼端渦の数値シミュレーションである。本システムの支配方程式は３次元Reynolds平均薄層Navier-Stokes方程式で、TVD風上スキームによる有限体積法を用いて解く。翼周りのC-O型格子は24のサブドメインに均等分割する。サブドメイン毎のそれぞれのPE(要素計算機)に計算負荷を分割配分し並列演算を実行する。1100万の格子点を用いた翼端に発生する渦の格子形成を示した。, Three dimensional Navier-Stokes simulations about the wing of the Boeing 747-200 have been carried out using a parallel vector computer called NWT (Numerical Wind Tunnel) at the National Aerospace Laboratory, Japan. The main objective of this study is to simulate the wing-tip vortex. The governing equations are three dimensional Reynolds-averaged thin-layer Navier-Stokes equations which are discretized by finite volume method with a TVD (Total Variation Diminishing) upwind scheme. The C-O type of grid around the wing has been decomposed equally into 24 subdomains. Parallelization has been done by means of sharing the load of computation for each subdomain to each PE (Processor Element). Computation using a grid with 11-million grid-points shows an existence of a strong tip vortex induced by the wing., 資料番号: AA0000294010, レポート番号: NAL SP-30

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, 完全航空機形状周りの3次元Navier-Stokesシミュレーションを航空宇宙技術研究所の並列ベクトル計算機NWT(数値風洞)を用いて実行した。本システムの支配方程式は３次元Reynolds平均薄層Navier-Stokes方程式で、TVD風上法による有限体積法を用いて解いた。格子点分配法(point charge distribution method)を用いて、完全航空機形態周りのマルチドメイン構造格子による格子形成を行った。, A three-dimensional Navier-Stokes simulation about a complete aircraft configuration has been carried out using the parallel vector computer called NWT (Numerical Wind Tunnel) at the National Aerospace Laboratory, Japan. The governing equations are three dimensional Reynolds-averaged thin-layer Navier-Stokes equations which are discretized by finite volume method with a TVD (Total Variable Diminishing) upwind scheme. The multidomain structured grid around the complete aircraft configuration has been generated by the point-charge distribution method., 資料番号: AA0000294009, レポート番号: NAL SP-30

｢数値風洞(NWT)｣と呼ばれる航空宇宙技術研究所の並列ベクトル計算機を用いてONERA-M5模型のまわりの遷音速流れの3次元ナビエ･ストークス計算を行った。計算プログラムを2つの方法により並列化した。物体のまわりの計算空間を14のブロックに分割し、第1にそれぞれのブロックを1つの要素プロセッサに割り当てる。第2に、それぞれのブロック内で並列化を行う。並列化の効果は、ナビエ･ストークス方程式を解くためのCPU時間を計ることにより評価した。計算結果の評価も併せて行った。, Three dimensional Navier-Stokes computation of transonic flows about ONERA-M5 (Office National d'Etudes et d'Recherches Aerospatials-M5) has been carried out using the NAL (National Aerospace Laboratory) parallel vector computer called 'NWT' (Numerical Wind Tunnel). The computational space around ONERA-M5 is decomposed into 14 blocks. The code is parallelized in two methods. The first one is sharing the computation for each block to each Processor Element (PE). The second one is combining the first one and the parallelization in each block. The performance of the parallelization is estimated by measuring the cpu-time for solving the Navier-Stokes equations. Validation of the computational results has also been done., 資料番号: AA0004174034, レポート番号: NAL SP-27

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

極超音速実験機のマッハ5空力特性を0.5mJAXA極超音速風洞で取得した。機体模型の6分力計測を行うとともに、実験機の飛行環境を摸擬したCFD(Computational Fluid Dynamics)解析を実施した。結果として、迎角に対する空力係数の変化は風洞試験とCFDで概ね一致した。風洞試験においては、機体上面に大規模な剥離域が観察された。風洞試験とCFDにおける空力係数の違いは、剥離域の違いの影響と推測された。エンジンを装着した形状では、法線力係数と軸力係数が上昇した。, Mach 5 aerodynamic coefficients for a hypersonic experimental airplane are examined at JAXA 0.5 m hypersonic wind tunnel. Six DOF (Degree Of Freedom) forces for the model are measured and evaluated. CFD (Computational Fluid Dynamics) analyses for the same airframe are performed simulating real flight condition for the experimental airplane. As a result, variations of aerodynamic coefficients on angle of attack for both wind tunnel test and CFD analysis are consistent in general. Large separation region is observed on the upper surface of the body at the wind tunnel test. It is presumed that the difference of aerodynamic coefficients has occurred by the difference in flow separation between the wind tunnel test and CFD. Lateral force coefficient and axial force coefficient are raised with engine installed configuration., 資料番号: AA0063912000, レポート番号: JAXA-RR-07-047