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A CFD (Computational Fluid Dynamics) workshop was held in the 44th Aircraft Symposium, October 2006, in order to evaluate domestic CFD technology for civil transport aircrafts in high-lift configurations. JAXA provided an experimental data for comparison obtained in the JAXA large low-speed wind tunnel using a half model equipped with high-lift devices, fuselage, nacelle-pylon, slat tracks and Flap Track Fairings (FTF). This paper briefly presents the experimental model, objectives and test cases of the CFD workshop, summary of comparison, and obtained knowledge and issues on CFD and wind tunnel testing for high-lift configuration., 資料番号: AA0063742027, レポート番号: JAXA-SP-07-016

Through advanced detail simulation tests, adjusting model mass and control factors made successful measurement of aerodynamic force on a sphere at Reynolds numbers around critical one at the JAXA 60 cm magnetic suspension and balance system. The obtained averaged aerodynamic force for 8 seconds shows that the minimum drag coefficient after the critical Reynolds number reaches 0.062 and that any hysteresis phenomenon on drag coefficient with respect to Reynolds number was not observed around the critical one. Through the obtained fluctuations of aerodynamic force, the shrunk force fluctuating around the sphere axis parallel to main flow and some force swirling the sphere about the axis for a short time were observed at Reynolds numbers larger than the critical one. And some oscillating force on a plane including the axis was observed in the test of a sphere with double tapes for boundary layer trip in large Reynolds number region around the critical one., 資料番号: AA0063742005, レポート番号: JAXA-SP-07-016

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

Laminar-turbulent transition is a crucial phenomenon appearing in a variety of industrial applications. However the involved physical mechanisms as well as methods for reliable and accurate prediction of transition are still a matter of active research. In the present contribution, we give a brief overview on recent advances in the simulation and prediction of transitional and turbulent wall-bounded shear flows. The focus is on Large-Eddy Simulations (LES), which differ from Direct Numerical Simulations (DNS) by resolving only the large-scale, energy-carrying vortices of the fluid flow, whereas the fine-scale fluid oscillations, assumed to be more homogeneous, are treated by a SubGrid-Scale (SGS) model. The application of LES to flows of technical interest is promising and LES is getting more and more applied to practical problems. The main reason for this is that LES provides an increased accuracy compared to solutions of the (statistical) Reynolds-Averaged Navier-Stokes equations (RANS), while requiring only a fraction of the computational cost of a corresponding fully-resolved DNS. Nevertheless, LES of practical transitional and turbulent flows still require massive computational resources and the use of large-scale computer facilities., 資料番号: AA0063908005, レポート番号: JAXA-SP-07-026E

This report presents experimental data obtained in lowspeed wind tunnel testing for high-lift configuration aircraft model JSM (JAXA Standard Model). Two time of the testing were conducted at 6.5 m by 5.5 m Low-speed Wind Tunnel in JAXA (JAXA-LWT1) aiming to store validation data for CFD (Computational Fluid Dynamics) analysis and to observe flow physics on high-lift system. JSM is a half type model which assumes 17 percent similarity of a modern 100-passenger class regional jet airliner. The model is equipped with leading edge slat, double-slotted flap at the inboard and single-slotted flap at the outboard, flow-through nacelle, in addition, a circular fuselage and Flap Track Fairings (FTF) so that the detailed flow fields occurring in actual aircraft can be provided. At first, repeatability and sensitivity of the force and the moment data in the case the detail of the model was slightly changed were assessed as well as the accuracy was confirmed from large amount of data in the first and the second testing. Variation in lift performances between long-cowling nacelle and short-cowling nacelle was observed with couple of the measurements and visualizations. The results showed much better stall performances of the short-cowling configuration than the long-cowling one that was due to less interference between the nacelle pylon and the wing. Boundary layer tripping on flap could clarify the loss in aerodynamic performances. Finally, improvement of stall characteristics by a vortex generator and an efficient way to decide its installation location by the application of Kriging based Genetic Algorithm (GA) on nacelle was discussed., 資料番号: AA0063742028, レポート番号: JAXA-SP-07-016

Karman-vortex shedding induced by wake instability behind a 2-D airfoil is of interest at relatively low Reynolds numbers, to understand how a vortex street develops in the wake region, and also how an aeolian tone is generated therefrom; this is supposed to show a different flow mechanism from a high-Reynolds number flow that exhibits more complicated viscous-flow phenomena. In this study, we performed numerical simulations of a 2-D laminar flow past an NACA0012 airfoil. The computation was run on the simulation code being developed to predict an aero-acoustic problem by using a highly accurate numerical method. For the verification of our code, the results are compared with the experimental data obtained in a similar flow configuration and also with an available numerical study. The detailed flow visualization is presented to understand the mechanism of the wake instability that results in vortex shedding., 資料番号: AA0063908013, レポート番号: JAXA-SP-07-026E

