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In supersonic wind tunnels, the humidity content of the working fluid is known to have a significant effect on flow conditions in the test section. The intent of this paper is to evaluate the influence of moisture on the test section flow quality for the indraft supersonic wind tunnel located at the Muroran Institute of Technology. The static-pressure fluctuation is measured by means of an unsteady pressure sensor buried in a 10-degree-aperture cone model installed in the test section. The results show that supersonic flow at Mach 2 is contaminated by the self-sustained oscillation of a condensation shock wave for ambient relative humidity greater than approximately 50%. It is also found that for the Mach 2 flow, that the ratio of the static-pressure fluctuation to the dynamic pressure is less than 0.1%, if the absolute humidity is kept below a critical value of 2[g/m3]. Experimental observations also reveal that the three-dimensional boundary-layer transition process is hypersensitive to the influence of free-stream humidity., 特集 : 航空宇宙機システム研究センターにおける開発研究2 : 4. オオワシII飛行にむけて並びに関連技術の研究開発（試験設備系）

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県, Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, From the Mach 11 flight condition experiment result at HIEST, it was found that it was necessary to promote mixing to raise combustion efficiency. Therefore, the mixture state in the engine was investigated. The experiment conducted at supersonic wind tunnel. The mixture state was evaluated by gas sampling by the experiment without combustion. As a result, mixing promotion effect is match higher with the rocket side auxiliary injection than the cowl side auxiliary injection. Additionally, a mixture promotion effect by the incidence of the expansion wave to rocket exhaust was confirmed., 形態: カラー図版あり, 形態: PDF, Physical characteristics: Original contains color illustrations, Physical characteristics: PDF, 資料番号: AA0061856035, レポート番号: STCP-2012-035

平成27年度衝撃波シンポジウム (2016年3月7日-9日. 熊本大学 黒髪南地区）熊本市, 熊本県, Symposium on Shock Waves in Japan, FY2015 (March 7-9, 2016. Kumamoto University), Kurokami Chuo-ku, Kumamoto, Japan, 資料番号: SA6000049059, レポート番号: 2A1-5【若】

第41回「境界層遷移の解明と制御」研究会（2007年10月4日-10月5日）, Instability of the wake behind a plate is investigated theoretically with the complex ray theory. Three kinds of velocity distributions are comparatively used to evaluate effects of the basic-flow approximation on stability characteristics of the wake. Streamwise variation of the maximum velocity defect is approximately described by using a parabolic bridge between the value at the trailing edge and those of the far wake. Computations show the frequency selection through absolute-global instability as well as growth and decay of the global mode far downstream., 資料番号: AA0064237001, レポート番号: JAXA-SP-08-006

It is well known that acoustic noise emanated from an airfoil trailing edge is discrete. The mechanism on why the noise becomes discrete and how the frequency is selected remains unclear. To reveal the mechanism, the boundary layer near the trailing edge on the pressure side was artificially disturbed with the use of distributed roughness elements. As a result, the separation is swept away and an acoustic sound is suppressed. The experimental results indicate that the frequency of trailing-edge noise is determined in separation region where mean velocity distribution has an inflection point accompanying reverse flow. These are necessary conditions for absolute instability., 資料番号: AA0063908019, レポート番号: JAXA-SP-07-026E

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

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

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

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

Transition on the natural laminar flow wing of a non-powered scaled supersonic experimental airplane was measured by hot-film sensors and an infrared camera at Mach 2. In order to supress crossflow instability which dominates the transition on the highly swept wing, the pressure falls very rapidly in the vicinity of leading edge of the wing, so as to minimize the crossflow velocity components. The unit Reynolds number had no effect on the transition Reynolds numbers of the wing. The transition Reynolds number was 0.8 million at 70 percent semi-spanwise location at Mach 2 at the angle of attack of 2 deg. Transition location in flight was estimated by conducting transition measurement in a quiet low Reynolds number wind tunnel and in a variable unit Reynolds number wind tunnel which has moderately low turbulence. The transition at 70 percent semi-spanwise position at the angle of attack of 2 deg at Mach 2 in flight was predicted to be located downstream of 60 percent chordwise position., 資料番号: AA0032827008, レポート番号: NAL SP-54

