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Parallel blade-vortex interactions have been calculated using the finite difference lattice Boltzmann method of the compressible Euler model. The perturbed discrete Boltzmann equation based on a prescribed vortex method has been proposed in order to prevent a vortex from diffusing by numerical dissipation. The discretization of the governing equation is based on a second order accurate explicit Runge-Kutta time integration and a fifth order accurate upwind scheme which includes additional terms to capture shock waves clearly. Transonic flow around an airfoil without vortexes has been simulated to validate the perturbed discrete Boltzmann equation system. A surface pressure distribution and pressure contour lines around the airfoil have been compared with other numerical data, and good agreements have been obtained. As a simple model of parallel blade-vortex interaction, two-dimensional blade-vortex interaction has been calculated using the proposed numerical method. An instantaneous pressure coefficient, a time history of a lift coefficient and patterns of acoustic waves have been compared with other numerical results, and agreed with them very well. Mechanism of noise generation has been also captured from numerical results. Three-dimensional calculations of parallel blade-vortex interaction have been performed using the present numerical procedure. Time variations of surface pressure distributions have been compared with Euler calculation and experimental data, and good agreements have been obtained., 資料番号: AA0063742007, レポート番号: JAXA-SP-07-016

An interest in Busemann biplane has been revived and the feasibility of the supersonic biplane is now under investigation by many researchers. However, these studies are based mainly on CFD (Computational Fluid Dynamics) calculations and not validated by experiment so that a discussion on the wave cancellation mechanism between the two wings is very limited. The objective of this research is to measure the wing surface pressure distribution on a Busemann biplane and to study the two-dimensional interference phenomenon between the two wings. Since the Busemann biplane wings are so thin (the wing thickness ratio is 0.05) that it is difficult to install static pressure taps on the model. In this experiment, we used Pressure-Sensitive Paint (PSP) to measure surface pressure distributions on the model. PSP is an optical pressure measurement technique based on photochemical reaction called 'oxygen quenching'. This is a coating type sensor and considered only means to measure pressure fields on a thin airfoil model. In this study, wind tunnel experiment was conducted in a 60 mm x 60 mm indraft supersonic wind tunnel. The surface pressure distribution obtained using PSP were compared with CFD results and the Schlieren images. The complicated three-dimensional interference phenomenon between the two wings, including shock wave/boundary layer interactions was analyzed. On the basis of these experimental results, the wave cancellation mechanism of the Busemann biplane is discussed., 資料番号: AA0063742004, レポート番号: JAXA-SP-07-016

Investigation of a new high-speed zero-emission transportation 'Aerotrain' is proceeded mainly in Tohoku University and Miyazaki University. Since the aerotrain utilizes the ground effect, researches of the aerofoil section which can harnesses the ground effect effectively are important. The aerotrain moves along U-shaped guide way which has a ground and side walls, so it has many viscous interference elements. We analyze the flow through a primary and secondary aerofoils, to prevent the boundary layer separation for the improvement of the aerodynamic characteristics at low speed and near the ground. A small secondary aerofoil is equipped above the trailing edge to increase lift at takeoff and landing. We find out the most suitable location of secondary aerofoil by the boundary layer approximation analysis, and experiment using towing wind tunnel at the Sunrise-Beach Research Facility in Miyazaki., 資料番号: AA0063908015, レポート番号: JAXA-SP-07-026E

In this study, CFD (Computational Fluid Dynamics) analysis is performed on a JAXA 3D High-lift Configuration. The main object is to assess and improve the reliability in simulating the flow around High-Lift configuration. The lift prediction shows qualitatively good agreement with the experiment, but the accuracy of lift prediction has the matter on account of prediction around maximum lift and lift slope. Drag is estimated lager than the experiment consistently. Applying UMUSCL (Unstructured Monotone Upstream-centered Schemes for Conservation Laws), computational results show improvement in lift and drag predictions as result of spurious drag decrease. Increasing the mesh density, the large separation on flap and both lift and drag decrease are computed. The scale of separation of flap is sensitive to (especially surface) mesh resolution around flap, and affects wing circulation. Consequently, the pressure distribution is changed and predicted pitching moment is improved. Using SST (SuperSonic Transport) turbulence model, all aerodynamic coefficients come to the experimental data compared to the case of SA (Spalart-Allmaras) turbulence model. These results are also caused by the large separation on flap. The prediction of the separation on flap is important to accurate prediction of aerodynamic forces., 資料番号: AA0063742030, レポート番号: JAXA-SP-07-016

JAXAで基礎研究を進めている極超音速実験機(HYTEX: HYpersonic Turbojet EXperiment)の遷音速および超音速での空力特性を確認するため、10％スケール模型を用いて1m×1m超音速風洞(SWT1: Supersonic Wind Tunnel 1)および2m×2m遷音速風洞(TWT1: Transonic Wind Tunnel 1)において感圧塗料(PSP: Pressure Sensitive Paint)による圧力分布計測試験を行った。PSPはSWT1試験では複合PSPを用い、TWT1試験ではPSP(Pressure Sensitive Paint)/TSP(Temperature Sensitive Paint)塗り分けを用いた。SWT1およびTWT1で風洞試験を行い機体周りの詳細な圧力分布を取得した。本稿ではM=2.5、2.0、1.4、1.1、0.9、0.7でのPSP試験結果について、模型上面および下面での圧力分布とそれに関わる現象について考察を加える。また圧力孔計測データとの比較も行なった。, 資料番号: AA0063897005, レポート番号: JAXA-SP-07-019