This numerical study examines the supersonic rectangular-nozzle flow by focusing attention on the development of secondary flow and corner-flow and their effects on the boundary layer instability and transition. The results show that secondary flows appear in the side wall boundary layers and form into streamwise vortices in the corner region immediately behind the throat station where the boundary layer momentum thickness Reynolds number is 400. The streamwise vortex increases in number downstream and induces a local normal-to-wall flow as large as 4 percent of the local mainstream. The present simulation is in good agreement with the corresponding experiment, in particular, on the appearance of cross-flow instability on the side wall boundary layer and the streamwise location where the streamwise vortex first appears. The corner-flow is found to be of extremely unstable nature suggesting that the turbulent transition observed is directly caused by the corner-flow disturbances., 資料番号: AA0063608008, レポート番号: JAXA-SP-07-009

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, 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: 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

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, 資料番号: AA0063609063, 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, A flight test of a supersonic experimental airplane (NEXST-1: National EXperimental Supersonic Transport-1) was conducted successfully by JAXA in October 2005. In this study some aerodynamics design concepts were verified by comparing the flight test results with CFD (Computational Fluid Dynamics) analysis using JAXA's CFD codes, a structured mesh code UPACS (Unified Platform for Aerospace Computation Simulation) and an unstructured mesh code TAS (Tohoku university Aerodynamic Simulation). After the CFD codes was verified using a wind tunnel test results, the CFD results were compared with the flight test results. The aeroelastic deformation and the boundary layer transition data of flight test were taken into account to improve CFD analysis. The effect of the design concept of the NEXST-1 was also confirmed by CFD analysis., 資料番号: AA0063609072, 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, 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

An advanced Large Eddy Simulation (LES) technique with high accuracy and resolution is proposed with the Approximated Deconvolution Method approach, the generalized characteristic interface conditions, and the generalized Navier-Stokes boundary conditions. Several applications to flows around 2-D NACA0012 aerofoil are presented and discussed. Numerical results show that this method can capture the laminar to turbulent transition phenomena in the order of Reynolds number of 106. In order to validate the present LES technique in more detail, the results are compared with existing wind-tunnel experiments and linear stability analyses with the eN method and discussed., 資料番号: AA0063156020, レポート番号: JAXA-SP-06-013

In this paper, the gamma-Re(sub theta) transition model proposed by Menter et al. is investigated. Correlations not described in the original paper are reproduced based on other literature and validated by several test cases: T3A- (Tu = 1 percent), T3A (3 percent), T3B (6 percent) for flat-plate boundary layers with bypass transition, and Schubauer & Klebanof (0.2 percent) for flatplate boundary layer with natural transition. T3LC test case is used for separated flow transition which is important for RANS simulation applied to separated flow problems. Transition position on flatplate for each free-stream turbulent intensity has good agreement with experiment, although transition lengths estimated for low free-stream turbulent intensity are too short. For T3LC test case, the length of laminar separation bubble is correctly predicted by using the present transition model., 資料番号: AA0063156010, レポート番号: JAXA-SP-06-013

To obtain a better understanding of the mechanism for supersonic boundary layer transition, experiments are made at Mach number 2.0 by using a small supersonic wind-tunnel, with focusing efforts on finding non-linear disturbances governing the transition. We examine a boundary layer along the nozzle wall and its transition process by means of quantitative schlieren optical system. It is found that vortical structures our schlieren system visualizes are mostly similar to each other in shape, being stick-like and inclined at about 30 to 70 degrees to the wall, and start to grow often forming a group at streamwise stations where the momentum thickness Reynolds number is around 580 and the transition proceeds as they increase in number to cover the whole boundary layer. We examine the statistical feature of such oblique vortical structures using 1,800 samples and show that these structures are important non-linear disturbances as the mechanism of supersonic boundary layer transition., 資料番号: AA0063156009, レポート番号: JAXA-SP-06-013

An advanced Large Eddy Simulation (LES) technique with high accuracy and resolution is proposed. This method solves compressible 3-D Navier-Stokes equations with the Approximated Deconvolution Method approach as LES approach and has several sophisticated techniques such as highly tuned compact schemes, the generalized characteristic interface conditions, and the generalized Navier-Stokes boundary conditions. Several applications to flows around 2-D NACA0012 aerofoil are presented and discussed. Numerical results show that this method can solve very complicated subsonic flow including the laminar flow separation at the leading edge region, the transition in the bubble and the reattachment. It also has capability to capture the natural laminar to turbulent transition phenomena in the order of Reynolds number of 10(exp 6). In order to validate the present LES technique in more detail, the results are compared with existing wind-tunnel experiments and linear stability analyses with the e(sup N) method and discussed., 資料番号: AA0063154022, レポート番号: JAXA-SP-06-010