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

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, 超音速輸送機を目的とした無推力実験機の自然層流翼上の遷移測定をマッハ数2.0と1.2で行った。翼弦沿いに設置した熱フィルム流速計に対する出力信号の振幅中の摂動増加が層流から乱流境界層への遷移を示す。熱フィルム流速計およびPreston管による測定結果は遷移点が翼の半弦の下流に位置していることを明らかにした。, Transition measurement on the natural laminar flow wing of non-powered experimental airplane for supersonic transport was conducted at Mach 2.0 and 1.2. Increase of perturbation in amplitude of output signals for hot-films placed along the wing chord indicates transition from laminar to turbulent boundary layer. Results by hot-film and Preston tubes proved that transition point is located downstream of the half chord of the wing., 資料番号: AA0028637019, レポート番号: NAL SP-48T

第4回学際領域における分子シメージングフォーラム (2008年10月9日. 早稲田大学), 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064305005, レポート番号: JAXA-SP-08-012

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, 流線曲率モードより横流れ進行モードを分離するため、円盤上の孔から点源撹乱を導入した静止空気中の回転円盤流の実験的研究を行った。実験結果は、両方のモードの明確な分離が、点源の下流の楔状領域で、分散関係に従って現れることを示していた。また、増幅していく撹乱の特性については、複素特性曲線法を用いた線形安定性理論とよく一致した。本論文の最後では、斜め円柱の場合と比較として、なぜ回転円盤流や後退翼では、横流れ不安定の定在モードが支配的となりがちなのかを議論した。, An experimental investigation in rotating-disk flow in still air was made to separate the crossflow traveling mode from the streamline-curvature mode introducing a point-source disturbance through a hole on the disk. The experimental results show that the definite separation of both modes appeared in wedge-shaped region downstream of the point source in accordance with the dispersion relation. It is also shown that the characteristic features of growing disturbances are in good agreement with the linear stability theory making use of the method of complex characteristics. In the last part of this paper, the reason is discussed why the stationary mode of the crossflow instability is prone to dominant appearance on rotating disk and swept wing compared to yawed circular cylinder., 資料番号: AA0001960006, レポート番号: NAL SP-43

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県, Space Transportation Symposium FY2013 (January 16-17, 2014. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000016041, レポート番号: STCP-2013-041

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県, Space Transportation Symposium FY2013 (January 16-17, 2014. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000016026, レポート番号: STCP-2013-026

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県, Space Transportation Symposium FY2013 (January 16-17, 2014. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000016042, レポート番号: STCP-2013-042

航空宇宙技術研究所 26-28 Mar. 1997 東京 日本, National Aerospace Laboratory 26-28 Mar. 1997 Tokyo Japan, 斜め円柱における3次元境界層の不安定性を実験的に検討した。流れの不安定特性を明らかにするため円柱上の点源から人為的に非定常撹乱を導入した。横流れ不安定の定常モードの影響により振幅は変調するが、レイノルズ数が線形安定理論の予測と一致するある一定の範囲内では、流線曲率不安定が横流れよりも支配的であることを確認した。, Instability of the three-dimensional boundary layer on a yawed circular cylinder is investigated experimentally. An unsteady disturbance is artificially introduced from a point source on the cylinder to clarify instability characteristics of the flow. Streamline-curvature instability is confirmed to dominate rather than cross-flow instability in a certain range of Reynolds number in agreement with the prediction of linear stability theory, although the amplitude is modulated by the effect of stationary mode of cross-flow instability., 資料番号: AA0001320001, レポート番号: NAL SP-36

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県, Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, With the aims of creating and validating innovative fundamental technologies for high-speed atmospheric flights, a small-scale supersonic experimental vehicle is being developed as a flying test bed. The configuration M2006 with a cranked-arrow main wing and twin engines was designed and its aerodynamic characteristics were verified through wind-tunnel and flight tests. A new configuration M2011 with a main wing and tails of similar shapes and a single GG-ATR engine was proposed. Intensive wind-tunnel tests were carried out in transonic and supersonic regimes. It was found that the overall drag coefficient increases as the intake passage is throttled and that the airflow through the intake seems to be squeezed by the development of the boundary layer. In addition, the tail steering box located at the vertical tail root has little influence on overall drag., 形態: カラー図版あり, 形態: PDF, Physical characteristics: Original contains color illustrations, Physical characteristics: PDF, 資料番号: AA0061856059, レポート番号: STCP-2012-059

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県, Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 形態: カラー図版あり, 形態: PDF, Physical characteristics: Original contains color illustrations, Physical characteristics: PDF, Translation title: Design of 2nd generation airframe of supersonic unmanned plane developed at Muroran Institute of Technology, 資料番号: AA0061856058, レポート番号: STCP-2012-058