An aerodynamics simulation code using high-order Spectral Volume (SV) method has been developed, and flow simulations around JAXA high-lift configuration model were conducted on the Earth Simulator. The obtained results of the Euler computation are compared with those of a conventional unstructured method and the available wind-tunnel data. The computed pressure profiles fairly agree with the experimental data on the relatively coarse mesh, although viscous effects were all neglected in the calculation. In particular, trailing vortices from various high-lift devices are clearly captured even in the down stream region where they are likely to be vanished due to inherent numerical viscosity in the conventional unstructured mesh methods. This demonstrates that the present high-order unstructured mesh method will be capable of capturing various flow features accurately while it retains the desired geometrical flexibility. It is also shown that the developed code achieves high computing performance on the Earth Simulator., 資料番号: AA0063742032, レポート番号: JAXA-SP-07-016

In this study, the flow characteristics of UT (University of Tokyo)-Kashiwa hypersonic wind tunnel are obtained by measuring the Pitot pressure distribution. In this wind tunnel, the Mach 7 and 9 hypersonic nozzles are inherited from its predecessor facility in UT-Komaba Campus. From a viewpoint of the operation safety and cost, the maximum stagnation pressure at Kashiwa facility is 950 kPa, which is much lower than at Komaba facility (5 MPa). To evaluate the uniformity of the freestream of this tunnel, the Pitot pressure traverse was conducted for the Mach 7 nozzle by using the rake of 11 Pitot tubes with 20 mm intervals in the radial direction. The measurement has been done at every 30 or 45 degrees in the circumferential direction and also at some streamwise locations. The Mach number calculated from the settling chamber pressure and the test-section static pressure is nearly seven, which is almost the same as the Mach number obtained from the Pitot pressure (about 14 kPa) on the center line. The Pitot traverse shows that the uniform flow core with 120 mm diameter is formed both at the nozzle exit (200 mm in diameter) and at the center of the test section. Such ratio of the core diameter to the nozzle exit diameter seems quite reasonable, comparing to other existing hypersonic wind tunnels. The cross-sectional Mach number distribution shows slight deviation from the axi-symmetric pattern because of the misalignment of the nozzle and the test section to the tunnel center line. Consequently, the highly uniform Mach 7 flow is obtained in this facility, though the nozzle is used under the off-design operation pressure, that is, the off-design Reynolds number., 資料番号: AA0063742009, レポート番号: JAXA-SP-07-016

Aerodynamics around unsteady flapping wings is analyzed by using CFD (Computational Fluid Dynamics) techniques. Two types of flapping wing are modeled and analyzed referring to the hovering motion of hawkmoth and honeybee. Multi-block technique is used to make the suitable calculation grids for wing and body, and the grid-overset technique is used for the interpolation of physical values between those grids. CFD results show several vortices are generated at the leading edge, tip, and the trailing edge of flapping wings which comprise the complex flow fields around the wings and a body. The analysis also clarified the leading edge vortex significantly contributes to the generation of lift. Subsequent acoustic analysis is conducted referring to the CFD results to simulate flapping sound of honeybee. Directivity of sound propagation is recognized showing strong sound propagation in the front direction. Consequently, many features of complex flow around the flapping wing and its sound generation are clarified both qualitatively and quantitatively by using CFD and acoustic techniques., 資料番号: AA0063742022, レポート番号: JAXA-SP-07-016

Pressure-Sensitive Paint (PSP) measurement system was developed in the JAXA 6.5 m x 5.5 m low-speed wind tunnel. Pressure images on a High-Lift-Device (HLD) model, which was built for R&D (Research and Development) for next generation of civil transport aircraft, were measured by the low-speed PSP system. The dependencies of pressure patterns on flow speed and angle of attack were investigated. Pressure distribution peculiar to the HLD model were clearly visualized. The measurement accuracy of the present low-speed PSP system was approximately 0.16-0.2 in Cp at 60 m/s. These results indicated that this PSP system was a practical measurement tool to acquire pressure images on an aerodynamic model at low-speed., 資料番号: AA0063607002, レポート番号: JAXA-SP-07-005

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ 航空宇宙技術研究所 20011203-20011205 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport National Aerospace Laboratory 20011203-20011205 Tokyo Japan, The high-speed wind tunnel tests of the NEXST-1 (National EXperimental Supersonic Transport-1) were conducted in 1999 at the 1 m x 1 m Supersonic Wind Tunnel of NAL. The main objectives of the tests were to confirm the CFD (Computational Fluid Dynamics)-and-inverse-method-based supersonic natural laminar flow wing design, to collect wind tunnel data to be correlated with CFD and flight test data of the NEXST-1 and to supply data for validation and further development of the CFD tools, other than to collect aerodynamic data to be used in the detailed system design phase of the NEXST-1. At M = 2.0 where the data were presented for the workshop, the force test data showed very smooth longitudinal characteristics. The pressure test results proved that the flat-type pressure distribution was achieved on the upper surface as was intended in the wing design, and the data also showed good repeatability and good agreement with ones of a larger model tested in ONERA., 資料番号: AA0063609057, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ 航空宇宙技術研究所 20011203-20011205 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport National Aerospace Laboratory 20011203-20011205 Tokyo Japan, This paper provides CFD (Computational Fluid Dynamics) flow analysis results of the NAL NEXST-1 (National EXperimental Supersonic Transport-1) for the 3rd NAL SST (SuperSonic Transport) CFD workshop. A hybrid unstructured grid system was employed for viscous aerodynamic analysis. Computations were conducted for both wind tunnel test condition and Flight Test condition. The CFD results showed very good correlation with wind tunnel test results., 資料番号: AA0063609059, レポート番号: JAXA-SP-06-029E