Breakdown process of the Klebanoff mode and origin of turbulence are experimentally investigated in detail. An acoustic excitation using a 2D roughness attached on the wall are used for T-S wave generation. Piezo-ceramic actuators are used to introduce the Klebanoff mode fluctuation. It is found that another shear layer, in which the spike phenomenon takes place and the turbulence starts, exists between the Lambda structures aligned in rows. The spreading of the turbulence along the shear layer induced by the Lambda structures is newly found., 資料番号: AA0049208011, レポート番号: JAXA-SP-05-012

The present experiment focused on revealing effect of scales and directional components of free stream turbulence on boundary layer transition, especially the growth rate of the disturbance energy in the streamwise direction. Free stream turbulence generated by a turbulence grid mounted upstream of the contraction has typical character of axisymmetric turbulence with strong anisotropy. The experimental results with the anisotropic free stream turbulence show the non-modal growth disturbance even at 0.7 percent turbulence intensity. The correlation between the filtered free stream fluctuation and growth rate of the disturbance suggests that the wall-normal velocity fluctuation in the free stream dominates the disturbance growth., 資料番号: AA0049208004, レポート番号: JAXA-SP-05-012

A series of wind tunnel tests were conducted to develop a boundary-layer tripping technique and also a detection technique for the boundary-layer state in transonic wind tunnels in ISTA/JAXA. A commonly used disk-type roughness element was used in the test and the conventional criterion for the effective roughness height was confirmed valid in the present case, also. A detection methodology for the boundary-layer state based on the quantitative measurement of the aerodynamic heating was applied to this experiment in a blow-down transonic facility, using both the Temperature Sensitive Paint (TSP) technique and the Infra-Red Thermograpy (IR) technique. The results indicate the methods are both useful in transonic experiments. Transition detection in a closed-circuit transonic wind tunnel was also achieved by varying the reservoir temperature, but it is found that the technique is not necessarily good for the boundary-layer state determination without any additional roughness for the reference purpose., 資料番号: AA0048470003, レポート番号: JAXA-SP-04-013

This paper presents current status and issues of aircraft drag prediction using numerical methods based on Reynolds-averaged Navier-Stokes equations (RANS) in JAXA. Several results on an airfoil and aircrafts show the prediction of aerodynamic drag is very sensitive to grid density and numerical schemes as well as transition in boundary-layer. Although state-of-the-art RANS CFD can not satisfy the requirement for aerodynamic design of aircrafts enough, it already provides almost the same reliability as wind tunnel tests in some cases. To overcome the issues in the prediction reliability, tighter cooperation between numerical and experimental methods is important besides efforts to improve numerical schemes and turbulence models., 資料番号: AA0047776019, レポート番号: JAXA-SP-04-002

Based on the idea that turbulence is the flow phenomena caused by the movement of vortex, the structures of turbulent boundary layer are discussed. The circulation is the most important factor for characterizing the flow structures. The large scale motions such as coherent structure of the outer layer and the burst structure of the inner layer can be explained by the interactions between vortices., 資料番号: AA0047776002, レポート番号: JAXA-SP-04-002

The results of an experimental investigation on the effect of a spanwise oscillating Lorentz force on the near-wall turbulent structures for drag reduction are presented in this paper. Comparing with the case of turbulent control by spanwise wall oscillation, a parameter called equivalent spanwise wall velocity W(+)(sub eq) = StT(+)/(2pi Re(sub tau)), is defined and emphasized as a virtual spanwise wall velocity scale under Lorentz force oscillation. It is shown that when W(+)(sub eq) approximately equal to 10 to 15, up to 30 to 40 percent skin frictional drag reduction could be achieved just at the downstream of the control area. Meanwhile, an oscillation amplitude investigation suggests that to achieve maximum skin friction reduction, the oscillation Lorentz force should be able to shift the near-wall streamwise vortices into spanwise direction by a peak to peak displacement of about 100 wall units during each period. A conceptual model to explain the mechanism for drag reduction under the spanwise Lorentz force oscillation is also presented., 資料番号: AA0047776013, レポート番号: JAXA-SP-04-002

The work described here relates to the method measuring the boundary layer transition at large-scale and low-speed wind tunnel. The objective of the research is to improve the test technique and to know a precise procedure about transition measurements, especially for an aircraft model at such type of wind tunnel. At JAXA 6.5 m x 5.5 m Low-Speed Wind Tunnel (LWT1), the boundary layer growing on the left-wing of the ONERA-Model was investigated using two kinds of the quantitative methods (Preston Tube and Hotfilm Sensor) and also two kinds of the visualizing methods (Temperature-Sensitive Liquid Crystal, Infrared Camera). Throughout the experiment, a lot of knowledge were obtained. Transition points investigated by four kinds of methods were correctively agreed., 資料番号: AA0047776001, レポート番号: JAXA-SP-04-002