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県, Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, The purpose of this study is to accomplish combustion in RBCC engine operating in hypersonic flow at Mach 11 flight condition. Auxiliary injection for combustion enhancement in RBCC engine was tested experimental at JAXA-Kakuda Space Center using high enthalpy shock tunnel (HIEST). By injecting a portion of fuel besides in residual fuel in the rocket exhaust, combustion of residual hydrogen in the rocket exhaust was attained., 形態: カラー図版あり, 形態: PDF, Physical characteristics: Original contains color illustrations, Physical characteristics: PDF, 資料番号: AA0061856040, レポート番号: STCP-2012-040

第43回（2008年10月2日-10月3日）「境界層遷移の解明と制御」研究会; 第44回（2009年3月26日-3月27日）「境界層遷移の解明と制御」研究会, 43rd JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (October 2-3, 2008) ; 44th JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (March 26-27, 2009), The compressible flows around 2-D airfoils are simulated by using a high-order numerical scheme. In this study, the generation of trailing-edge noise is directly reproduced as the fluctuation in pressure field. Two types of airfoils are investigated: NACA0012 and NACA0006. The obtained results on the vortex-shedding frequencies and the growth rate of velocity fluctuation confirm that the tone noise frequency is determined by the wake instability in both cases. The dipole sound scattered by the acoustically non-compact airfoils shows multiple lobes in the far-field directivity. By imposing non-zero angle of attack, very weak instability waves has little effects on the development of Karman vortex shedding in the present calculation., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064575018, レポート番号: JAXA-SP-09-014

第43回（2008年10月2日-10月3日）「境界層遷移の解明と制御」研究会; 第44回（2009年3月26日-3月27日）「境界層遷移の解明と制御」研究会, 43rd JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (October 2-3, 2008) ; 44th JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (March 26-27, 2009), Aerodynamic noise generated from 2-D airfoils is of interest in conjunction with the development of Karman-vortex shedding in the wake region. In this study, the computational aspects are discussed to reproduce the trailing-edge noise using acoustic analogies, based on the flow field obtained by a compressible flow solver with higher-order numerical schemes. The comparative study indicates that less-expensive acoustic analogies fail in representing the acoustically non-compact sound source with incompressible flow field However, by using a compressible near-field solution, the FW-H equation successfully predicts the far-field sound pressure directivity, for an accurate pressure fluctuation is provided at the airfoil surface., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064575001, レポート番号: JAXA-SP-09-014

第43回（2008年10月2日-10月3日）「境界層遷移の解明と制御」研究会; 第44回（2009年3月26日-3月27日）「境界層遷移の解明と制御」研究会, 43rd JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (October 2-3, 2008) ; 44th JAXA Workshops on "Investigation and Control of Boundar-Layer Transition" (March 26-27, 2009), This brief note presents my research activities since I moved in the National Aerospace Laboratory as a researcher in 1979. During this period, my study was consistently made on the impregnable philosophy of "the basic research leading to promising applied research and the applied research developing on basic research"., 資料番号: AA0064575019, レポート番号: JAXA-SP-09-014

平成21年度宇宙輸送シンポジウム(2010年1月14日-15日, 宇宙航空研究開発機構宇宙科学研究本部), Space Transportation Symposium (January 14-15, 2010. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)), Sagamihara, Kanagawa Japan, 形態: CD-ROM1枚, Note: One CD-ROM, 資料番号: AA0064741020

第41回「境界層遷移の解明と制御」研究会（2007年10月4日-10月5日）, The trailing-edge noise from 2-Dimensional airfoil is experimentally investigated to reveal the mechanism of the frequency selection. The splitter plate is placed at the trailing-edge of the airfoil to produce the no sound emitted condition. At this condition, the instability wave with broad band spectrum which central frequency almost coincides the tonal noise frequency is clearly observed on the pressure side. The artificial acoustic disturbances which are the feedback of the instability wave on the airfoil are introduced at no sound emitted condition, and it is observed that the broad-band spectrum become discrete soon after the feedback is excited. These results confirmed that the trailing-edge noise is the consequence of the acoustic feedback loop between a vortex shedding at the trailing edge and an instability wave on the pressure side of the airfoil., 形態: カラー図版あり, 資料番号: AA0064237013, レポート番号: JAXA-SP-08-006