航空宇宙技術研究所 17-19 Jan. 2000 東京 日本, National Aerospace Laboratory 17-19 Jan. 2000 Tokyo Japan, 1997年、航空宇宙技術研究所(NAL)は、次世代超音速輸送機開発のための技術的基盤の確立を目的に、縮尺超音速実験機を開発する計画を開始した。2002年の初実験飛行に向けて無動力実験機の開発が進められている。実験機をロケットブースタにより打ち上げる。ロケットブースタを分離の後、実験機はマッハ数2.0、高度18.5kmで計測飛行を開始する。実験機の空力設計において、三菱重工業(株)(MHI)は全体設計を担当し、NALはその高度なCFD(計算流体力学)コードを用いて主に翼および翼-胴空力接合の設計を行った。また、MHIは打上げ形態の空力設計を担当した。MHIの設計作業においては、風洞試験データを主に使用した。実験データのいくつかを検討し、CFD解析を用いて修正した。さらに、実験機およびロケットブースタの複雑な流れ場の物理的現象を理解する上で、CFD解析が重要な役割を果たした。, In 1997, National Aerospace Laboratory (NAL) started a program to develop a scaled supersonic experimental airplane to establish the technological basis for the development of the next generation supersonic transport. A non-powered experimental airplane is being developed for the first experimental flight in 2002. The airplane is launched by a rocket-booster. After the rocket-booster is separated, the airplane starts the measurement flight at Mach 2.0 and altitude 18.5 km. In the aerodynamic design of the experimental airplane, Mitsubishi Heavy Industries, Ltd. (MHI) is in charge of the overall design and NAL mainly conducted the wing and wing-body aerodynamic integration design with their sophisticated CFD (Computational Fluid Dynamics) code. MHI is also in charge of the launch configuration aerodynamic design. In the MHI design work, wind tunnel test data were mainly used. Some experimental data were checked and corrected using the CFD analysis. CFD analysis also played an important role in understanding physical phenomena of the complicated flow field of the experimental airplane and the rocket-booster., 資料番号: AA0028638028, レポート番号: NAL SP-49T

SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design 航空宇宙技術研究所 19980316-19980317 東京 日本, SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design National Aerospace Laboratory 19980316-19980317 Tokyo Japan, 資料番号: AA0063609004, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, This paper describes the outline of the CFD (Computational Fluid Dynamics) analysis workshop on the NEXST-1 (National EXperimental Supersonic Transport-1) flight test, which was held on October 13, 2006. The NEXST-1 flight test was conducted to validate JAXA's CFD-based aerodynamic design concepts on a future supersonic transport aircraft. The comparison of the flight test data and CFD analysis results validated the design concepts qualitatively, but some quantitative differences were also found. To understand the true effects of the design concepts, those differences should be investigated in detail. As one of the best ways for such investigation, present CFD analysis workshop was planned and the validation of JAXA's CFD results was selected as a main theme. In this paper, first of all, several discussion points to be investigated were summarized. Then, analysis subjects are also described. Four parties finally participated in this workshop and compared their original CFD results with the flight test data to investigate the discussion points., 資料番号: AA0063609066, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, Software for aerospace vehicle surfaces is applied and further developed as a Parameterized Geometry Preprocessor suitable for aerodynamic design and optimization. Airfoil families are designed with knowledge based parameters and their variations are used for optimization strategies, as well as with spanwise parameter variations to shape arbitrary lifting wings. Integration with fuselage results in a one-block surface grid for wing-body configurations which is found practical for rapid prototype design. Generation of scalar distributions along surface gives target pressure functions to support the input of inverse design codes. Case studies derived from the NEXST (National EXperimental Supersonic Transport) supersonic transport project and other concepts illustrate the value of this approach., 資料番号: AA0063609069, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, In this study, flow computations were performed on the NEXST-1 (National EXperiment Supersonic Transport-1) with unstructured mesh solver. Two different meshes and two different turbulence models were used. The computational results were compared with NEXST-1's flight data. As a whole, the computational result showed a good agreement with flight data, except the pressure distribution on the lower surface in alpha-sweep cases., 資料番号: AA0063609074, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, Flight test of a supersonic experimental airplane was performed by Japan Aerospace Exploration Agency to improve advanced aerodynamic design technologies for the next generation SST (SuperSonic Transport). The experimental airplane was designed to reduce the drag on a supersonic cruise condition. Surface pressure distributions, boundary layer transition locations and aerodynamic forces in the flight test were obtained to demonstrate the aerodynamic design concepts. The flight test results evaluated by comparing with the prediction results obtained by CFD (Computational Fluid Dynamics) based optimum design tools. The flight test results relatively corresponded to the CFD results. Aerodynamic design concepts for drag reduction were demonstrated qualitatively and quantitatively by the NEXST-1 (National EXperimental Supersonic Transport-1) flight test., 資料番号: AA0063609065, レポート番号: JAXA-SP-06-029E

航空宇宙技術研究所 17-19 Jan. 2000 東京 日本, National Aerospace Laboratory 17-19 Jan. 2000 Tokyo Japan, 超音速輸送機に対するNALの実験(NEXST(小型超音速実験機計画)-1)の第1段階において、超音速航行時での抵抗軽減のための空力設計に焦点が当てられた。CFD(計算流体力学)が、この研究には大変有益であると考えられた。設計段階で4つの空力概念が適用された。(1)矢羽根平面型、(2)ねじれ翼、(3)面積則胴体、(4)自然層流(NLF)翼である。NLFの概念は摩擦抵抗低減のために考案された。NLFは今までSSTに適用されたことはなく、このプロジェクト独自の試みである。NLFを実現するために、先ず、超音速航行速度での翼面上圧力分布を設計し、超音速微小擾乱方程式にもとづく逆設計手法を適用した。10回以内の繰り返し計算で、目標の圧力分布を得ることに成功した。また、遷移点は非圧縮性および圧縮性境界層遷移点予測の計算機コードを用いて予測し、良い結果を得た。以上のことから、CFDは解析の手法としてのみならず設計の手法として用いることができることが分かった。NEXST-1の詳細設計は1999年中に終了した。次のCFD設計の目標はNEXST-2であり、現在、基本設計段階にある。, At the first stage of NAL's experiment for supersonic transport (NEXST (National Experimental Supersonic Transport)-1) project, the aerodynamic design focused on reducing the drag at the supersonic cruising. The CFD (Computational Fluid Dynamics) was supposed to be a powerful tool for the study. Four main aerodynamic concepts were applied in the process of designing; (1) arrow type planform, (2) warped wing, (3) area-ruled fuselage, and (4) wings with a Natural Laminar Flow (NLF). The NLF concept was devised for the friction drag. It has never been applied to SST up to now, which is one of original idea for this project. To realize the NLF on the wing at the supersonic cruising, firstly a pressure distribution on wing surfaces at supersonic cruising speed was designed and secondly an inverse design method based on the supersonic small perturbation equation was applied. After less than ten iterations, the target pressure distributions were obtained. Then the transition positions were estimated successfully by the CFD code of the incompressible and compressible boundary layer transition prediction. It must be noted that the CFD code was useful not only as an analytical tool but also as a method for the designing. The critical design of NEXST-1 has been completed in 1999. The next target of CFD based design is NEXST-2, which is now on the preliminary design phase., 資料番号: AA0028638010, レポート番号: NAL SP-49T