Results of the fundamental performance measurement of newly designed and constructed fluid dynamics research facility Hyuga Aerodynamic Research by Towing (HART) are reported. Outstanding features of the present facility are its ultra low turbulence intensity nature and capability of measuring wing-in-ground effect performance. Using this facility, high accuracy turbulent transition research, under floor flow measurement of vehicles, wing-in-ground effect research and others researches are possible., 資料番号: AA0047776018, レポート番号: JAXA-SP-04-002

An experimental investigation for an efficient tripping method of the hypersonic boundary layer is conducted with a 5 deg half angle sharp cone in ISTA/JAXA phi0.5 m hypersonic wind tunnel. Aerodynamic heating distributions and fluctuations of wall pressure are measured both with and without roughness elements. A set of transition data and wall pressure fluctuation measurements in the wind tunnel are obtained. A two-dimensional wavy roughness of a wave length approximately as two times much as the estimated boundary layer thickness doesn't show any effects on the location of the boundary layer transition, while in the case where the boundary layer thickness is approximately 1.25 times of the half wavelength, a clear delay in transition is observed. Effects of conventional three dimensional roughness elements on tripping effectiveness are also examined., 資料番号: AA0047776012, レポート番号: JAXA-SP-04-002

摩擦抵抗を軽減した自然層流翼を基本概念として設計されたSSTロケット実験機の実証飛行試験に先立ち、遷移予測コードの検証用データベースを構築するため、ONERA回流式大型超音速風洞S2MAにおいて全機模型の遷移過程を計測した。本風洞では国内の風洞で実現されない飛行条件に近い高レイノルズ数が達成され、またマッハ数、風洞総圧、迎角および横滑り角などが制御可能である。多点熱膜流速計から決定された遷移点は、赤外線画像の計測結果とよく一致し、レイノルズ数の増加に伴って上流へ移動することが確かめられた。前縁から遷移点までの距離に基づく遷移レイノルズ数は、マッハ数を2に固定した場合、主流レイノルズ数には依存しなかった。一方、迎角依存性は、内翼と外翼で異なっていた。内翼では、設計迎角において最も遷移点が下流に移動した。それに対し外翼では、5度までの範囲で迎角が大きいほど遷移が遅くなった。これらの迎角依存性を導く遷移機構は、圧力分布から理解可能である。また、主流の総圧変動はマッハ数によって0.1〜0.35％と大きく変化したが、風洞総圧には依存しなかった。, 資料番号: AA0047428004, レポート番号: JAXA-SP-03-003

Using the prediction system of boundary-layer transition, the spatial growth of traveling waves is examined in the boundary layer over the outer wing of the NEXST-1 semispan model. It is found from the computational results that TS waves are dominant to transition on the wing. Moreover, it is noted that TS waves directed to the leading edge are amplified even in the accelerated boundary-layer flow near the leading edge. However, the growth of TS waves is suppressed around the designed angle of attack., 資料番号: AA0045948035, レポート番号: NAL SP-57

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

A series of experiments was conducted on high enthalpy effects on hypersonic boundary layer transition in T5 hypervelocity shock tunnel. This series concerns transition of the attachment line boundary layer on a swept cylinder. The observed trend of transition Reynolds number with total enthalpy shows strong transition delay at a larger sweep angle for carbon dioxide, while no significant effect is observed in nitrogen., 資料番号: AA0045947028, レポート番号: NAL SP-56

A numerical research is carried out in order to clarify former experimental attempts to introduce various waves into the boundary layer using piezo-actuators. The numerical results are in good agreement with the experimental results. From the information which can be obtained only by the numerical approach, such as vorticity components near the wall, it is shown that the process of waves generated into the boundary layer is complicated, but it can be explained., 資料番号: AA0045947014, レポート番号: NAL SP-56

Small disturbances superimposed on a growing boundary-layer flow are governed by partial differential equations, but a simpler estimation of stability characteristics will be done if the exact disturbance equations are reduced to an eigenvalue problem of the corresponding ordinary differential equations. In this paper, a simple model of the ordinary differential system is proposed, which includes the most important terms of boundary-layer growth and wall curvature., 資料番号: AA0032827005, レポート番号: NAL SP-54