第42回「境界層遷移の解明と制御」研究会（2008年3月31日-4月1日）, The instability of axisymmetric laminar wake behind a body of revolution whose cross section is a NACA0018 airfoil is studied experimentally under a natural disturbance condition. Magnetic Suspension and Balance Systems (MSBS) are used to support the axisymmetric body in order to avoid undesirable influences of mechanical supports on the disturbance development. A multi-hotwire-sensor probe is used to identify helical instability modes. The experiment is conducted mainly at a Reynolds number based on the maximum diameter Re = 1.9 × 10(exp 4). It is found that the wake is convectively unstable despite that the flow is slightly reversed near the trailing edge. Spatially-growing instability waves are found to be helical modes with azimuthal wavenumber of 1 as predicted by the linear stability theory for the axisymmetric wake. The spatial growth rate and amplitude distribution obtained experimentally agree well with those calculated from the linear stability theory., 形態: カラー図版あり, 資料番号: AA0064237015, レポート番号: JAXA-SP-08-006

資料番号: NALSP0015021, レポート番号: NAL SP-15

The three-dimensional boundary-layer experiments were conducted on a 45 degrees swept wing in the ASU unsteady Wind Tunnel. Crossflow-dominated transition was produced on a very smooth surface with contoured end liners to simulate infinite swept-wing flow. The model surface was carefully polished and buffed up to plus or minus 3 mum to minimize roughness effects on transition. Fixed-wavelength vortices, traveling wave and inflectional breakdown were subsequently observed prior to turbulent state as before, but the transition location moved further 10%-20% downstream of the chordlength. In order to control traveling wave and inflectional breakdown, two kinds of acoustic sounds, which tune up the traveling wave and breakdown, respectively, were introduced visualizing the surface flow via two-or three-dimensional roughness elements. No acoustic effect was observed in visualization., 資料番号: NALSP0015020, レポート番号: NAL SP-15

資料番号: NALSP0012017, レポート番号: NAL SP-12

資料番号: NALSP0011010, レポート番号: NAL SP-11

資料番号: NALSP0011023, レポート番号: NAL SP-11

e(sup N)法を用いた2次元および3次元境界層の乱流遷移位置の予測において外乱の影響を定量的に評価するため、一様流の残留乱れTuおよび騒音レベルSPLと遷移位置における撹乱の空間増幅率Nの相関関係を実験的に調べた。気流特性が異なる3つの風洞で同一の矩形翼NACA0012模型の遷移計測を実施した。その結果、模型を2次元と3次元形態のどちらに設置しても、NはTuの対数とSPLの1次関数で表されるが、外乱に対する感度は翼断面や後退角に依存することがわかった。この相関関係を用いることにより2次元の場合には遷移位置の予測精度が翼弦長の10％程度向上することが期待される。また低亜音速境界層の遷移位置検出手法として、赤外線力メラを用いた表面温度分布計測が有効であり、横流れ定在渦も可視化可能であることを示した。, A series of experimental investigations have been carried out to examine the influence of free-stream turbulence and acoustic noise on laminar-to-turbulent transition in boundary layers on an NACA0012 airfoil. The quantitative evaluation of the influence on the total growth rate N of the most amplified disturbance obtained from the e(sup N) method at the transition location is made in an empirical relation between the total growth rate N and both the residual turbulence Tu and the sound pressure level SPL. Consequently, the transition location is detected as a linear equation N = C(sub Tu) x ln(Tu) + C(sub SPL) x SPL + C for not only 2-D case with zero angles of both sweep and attack but also 3-D case with non-zero angles of both sweep and attack, though each coefficient in N depends on geometrical configuration of airfoil such as cross-sectional shape and swept angle. Accuracy of predicted chord length is expected to be improved 10 percent by use of the obtained empirical relation. Surface temperature measurement technique by use of infrared camera is shown to be efficient and useful for transition detection in low subsonic flow. Usefulness is also shown in visualizing cross-flow stationary vortices growing in 3-D boundary layer., 資料番号: AA0049194000, レポート番号: JAXA-RR-05-017

This paper studies the possibility of strategy-proof rules yielding satisfactory solutions to matching problems. Alcalde and Barberá (1994) show that effcient and individually rational matching rules are manipulable in the one-to-one matching model. We pursue the possibility of strategy-proof matching rules by relaxing effciency to the weaker condition of respect for unanimity. Our first result is positive. We prove that a strategy-proof rule exists that is individually rational and respects unanimity. However, this rule is unreasonable in the sense that a pair of agents who are the best for each other are matched on only rare occasions. In order to explore the possibility of better matching rules, we introduce the natural condition of 'respect for pairwise unanimity.' Respect for pairwise unanimity states that a pair of agents who are the best for each other should be matched, and an agent wishing to stay single should stay single. Our second result is negative. We prove that no strategy-proof rule exists that respects pairwise unanimity. This result implies Roth (1982) showing that stable rules are manipulable. We then extend this to the many-to-one matching model.