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, Four parties participated in the 4th SST (SuperSonic Transport)-CFD (Computational Fluid Dynamics) Workshop. They analyzed aerodynamic characteristics of the NEXST-1 (National EXperimental Supersonic Transport-1) flight test according to their own CFD analysis plans. After the presentations by four parties, several viewpoints on the future studies were discussed at the workshop. This paper summarizes those discussion points, action items to be studied and JAXA's results on some of the action items after this workshop. As a conclusion, this workshop revealed present JAXA's CFD results were almost valid by comparing them with the flight test data, although some quantitative differences have still remained. Therefore, principal discussion points were focused on turbulence model, transition condition and effect of small parts in the comparison of drag polar curve between the flight test data and CFD results. As concerned with inviscid force characteristics, all CFD results showed a very good agreement., 資料番号: AA0063609076, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ 航空宇宙技術研究所 20011203-20011205 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport National Aerospace Laboratory 20011203-20011205 Tokyo Japan, CFD (Computational Fluid Dynamics) analyses of flow around NAL's non-powered supersonic experimental airplane are conducted. The results are compared with those of corresponding wind tunnel tests. In this paper, agreements and discrepancies between them are discussed in aspects of correspondence between calculation and test models., 資料番号: AA0063609058, レポート番号: JAXA-SP-06-029E

航空宇宙技術研究所 17-19 Jan. 2000 東京 日本, National Aerospace Laboratory 17-19 Jan. 2000 Tokyo Japan, NALの最初のSSTモデルに用いる翼の空力形状が、超音速逆設計法を用いて設計された。この手法は翼・胴体形状を取り扱い、各スパン位置での翼断面幾何形状を提供する。設計目標は、マッハ2の速度でのNLF(自然層流)翼の実現である。逆設計手法の最初の系を、いくつかの設計制約を満足するように修正しなければならい。この方法により、従来の設計法で設計したものよりずっと望ましい特性を持つように翼断面形状を設計した。NLF翼を実現する空力学については、修正した方法は良く機能する。設計翼の圧力分布は、目標圧力と良く一致した。制約については、その多くが修正した設計法を用いることで満たすことができる。しかし、厚さ制約制御の過程で困難さを見出した。圧力制約に対処する将来計画を論じた。, Aerodynamic shape of a wing for NAL (National Aerospace Laboratory)'s first SST model has been designed by a supersonic inverse design method. This method handles wing-fuselage configurations and provides wing section's geometry at every span for Navier-Stokes flowfields. The design target is an NLF (Natural Laminar Flow) wing at the speed of M(sub infinity) = 2.0. The original system of the inverse design method has to be modified so that several design constraints can be satisfied. By means of the method, a wing section shape has been designed to have much more desirable characteristics than that by the traditional design method. In terms of aerodynamics to realize an NLF wing, the modified method works very well. The pressure distribution of the designed wing shows good agreement with the target pressure. In terms of constraints, most of them can be satisfied in the modified design method; however, difficulty has been found during the process for thickness constraint control. The prospective strategy to cope with the thickness constraint is discussed., 資料番号: AA0028638012, レポート番号: NAL SP-49T

SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design 航空宇宙技術研究所 19980316-19980317 東京 日本, SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design National Aerospace Laboratory 19980316-19980317 Tokyo Japan, Aerodynamic shape of a wing for NAL's SST (SuperSonic Transport) has been designed by a supersonic inverse design method. This method handles SST's wing-fuselage configurations and provides wing section's geometry at every span. The design system consists of a new inverse problem solver and a Navier-Stokes simulation. The design procedure is iterative; the baseline shape is Successively modified as the process of the inverse problem solver and Navier-Stokes simulation is iterated until the pressure distribution given by the designed wing can be regarded to converge to the target one. The design target is a NLF (Natural Laminar Flow) wing for the wing-fuselage combination at the speed of M(sub infinity) = 2.0. Several design constraints have been intended to be satisfied. By means of the method, a better wing shape which has much more desirable characteristics has been designed than that by the traditional linear theory., 資料番号: AA0063609012, レポート番号: JAXA-SP-06-029E

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ: Toward the Future form NEXST-1 Flight Trial 宇宙航空研究開発機構 20061012-20061013 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport: Toward the Future form NEXST-1 Flight Trial Japan Aerospace Exploration Agency 20061012-20061013 Tokyo Japan, This paper provides CFD (Computational Fluid Dynamics) flow analysis results of the JAXA NEXST-1 (National EXperimental Supersonic Transport-1) for the 4th JAXA SST (SuperSonic Transport) CFD workshop. A hybrid unstructured grid system was employed for viscous aerodynamic analysis. Computations were conducted for the flight test condition, then computed and measured results are compared. Effects of some computational conditions, such as, choice of turbulence modeling, mesh density and numerical scheme, are studied., 資料番号: AA0063609075, レポート番号: JAXA-SP-06-029E

SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design 航空宇宙技術研究所 19980316-19980317 東京 日本, SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design National Aerospace Laboratory 19980316-19980317 Tokyo Japan, SJAC (Society of Japanese Aerospace Companies) Phase-2 SST (SuperSonic Transport) configuration was designed by JADC (Japan Aircraft Development Corporation) SST working group. CFD (Computational Fluid Dynamics) analysis was conducted on the wing-body configuration to obtain preliminary aerodynamic data including non-linear effects such as transonic effect and so on, which are not counted in the primary design procedure. The results and the information obtained though this study are going to be discussed on this paper., 資料番号: AA0063609013, レポート番号: JAXA-SP-06-029E

航空宇宙技術研究所 17-19 Jan. 2000 東京 日本, National Aerospace Laboratory 17-19 Jan. 2000 Tokyo Japan, 縮尺超音速実験機の境界層の遷移特性を数値的に解析し、同実験機の翼が自然層流(NLF)翼であることを確認した。従来の非圧縮性およびe(exp N)法に基づく新たに開発された圧縮性遷移予測法を用いて、以下の結果によりNLF特性が良く確認された。(1)NLF翼設計に適用された階段状の圧力分布が最適であることが分かった。(2)圧縮コードにより推定されたNの値をNASAでの実験値と比較して、NLF翼の広い層流領域が期待された。(3)ナビエ・ストークス方程式コードより予測される境界層分布は、圧縮性境界層コードに比べて小さなN値を与えた。(4)付着線コンタミネーションによる遷移は予測されなかった。, Boundary layer transition characteristics of a scaled supersonic experimental airplane were numerically analyzed to confirm its Natural Laminar Flow (NLF) wing design. Using conventional incompressible and newly developed compressible transition prediction codes based on the e(exp N) method, the NLF characteristics were well confirmed through the following results. (1) The step function type target pressure distribution applied to the NLF wing design was found to be optimum. (2) Comparing the transition N value estimated by the compressible code with experimental results obtained by NASA, wide laminar region of the NLF wing was expected. (3) Laminar boundary layer profiles estimated by a Navier-Stokes code led to smaller N value than that by the compressible boundary layer code. (4) No transition due to attachment-line contamination was predicted., 資料番号: AA0028638011, レポート番号: NAL SP-49T

SST-CFDワークショップ.次世代超音速機の数値シミュレーション利用技術に関する国際ワークショップ 航空宇宙技術研究所 20011203-20011205 東京 日本, SST-CFD Workshop. International Workshops on Numerical Simulation Technology for Design of Next Generation Supersonic Civil Transport National Aerospace Laboratory 20011203-20011205 Tokyo Japan, In the aerodynamic design of the National EXperimental Supersonic Transport (NEXST) program, a supersonic Natural Laminar Flow (NLF) wing design concept was originally developed. Before the flight test of the NEXST-1 airplane, more detailed transition analysis should be performed in order to validate the NLF wing effect. Therefore, the NAL-ONERA cooperative research project started in April 2000, because ONERA had great ability of analyzing transition phenomena and NAL had some experimental transition data in supersonic flow. In the transition analysis on the sharp cone with a half angle of 5 degrees, the nose cone, and the NLF wing of the NEXST-1 airplane, almost good cross validation of both ONERA's and NAL's e(sup N) codes was obtained. Then, there was also high correlation between the S2MA wind tunnel test results and both laboratories' predictions under the assumption of a critical transition N value of 6. Moreover, about 40 percent chordwise region was estimated to be laminar at flight condition by NAL's code. In addition, a risk of transition due to attachment-line contamination was predicted at inner wing, using the well-known Poll's criterion., 資料番号: AA0063609064, レポート番号: JAXA-SP-06-029E

A fast heat transport phenomenon occurring in supercritical fluids, known as the 'piston effect', has been studied in relation to the sound wave generation in CO2. A periodic sound wave following the pulse sound wave generated by the piston effect, has been found. The amplitude of the periodic sound was very small, and the frequency was determined by the adiabatic compressibility., 資料番号: AA0063349058

Nowadays, the development of Supersonic Transport (SST) is attracting international attention again. However, in supersonic flights, we face the problem of 'sonic boom', which is induced by a shock wave. To alleviate the problem of sonic boom, the idea of reducing shock wave using Busemann biplane has been proposed. The interaction of shock waves between the two wings cancels out or reduces the shock waves created by individual wings. This research discusses two issues about Busemann biplane. One is the unstarted condition during acceleration. The other is the three-dimensional configuration of Busemann biplane. Unstarted condition happens due to the unsatisfied condition with Kantrowitz Donaldson Criteria. Unstarted condition that holds in the criteria is confirmed by analyzing accelerate phenomenon using Computational Fluid Dynamics (CFD) analysis. The other important issue is that the shock front should be planar to cancel out the shock waves between two wings. The caret wing induces a planar shock front. Three caret wing configurations are analyzed to examine how the planar shock front are created using CFD analysis., 資料番号: AA0063154019, レポート番号: JAXA-SP-06-010

A supersonic jet into complex geometry and its accompanying acoustic field have been investigated by using Euler Equation/Linearized Euler Equation (LEE) hybrid methodology here. As a first step, we have examined its performance by computing traveling sinusoidal waves. Then, we found that the hybrid methodology achieves higher accuracy than the Euler computation when combined with a fifth-order WENO (Weighted Essentially Non-Oscillatory) scheme. Based on this knowledge, we have conducted both of fully third-order Euler/LEE and third-order-Euler/fifth-order-LEE hybrid computations. According to the results, the computed acoustic field is mainly composed of three waves having distinct frequencies: a high frequency wave (0.5 Hz) which seems to be caused by interferences from the wall, a medium frequency wave (2.2 Hz) corresponding to first-mode eigenfrequency of the jet exit geometry, and a low frequency wave (10 Hz). The computed pressure values are rather different between the two results, although these frequencies are not so much. This fact implies that the more we extend the region of fifth-order computation, the more the computed pressure values approach to physically correct ones in quantitative view., 資料番号: AA0063154016, レポート番号: JAXA-SP-06-010