Through systematic investigation concerning the crossflow dominant boundary layer transition, the complicated transition mechanism was clarified experimentally. Namely, transition starts with stationary streamwise vortex (crossflow vortex) generation. This is called 'primary instability'. Then, primary instability induces three unsteady instabilities, F(sub 1), F'(sub 1)(relatively low frequency), F(sub 2) (relatively high frequency). F(sub 1) travels to crossflow direction, F(sub 2) travels to streamwise direction, and F'(sub 1) travels to the opposite direction with F(sub 1). It seems to be that F(sub 2) enhances the transition into full turbulent. Taking into consideration of such flow mechanism, a prediction method has been developed by introducing C(sub 1), C(sub 2) which is valid in general three-dimensional boundary layer transition. Knowing the threshold value for critical condition either by experimentally, or numerically, it is possible to predict transition onset simply calculating potential flow information. A unique and practical selective suction method to delay the transition with less suction power. It was possible to delay the transition with one third of the power of uniform suction., 資料番号: AA0032827009, レポート番号: NAL SP-54

In the experiments conducted at FHI, boundary-layer transition was observed on the outer wing of the SST half model. In many cases, boundary-layer transition on a swept wing has been caused by the growth of crossflow disturbances. Then, the spatial growth of crossflow disturbances on the outer wing is investigated using the prediction system of boundary-layer transition. The results show their growth is too small to cause transition. It is fairly possible that the transition observed in the experiments was caused by TS waves., 資料番号: AA0032819046, レポート番号: NAL SP-53

Infrared-image technique has been prevailingly applied to diagnose laminar-turbulent transition in compressible boundary layers. The purpose of the present paper is to establish this technique for low subsonic flow. To produce measurable temperate difference between laminar and turbulent states above ambient temperature, the model was pre-heated by a hair-drier prior to each measurements. Infrared image captured in two-dimensional boundary layer was quantitatively reduced to evaluate transition location. It was found that reduced results are in good agreement with those by means of conventional methods such as hot-wire anemometer and Preston tube., 資料番号: AA0032827015, レポート番号: NAL SP-54

The laminar-turbulent transition of spherical Couette flow between two independently rotating concentric spheres for the gap ratio 0.14 where the Taylor instability occurs was investigated by simultaneous spectral and flow-visualization measurements. The conclusions are as follows. Rotation frequencies of disturbances do not so depend on the ratio of the outer- to the inner-sphere rotating Reynolds number if they are non-dimensionalized by the mean rotation frequency of both spheres. Herring-bone waves and wavy shear waves appear with increasing rotation effect for the co-rotating spheres, while spiral vortices and twists appear with increasing shear effect for the counter-rotating spheres. Flow regimes and their occurrence-region in the plane of the inner- and outer-sphere rotating Reynolds numbers differ by the Reynolds-number history., 資料番号: AA0032827006, レポート番号: NAL SP-54

The laminar-turbulent transition of shear flows is an old subject. The linear growth of small-amplitude disturbance and nonlinear interaction of grown-up fluctuation have been well-documented. The remaining and most important process is the randomization of ordered fluctuations. In order to describe the process a quantity called randomness factor is needed. This concept is useful not only for fluiddynamics but also for many transition processes. Subjects discussed here include physical phenomena, plants, animals and social problems., 資料番号: AA0032827011, レポート番号: NAL SP-54

The effect of accuracy of boundary layer calculations on the prediction of laminar-turbulent transition of boundary layer is examined by comparing the two scheme of boundary layer calculations, using a similarity equation and a Navier-Stokes code. The flow field is a boundary layer around an NACA0015 wing model. The Orr-Sommerfeld equations are used for the local stability analysis and the N factor of the e(sup N) method is used for a criterion. From the present study a slight difference of velocity profiles causes a significant difference of an end result., 資料番号: AA0032819048, レポート番号: NAL SP-53

航空宇宙技術研究所 27-29 Sept.2000 東京 日本, National Aerospace Laboratory 27-29 Sept.2000 Tokyo Japan, e(exp N)法を使用した遷移位置の予測には、空間成長速度Nに及ぼす外部擾乱の定量的評価が強く求められるため、2次元翼流れ内の層流-乱流遷移に及ぼす音響騒音の影響を実験的に検討した。最初に、同一モデルに対する自然遷移を異なる5風洞中で調べた。各因子の貢献度は分離不可能であるが、大音響騒音および/または大残留主流乱れを用いた風洞では早期遷移を観測した。そこで、測定胴内に設置したスピーカを使って人工音響騒音を導入した。大きな励起が遷移過程を促進することを確認した。Nと音圧レベル(SPL)間の経験的関係N=-0.6×SPL+74を提案した。, Influence of acoustic noise on laminar-to-turbulent transition in 2-D wing flow is investigated experimentally, since the quantitative evaluation of external disturbances on spatial growth rate N is intensively required to predict the transition location by the use of the e(exp N) method. At first, natural transition on an identical model is examined in five different wind tunnels. Early transition is observed in the wind tunnels with large acoustic noise and/or large residual free-stream turbulence, although the contribution of each factor is not separable. Therefore artificial acoustic noise is introduced by means of a loud speaker installed in the settling chamber. Larger excitation is confirmed to facilitate the transition process. An empirical relation between N and Sound Pressure Level (SPL) is proposed as N = -0.6 x SPL + 74., 資料番号: AA0028637014, レポート番号: NAL SP-48T

27-28 Sept. 1999 (25th). 27-28 Mar. 2000 (26th), CFD(数値流体力学)手法を用いて自然層流翼に適用することを意図したNALのSST(超音速輸送機)計画に関連して境界層遷移計測を実施した。実験の目的は最先端的な境界層遷移に関するデータベースを構築して翼上の遷移点予測の参考とし、また高度な計測技術を確立して航空機の飛行試験に適用することにある。本論文では、代表的な2つの予備実験の結果を報告した。ひとつはビーチクラフト機を使った飛行試験で、低乱流環境下における遷移過程を把握し、飛行試験の諸課題を明らかにするために実施した。Tollmien-Schlichting波が前縁近傍領域で検出され、スペクトルの振幅分布がSALLY(3次元層流境界層安定性解析)コードを使った数値予測値と非常によく一致したことから、風洞内で行う遷移計測が好結果を得ることが分かった。もうひとつは、超音速風洞内においてSST実験機の半裁模型に対して行った計測である。設計点、すなわち、迎角α=2.7度の時のマッハ数M=2では、層流状態は少なくともX/C=0.5に耐えることが分かった。自然層流翼はSST実験機の飛行試験で実現すると期待されている。, Boundary-layer transition measurements have been conducted associated with the NAL SST (Supersonic Transport) project, which is designed to apply natural laminar flow wing using CFD (Computational Fluid Dynamics) technique. The purpose of experiments is to construct the database on boundary-layer transition most advanced which will be referred to predict the transition point on the wings, and to establish high quality measurement techniques, which will be applied to the flight test of the airplane. In this paper, the results of two typical preliminary experiments are introduced. One is flight tests using Beachcraft airplane, which are performed to grasp the transition process under low disturbance environment and to clarify the problems of flight test. Transition measurement as conducted in wind-tunnel was found to be successful, because Tollmien-Schlichting wave is detected at the region near the leading edge, and its amplitude distribution of the spectrum is in very good agreement with the numerical prediction by SALLY (3D laminar boundary layer stability analysis)-code. Another is the measurement on a half-model of SST experimental airplane in a supersonic wind tunnel. At the design point, i.e. Mach number M = 2 at an angle of attack alpha = 2.7 deg, it is found that the laminar state sustained at least to X/C = 0.5. Natural laminar flow wing is expected to realize on the flight test of SST experimental airplane., 資料番号: AA0028636001, レポート番号: NAL SP-47

27-28 Sept. 1999 (25th). 27-28 Mar. 2000 (26th), 以下の主題を論じた。加速乱流境界層、大気積雲対流と乱流渦、安定成層乱流、乱流遷移過程の力学機構、剥離流の前兆信号、渦輪の変形過程、Goertler渦、壁乱流構造に及ぼす境界層吸込み効果、マイクロバブル、球Couette系の層流乱流遷移、剥離流れに及ぼす速度変動効果、net型渦列、および音による境界剥離制御。, The following topics were discussed: turbulent boundary layer in accelerated flow, atmospheric cumulus convection and turbulent vortex, stably stratified turbulence, dynamic mechanism of turbulent transition process, precursor signal before flow separation, vortex ring in deformation process, Goertler vortex, effect of boundary layer suction on wall turbulence structure, microbubble, laminar-turbulent transition of spherical Couette system, velocity fluctuation effect on separated flow, net type vortex street, and acoustic control of boundary separation., 資料番号: AA0028636000, レポート番号: NAL SP-47