This paper provides the results of numerical analysis on the sound radiation and frequency characteristics of a rocket frame deflector. In general, a deflector is built in a rocket Launch Pad (LP) in order to reduce the influences of exhaust plume on a rocket body. The rocket frame deflector is expected to be a low frequency sound source because of its large dimension. Since the low frequency sound wave is dangerous for rocket body at its natural frequency, it is necessary to make the frequency characteristics of deflector clear. The characteristic frequency analysis of a modeled rocket frame deflector is conducted using a FEM (Finite Element Method) solver. It can be seen from the FEM analysis that the strong resonance occurs at low frequency in the modeled deflector. Three dimensional numerical analysis of sound radiation from deflector including rocket plume inflow is also conducted on a modeled LP using a hybrid code of an Euler flow solver and a Linearized Euler solver. From the sound radiation analysis, the characteristic sound wave from a modeled deflector has low frequency., 資料番号: AA0063154020, レポート番号: JAXA-SP-06-010

A direct numerical simulation of a supersonic wake undergoing transition to turbulence has been performed. The compressible Navier-Stokes are solved to study the acoustic generation mechanism for a Mach 2 wake. High-order upwind-biased compact schemes are used for spatial derivatives and a fourth-order Runge-Kutta scheme for time advancement. Navier-Stokes characteristic boundary conditions are employed in the vertical direction and periodic boundary conditions in the streamwise and spanwise directions for a time-developing DNS. The laminar wake was forced with a combination of the 2-D most unstable mode and a pair of oblique subharmonic unstable modes, which were obtained from linear stability theory for the viscous compressible equations. Forcing with a pair of oblique subharmonic modes yields streamwise/vertical counter rotating vortices in the saddle region of the plane wake. As streamwise/vertical counter rotating vortices evolve outside, their self-induction causes inclined braidlike structures to form in the wake, which are similar to the observations in the experimental supersonic flat plate wake transition. Acoustic waves of the plane wake are generated when two-dimensional rollup structures appear and rotate in a coherent fashion in the wake. 3-D elliptical structures play an important role for sound generation in the plane wake., 資料番号: AA0063154018, レポート番号: JAXA-SP-06-010

This paper introduces a computational technique to compensate for the added numerical diffusion that is generated when uniform Cartesian coordinates are used to describe the flow around bluff bodies. Because of the staircase-like representation of the surface object, it was found that the added surface 'roughness' causes larger than expected separation region for some test cases (flow around a sphere, and flow around a cylindrical body). In order to control the velocity profile in the boundary layer, a new parameter blvr (boundary layer velocity ratio) is define, and it is used to set the negative value of the viscosity along the surface. Extensive visualizations of flow past bluff bodies are performed using the present technique. Numerical solutions of the governing Navier-Stokes equations are carried out in a uniform Cartesian coordinates using a multi-directional finite difference scheme with a third-order upwinding. No explicit turbulence model is incorporated into the model, and the dependence of the solution on the blvr parameter is investigated., 資料番号: AA0063154006, レポート番号: JAXA-SP-06-010

In this study, the pressure estimation system was developed to enhance the value of PIV (Particle Image Velocimetry) measurement technique. The system conducts CFD (Computational Fluid Dynamics) computation to estimate pressure using PIV velocity data. The Poisson equation for pressure is the basic equation for the computation. After developing the CFD, the pressure estimation system was verified by using Navier-Stokes simulation results. Instead of PIV data, the velocity data of the simulation results were used as input data, and pressure values were calculated by the system. Then, the calculated pressure values were compared with those resulted by the Navier-Stokes simulation. After the validation check, pressure estimation using experimental data was done. The experimental data came from the experiment of wake measurement of a cranked arrow wing in JAXA's low-speed wind tunnel. Moreover, the influence of the measurement error to the system was evaluated. The evaluation can predict the correlation between measurement errors in velocity data and errors in estimated pressure values. The artificially generate errors were distributed over the original input velocity data. Then, analysis was done comparing pressure values obtained without the error and those with the error. It was found that input error magnitude reduced on the calculated pressure value., 資料番号: AA0063154011, レポート番号: JAXA-SP-06-010

Computations of opposing jet flow from the vertical landing rocket vehicle are performed by using Large Eddy Simulation technique. Calculated results are presented for the cases with and without jet ejection respectively, and are compared with experimental data obtained by several wind tunnel testings for the purpose of code validation. For the case without jet ejection, it is shown that dominant mean flow structures around the model are nearly reproduced in this calculation. Calculated time-averaged pressure coefficient distribution duplicates well with that of measured. For the case with jet ejection, although a quantitative agreement of measured pressure coefficient values with those given by the present calculation is yet to be accomplished, the general trend in the measured pressure coefficient distributions was reproduced well in the present calculation. It is also found that calculated characteristics of power spectrum density of pressure coefficient are similar to those of measured for the case with jet ejection., 資料番号: AA0063154026, レポート番号: JAXA-SP-06-010

In recent years, development of re-usable space transportation system is promoted. However, in order to realize it, development of the base technology is indispensable, particularly, aerodynamic heating and characteristics with strong real gas phenomena should be comprehended more precisely. Under such circumstances, acquisition of detailed experimental data like drag and heat transfer over hypersonic bodies is indispensable for the development of reliable space vehicles and thermal protection system (TPS), and also for the validation and progress of numerical estimation. In the present study, based on the numerical results a blunted double cone model (heat model) was designed and manufactured as the aerodynamic configuration design of re-entry capsules. The aim of this experiment is to obtain detailed data of heat flux and pressure distribution as well as the flow structures over the heat model. Coaxial thermocouples and pressure sensors are mounted at the frontal and base region thoroughly, and surface temperature and wall pressure are measured., 資料番号: AA0049500019, レポート番号: JAXA-SP-06-006