27-28 Sept. 1999 (25th). 27-28 Mar. 2000 (26th), 境界層遷移に影響を与える擾乱の実際の物理的機構を理解するために、境界層内の層流-乱流遷移の位置に影響する音波、乱流、および表面欠陥や表面振動に起因する擾乱などの主流撹乱を研究した。実験は、アルミ合金板から作製した2枚の平板モデルを用いて低乱流閉回路風洞中で実施した。遷移の初生は表面摩擦の上昇を計測して決定した。主流乱れはモデルの上流に設置したスクリーンを使って発生させ、主流音響擾乱は平板から下流に設置したスピーカを用いて導入した。主流および境界層における計測は自然バックグラウンド擾乱および音響励起下で行った。人工音場の影響下における不安定なTS(Tollmien-Schlichting)波発生源と遷移の課題を研究するため、平板境界層を用いた実験では音波受容性の計測に新規方法を適用した。平滑平板境界層の前縁受容性を定性的に研究するための実験を行った。スクリーンにより増強された主流乱れは境界層の内側で典型的な振動分布の変形を起こすことが分かった。遷移の筋書きは低撹乱主流で観察した遷移と類似している。主流乱れは音響受容性を変化させず、また局所受容性の全係数は上昇乱流下では主流撹乱の音波成分の転換により主として決まることが明らかになった。, The free stream disturbances, such as sound, turbulence and the disturbances due to the defects of surface and the surface vibrations, which are to influence on the location of laminar-turbulent transition in the boundary layers, have been studied in order to understand real physical mechanism of disturbances to affect boundary layer transition. The experiments were conducted in a low turbulence close-circuit wind tunnel using two flat plates model made of aluminum alloy sheet. The onset of transition was determined by measuring the rise in skin friction. The free stream turbulence was generated by means of the screen placed upstream from the model and the free stream acoustic disturbances were introduced using a loudspeaker placed downstream from the flat plates. Measurements in the free steam and boundary layer were carried out with natural background disturbance as well as under acoustic excitation. A new method to measure the sound receptivity was applied in experiments with flat plate boundary layer in order to study a problem of unstable TS (Tollmien-Schlichting) waves origin and transition under the influence of an artificial sound field. The experiments on the qualitative study of leading edge receptivity of the smooth flat plate boundary layer were carried out. It was found that the free stream turbulence enhanced by means of a screen had led to typical distortion of oscillation profiles inside the boundary layer. The scenario of transition was similar to that was observed in low-disturbed free steam. It is revealed that the free stream turbulence does not change the acoustical receptivity, and that the total coefficient of the localized receptivity is mainly determined by conversion of sound component of free stream disturbances under the elevated turbulence., 資料番号: AA0028636018, レポート番号: NAL SP-47

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, ワークショップの各部会における以下の課題について発表された研究が報告されている。(1)一般講義では、「次世代SSTに向けての研究」および「NALのSST研究プロジェクト」、(2)一般部会では、「SSTの設計とCFDに対する必要条件」(3)技術部会1では、「NALのCFD支援設計研究」、「数値最適化」および「多分野にわたる研究」、(4)技術部会2では、「設計のための次世代コンピュータシステム」および「CFDにおける基本的問題」。, Studies on the following topics in each section of the workshop are reported: (1) 'research toward the next generation SST' and 'NAL's SST research project' as general lectures; (2) 'designing SST and requirement on CFD' for the general session; (3) 'NAL's CFD-aided design research', 'numerical optimization' and 'treatment of multidiscipline' for the technical session 1; (4) 'next generation computer system for designing' and 'fundamental problems in CFD' for the technical session 2., 資料番号: AA0028638000, レポート番号: NAL SP-49T

極超音速飛行実験機における澱点近傍、エレボン部及び胴体、安定板における空力加熱を計測するためにセンサを新規に開発し、ランプ加熱試験によりセンサ周りの精密な熱モデルを定めることで実飛行実験にそれぞれ適用可能であることを確かめた。検証された熱モデルを用いてHYFLEX飛行中の空力加熱計測をおこない、極超音速風洞試験結果に基づく予測、CFD解析などと比較した。澱点近傍C/Cセンサの飛行データはt=100sec付近において、風洞試験結果に基づく予測と比べ低い加熱が得られた。非平衡流解析コードによるCFD結果との比較により、観察された現象は空気の解離/再結合に伴う実在気体効果によるものと考えられ、触媒姓の温度依存性を考慮することで定性的に説明できた。エレボン部C/Cセンサでは胴体下面セラミックタイル部において観察された時期とほぼ対応して境界層遷移が見られ.た。胴体下面のセラミックタイル部においては境界層遷移および再層流化がみられた。無限斜め円柱付着線上遷移判定基準をもとに気流条件を評価した結果、セラミックタイルギャップ・ステップによる表面粗さ、迎角変化などを考慮すれば妥当な範囲で遷移及び再層流化が起こっていることがわかった。またその他のセラミックタイル部における空力加熱データから層流時における加熱率はおよそ予測値と一致した結果が得られたが、胴体下面前方付近ではC/Cノーズ部で見られた実在気体効果と考えられる現象と類似した低加熱の領域が観察された。, Aerodynamic heating on the Hypersonic Flight Experiment (HYFLEX) vehicle was measured using newly developed sensors, calibrated by lamp heating tests. These sensors were shown to be of use for flight measurement purposes. The results of the aerodynamic heating measurements on the vehicle were compared with predictions based on cold hypersonic wind tunnel tests and CFD (Computational Fluid Dynamics) calculations. The flight results on the nose cap were lower than the predictions based on the wind tunnel tests in the stagnation region of around t = 100 sec. Comparison with real gas CFD results assuming a fully catalytic wall and non-catalytic wall indicate that this discrepancy is due to the effect of dissociation of the air, and can be explained by the dependence of catalycity on surface temperature. For the measurements on the elevon, an increase in aerodynamic heating caused by boundary layer transition was observed immediately prior to it being observed on the windward fuselage. Flight data on the windward side of the afterbody show boundary layer transition and following laminarization, which may be explained qualitatively by the transition criteria for the attachment line of the yawed cylinder, taking into account the effect of nose bluntness., 資料番号: AA0028660000, レポート番号: NAL TR-1415

Navier-Stokesコード、線形放物型安定方程式(PSE)に基づく2種類のコード、およびe(exp N)法コードからなる境界層遷移予測システムが開発された。予測システムは一般曲線座標で安定化されており、実用形状への適用が可能である。このシステムを用いて、超音速流中に置かれた無限後退円柱の境界層の遷移予測が行われた。得られた結果からシステムの正常な動作が確認され、また境界層の非平行性が斜めTS(Tollmien-Schlichting)波に与える影響も明らかになった。さらに、横流れ擾乱に対する解析によって、PSEコードがトランジェットや振動といった特性を持つことがわかった。, A system for prediction of boundary-layer transition is developed which consists of the Navier-Stokes code, two kinds of codes based on the linear Parabolized Stability Equations (PSE), and the e(exp N)-method code. The prediction system is formulated in generalized curvilinear coordinates, which makes the system applicable to practical configurations. The system is used to predict the transition of the boundary layer of an infinite swept cylinder placed in a supersonic flow. The results obtained validate the system and clarify the effect of the nonparallelism of the boundary layer on an oblique TS (Tollmien-Schlichting) wave. Furthermore, the characteristics of the PSE codes such as transients and oscillations are found through the analysis of a crossflow disturbance., 資料番号: AA0028648000, レポート番号: NAL TR-1397T

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, チャンネル流の層流-乱流遷移を並列計算機上で直接数値シミュレーションを行った。シミュレーションは、基本流と様々なレイノルズ数に対応する異なるエネルギーレベルの小さな撹乱の和で与えられる初期流れで開始する。超臨界レイノルズ数の場合、以下の2つの機構により遷移が起こる。ひとつは初期撹乱が非常に小さい場合、流れの中で成長するトルミーン・シュリヒティング波であり、もうひとつは初期撹乱が大きい場合、流れの方向に軸を持つ縦渦の過渡的増幅である。一方、亜臨界レイノルズ数では、初期撹乱が非常に大きい場合にのみ過渡的増幅により遷移が起こる。過渡的増幅型遷移を引き起こす初期撹乱のエネルギーレベルは、レイノルズ数の-7/2乗に比例することが見出された。, Laminar-turbulent transition of channel flow is directly simulated on a parallel computer. The simulations are started with initial flows given as the basic flow plus small disturbances with different energy levels for various Reynolds numbers. In the case of supercritical Reynolds numbers, the transition is triggered by two mechanisms: one is Tollmien-Schlichting waves growing in the flow when initial disturbances are very small, and the other is transient growth of stream-wise vortices when initial disturbances are rather large. On the other hand, for subcritical Reynolds numbers, transition is triggered by the transient growth only when initial disturbances are considerably large. It is found that the energy levels of initial disturbances, which can trigger the transient-growth type transition, are proportional to -7/2 power of the Reynolds number., 資料番号: AA0001961054, レポート番号: NAL SP-44

航空宇宙技術研究所 7-9 Oct. 1998 (23th). 24-26 Mar. 1999 (24th) 東京 日本, National Aerospace Laboratory 7-9 Oct. 1998 (23th). 24-26 Mar. 1999 (24th) Tokyo Japan, 以下の主題について論じた。チャネル流遷移、3次元不安定性、衝撃自由剪断層不安定性、回転ゲルトラー流れ、e(exp N)法、斜め円柱境界層、縮流胴境界層遷移、回転円盤流、横流れ渦、絶対不安定性、圧縮性縦渦、超音速剪断層、衝撃波-境界層干渉、超音速境界層、小型超音速実験機、内部渦構造、2次渦、表面圧力変動、空音力強度、乱流境界層、円柱渦、剥離制御、等方性乱流、摩擦抵抗低減、小型超音速風洞、および強制磁気再結合。, The following topics are discussed: channel flow transition, three dimensional instability, impinging free shear layer instability, rotating Goertler flow, e(exp N) method, yawed cylinder boundary layer, laminar-turbulent transition, rotating disk flow, cross flow vortex, absolute instability, compressive streamwise vortex, supersonic shear layer, shock-boundary layer interaction, supersonic boundary layer, scaled supersonic experimental airplane, internal vortical structure, secondary vortex, surface pressure fluctuation, acoustic intensity, turbulent boundary layer flow, columnar vortex, control of separation, isotropic turbulence, friction drag reduction, supersonic wind tunnel, and forced magnetic reconnection., 資料番号: AA0001960000, レポート番号: NAL SP-43