Rocket Based Combined-Cycle (RBCC) engines are currently being explored as advanced propulsion for space transportation. JAXA has been conducting RBCC engine research by using various experimental facilities. In order to clarify the experimental results and contribute to the improvement of designing, the analysis of the RBCC engine in an ejector-jet mode was carried out using the CFD (Computerized Fluid Dynamics) code developed in-house for unstructured grids. CFD replicated the basic flow structures in regard with barrel shocks and pseudo-shock waves. Two cases of Mach numbers (i.e., M = 1, 1.1) entering to the combustor were simulated and remarkable differences were found on the subsonic areas in the flowfields. The effect of Mach number was found to be one of the important parameters for better match to the experimental data. Vorticity and mixing were also investigated. In addition to the transverse vorticity, the streamwise vorticity were generated in the air flow, and they stirred and affected the nitrogen/air mixing., 資料番号: AA0049500027, レポート番号: JAXA-SP-06-006

A 3D flutter of a sweptback wing is numerically investigated by coupling inviscid flow simulations with a bending-torsion beam structure model. The beam model of a 3D wing is a natural extension of a widely used 2D 2-degree-of-freedom structure model and it considerably simplifies the structure analysis. Finite-volume unstructured CFD is used with the aim of applying various configurations in the future. The grid is dynamically deformed in accordance with the wing motion. A truss-analogy is applied to determine the new positions of interior points when the displacements of boundary points are specified. A well-known experiment using the AGARD standard aeroelastic configuration called 445.6 wing is selected for the validation. The developed methods successfully simulated the experimental flutter boundary., 資料番号: AA0049212015, レポート番号: JAXA-SP-05-017

In practical computations with structured grids around complex bodies, singular points can be frequently found where an abrupt grid change exists. These singularities pose troublesome problems when some high order and high resolution scheme is applied. An excellent theory has been proposed, which solve the above singular problem by decomposing a computational domain into two blocks along a line or surface which contains the singular points and by imposing accurate characteristic-based interface conditions at the block interface. However, the original theory has limitations on the combination between the adjacent computational coordinate definitions, and these two coordinates have to be the same direction and the same coordinate index at the block interface. For more flexible coordinate arrangement, the original characteristic interface treatment is extended and generalized. Consequently, the coincidence of the computational coordinates in the adjacent blocks becomes unnecessary, and robust multi-block computations can be successfully realized. Numerical test analysis is performed for verification of the new theory, and excellent performance is confirmed., 資料番号: AA0049212005, レポート番号: JAXA-SP-05-017

To investigate hypervelocity scramjet characteristics, numerical simulation was applied to combusting flows in a M12-02 HM scramjet engine at a free-stream stagnation enthalpy of 5 MJ/kg and 7 MJ/kg, which corresponds to the flight Mach number condition of 10 and 12. The numerical results were compared with the experimental results obtained by using a large free-piston shock tunnel to verify the numerical procedure. Good agreement was obtained between the numerical and experimental results of the wall pressure distribution. The results gave detailed information on the ignition process in the combustor. To investigate influence of the uncertainty in the test flow composition on the combustion process in the engine, numerical results computed assuming the free-stream condition with and without a small amount of atomic oxygen were compared. At 5 MJ/kg condition, existence of atomic oxygen in the free-stream affected ignition process and better agreement with the experimental results was obtained in the case of the free-stream condition without atomic oxygen than in the case with atomic oxygen. The numerical simulation showed that the combustor gas temperature exceeded 2,600 K at near the combustor exit and loss in the net combustion heat release due to thermal dissociation of the combusting gas was still one of major causes to drop the combustor performance., 資料番号: AA0049212013, レポート番号: JAXA-SP-05-017

When computational simulation for low Mach number flows is conducted with a compressible flow solver, its performance deteriorates in terms of both convergence speed and accuracy. The deterioration of accuracy is caused by excessive artificial viscosity in low Mach flow regions. In order to improve convergence of an iterative method, a local preconditioning method for low speed flow is frequently used. The preconditioning method minimizes the wave speed range between a particle wave and an acoustic wave. When the preconditioning method is applied to an upwind scheme, its artificial viscous terms need to be modified by preconditioning matrix. This modification cures the accuracy problem of calculated flows in the low-Mach flow regions. In this paper, the effects of the preconditioning method on unsteady calculation of flows around an airfoil are discussed., 資料番号: AA0049212026, レポート番号: JAXA-SP-05-017

To improve scramjet combustor performance at the hypervelocity condition corresponding to flight Mach number from 10 to 15, optimization possibility of the combustion gas temperature was investigated based on combustion experiment of a large-scale scramjet engine using a large free-piston driven shock tunnel. A previous engine model was designed so that the gas temperature at the combustor entrance would become around 1,200 K at the hypervelocity condition to prevent excessive loss in the net heat release due to thermal dissociation of the combustion gas. The results showed that the pressure rise due to combustion was large at the beginning of the test time but dropped drastically as time progressed. It was considered that the heat release obtained in the combustor was not sufficient to maintain intensive combustion. Following the discussion above, a new scramjet engine model (M12-02) was designed with an aim of promoting combustion heat release and establishing steady intensive combustion. The inlet compression ratio of the M12-02 WZLS increased to raise the static temperature and density of the flow at the combustor entrance. As a result of the design modification, the pressure rise due to combustion increased and the duration of intensive combustion was extended. The combustor performance became the highest at stagnation enthalpy of 7 MJ/kg, which was the design condition of the M12-02, and it was considered that the remarkable improvement in the combustor performance was achieved by establishing both adequately low gas temperature in the combustor and fairly stable combustion. Application of a Hyper-Mixer (HM) injector to the M12-02 was attempted with an aim of further improvement in the combustion stability and the combustor performance. The combustion test results showed that the HM injector was quite effective in establishing steady intensive combustion at the hypervelocity condition. In addition to the improvement in the mixing rate, capability in ignition and flame-holding at very low gas temperature condition was improved and high combustor performance was obtained at the low enthalpy condition. However, the combustor performance with the HM injector dropped as the stagnation enthalpy increased and became almost the same as that with a simple back-step injector at the enthalpy condition of 7 MJ/kg or higher., 資料番号: AA0049056017, レポート番号: JAXA-SP-05-010

This report summarizes recent progresses of the research on CFD simulations for turbopumps of a rocket engine. The results for the cavitating flow in an inducer, the internal flow of a centrifugal pump for a LNG engine test bed, and the flow in a tandem inducer will be presented, comparing with experimental results., 資料番号: AA0049056011, レポート番号: JAXA-SP-05-010

An extendible nozzle is considered to be a feasible device to improve the performance of booster engines because it has possibilities to provide altitude compensation and to achieve higher specific impulse. The booster engines with the extendible nozzle have to deploy the extendible nozzle during firing. For the design of the extendible nozzle and its driving mechanics, it is required to clarify the characteristics of dynamic loads acting on the extendible nozzle. In order to clarify the characteristics of dynamic loads, firing tests on a high altitude test stand were conducted with varying the ambient pressure to simulate the altitude change. The thrust and side loads acting on the extendible nozzle were evaluated quantitatively by use of a supporting system consisted of six rods. Using the measured load data and the calculated nozzle pressure distribution data the history of flow separation point was estimated. It was found that the large inverse thrust induced during the transition between full-flow and fully separated flow was caused by the pressure balance around the extendible nozzle., 資料番号: AA0049056009, レポート番号: JAXA-SP-05-010

3D CFD simulation was performed to examine the internal flow in a LNG turbopump. Rocket Engine Research Center of JAXA has been constructing an engine test bed, in which LNG (Liquid Natural Gas) is used as a propellant. The engine test bed has a LOX/LNG turbopump to pressurize propellants. The LNG turbopump section was built so far and the preliminary tests of this section have been started. The LNG turbopump uses an open impeller to meet the requirement of a low specific speed of 87 (m, cu m/min, rpm) and to avoid manufacturing complexity. A balance piston is adopted to keep thrust balance. The internal flow network is very complex and hard to estimate flow rates inside the pump in experiments. CFD analysis was performed for the LNG turbopump section to estimate internal flows and to predict performances of the pump. The calculated results were compared with experimental data and showed reasonable agreements., 資料番号: AA0048469036, レポート番号: JAXA-SP-04-012

CFD解析や風洞試験への適用を通して、ウェークトラバースによる空力特性推算法(後流評価技術)は、航空機の空力設計にとって重要な情報をもたらす空力設計診断ツールとして有用であることが明らかとなった。, 資料番号: AA0048140003, レポート番号: JAXA-SP-04-011

老朽化に伴い1995年から1997年にかけて測定部の大きさの変更を伴う大改修を行った航空宇宙技術研究所0.8m×0.45m高Re数遷音速風洞(第2遷音速風洞)を用いて風洞改修後あまり行われていなかった測定部2次元性および総圧変動特性にデータを取得した。2次元性の確認試験では改修前の2次元風洞および海外の風洞(NAE)のデータのあるBGK1翼型を用いたが、失速近傍まで中央断面は2次元性を保つこと、改修前の風洞と遜色のないデータが得られることが分かった。また総圧プローブを用いた総圧変動特性も改修前とほぼ同等で低マッハ数の総圧変動は集合胴圧の制御能力によると考えられる低周波成分が主であることが分かった。, 資料番号: AA0048140002, レポート番号: JAXA-SP-04-011

航空宇宙技術研究所6.5m×5.5m低速風洞の精度維持管理および風洞特性検討を目的として、ONERA-M5標準模型を用いた低速風洞試験を実施した。試験の結果、全機6分力、主翼・胴体表面圧力分布およびタフトによる気流可視化により、主要な目的である風洞の較正データおよび標準模型の低速空力特性を取得した。また同風洞における風洞試験データの信頼性向上を目指し、風洞壁境界の空力干渉量を古典的修正法により求め、この値の妥当性の検証をパネル法による数値解析結果との比較により行った。その結果、標準模型程度の大きさであれば風洞壁境界修正法は古典的方法で概ね十分であることを確認した。またスティング模型支持装置の模型への空力干渉量は、風洞壁境界修正量と同程度であるということがパネル法により推測された。さらに、主翼表面上にラフネスを貼り付けその粒径を変化させながら抵抗を計測し乱流遷移点を推定する試みを行い、同風洞において強制遷移を行う際の標準的なラフネス高さを定めた。, 資料番号: AA0048061008, レポート番号: JAXA-SP-04-008

Pressure-sensitive paint techniques make the measurement of global surface pressure distributions possible, which are different from point measurements by conventional pressure transducers. The present study was conducted to make the pressure-sensitive paint measurement practicable at the NAL 1 m x 1 m supersonic wind tunnel. Since the temperature of a model decreases with test time in the wind tunnel, the temperature dependence of luminescent intensity of the pressure-sensitive paint is a large source of error for pressure measurements. A temperature-correction method using a temperature-sensitive paint for the pressure-sensitive paint was applied. The temperature-corrected pressure-sensitive paint data were compared with pressure data using pressure transducers. As a result, good agreements were obtained between the pressure-sensitive paint data and pressure transducer data., 資料番号: AA0048061004, レポート番号: JAXA-SP-04-008

資料番号: AA0048062002, 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: AA0048062003, レポート番号: JAXA-SP-04-009