線形安定理論は、3次元境界層における横流れ不安定に対して時間依存型モードは定在モードに比べてより不安定であると予測している。3次元境界層の典型的な流れ場である静止流体中で回転する円盤流では定在モードが進行波モードに比べて優勢的に観察されているけれども、その理由については合理的な説明がなされていない。本論文では、これまでの実験的観察と理論的研究のリビューに加えて、壁面曲率項、流線曲率項および非平行性を含めた安定計算を実施して、その選択機構について議論している。この結果、壁面曲率項は横流れ不安定のモード選択機構に重要な役割を演じる、と同時に撹乱の初期振幅も重要であることが明らかにされている。, Linear stability theory predicts that the most unstable disturbance of crossflow instability in three-dimensional boundary layers is a time dependent mode rather than a stationary mode. Stationary rather than traveling mode has been predominantly observed in natural disk flow, spinning in still air, which is known to be a typical example of three-dimensional boundary layers. The present paper discusses this mode selection mechanism of crossflow instability on the basis of reviewing both experimental and theoretical research and making linear stability calculations inclusive of a wall-curvature term as well as streamline-curvature and non-parallel terms. This reveals that the wall curvature plays a key role in the mode selection mechanism of three-dimensional boundary layers and that initial amplitudes of unstable disturbances are of principal importance., ISSNの誤記: 1349-1148, Incorrect ISSN: 1349-1148, 資料番号: AA0047414000, レポート番号: JAXA-RR-03-022E

航空宇宙技術研究所 Sep. 1994 (15th). Mar. 1995 (16th) 東京 日本, National Aerospace Laboratory Sep. 1994 (15th). Mar. 1995 (16th) Tokyo Japan, 3次元境界層に固有な横流れ不安定場での点状連続攪乱の形成と成長を実験的に検討した。攪乱は、小さな孔から境界層流中に連続的に空気を噴出する空気噴流柱により、人工的に起こした。熱線測定の結果、静的モードも進行モードも楔形になる。これは、進行波が分散しながら正常方向にも局部外向流の方向にも伝搬するからである。, The formation and development of a point-source disturbance in a three-dimensional boundary layer subject to crossflow instability has been investigated experimentally. The disturbance was artificially initiated by an airjet column, which was continuously issued into the boundary layer flow through a small hole. Hot-wire surveys showed that both stationary and traveling modes evolve into a wedge shape as the traveling waves dispersively propagate in both directions approximately normal to local external flow., 資料番号: AA0000109012, レポート番号: NAL SP-28

低速風洞縮流胴に沿った境界層の遷移過程を可視化および熱線計測などにより実験的に研究した。縮流胴境界層の遷移過程は次のように幾つかの素過程に分けられる。レイノルズ数が低い場合、凹面領域にはゲルトラー不安定により、流れ方向に軸を持ち対になって回転する縦渦列が発生する。層流境界層はその縦渦列によって歪められて変曲点を持ち、2次不安定によって乱流へと遷移する。しかし、下流の凸面領域に至ると主流の加速によって境界層は再層流化され、縮流胴出口まで層流状態が維持される。レイノルズ数が高い場合、縮流胴入口における初期境界層はすでに乱流状態にあり、凹面領域を通過する間に境界層厚さはさらに増加するが、その下流の凸面領域では一旦層流化する。しかし、縮流胴出口付近において乱流斑点が自然発生することにより、境界層は再遷移する。測定部の主流流速を増加すると乱流斑点が発生する位置は上流側のより加速された領域へと移動する。乱流斑点は加速流中でも下流に向かって成長するので、レイノルズ数が高い場合には測定部の境界層が乱流になるのは避けられない。, The transition process of boundary layers on a low-speed wind-tunnel contraction wall was experimentally investigated using flow visualizations and hot-wire/hot-film measurements. The process has several classifications depending on the Reynolds number. For the low Reynolds number, a row of counter-rotating longitudinal vortex pairs is generated by the Gortler instability on the concave wall. The velocity profile of the laminar boundary layer distorted by the Gortler vortices has inflection points, and eventually the secondary high-shear-layer instability causes high-frequency velocity fluctuations. These growing disturbances lead the boundary layer to a turbulent state within the concave region of the contraction. However, the resultant turbulent boundary layer is laminarized by acceleration of the mean flow in the convex region free from centrifugal instability. For the higher Reynolds number, the initial boundary layer flow was already turbulent at the entrance of the contraction. The boundary-layer thickness increases downstream within the concave region. On the convex wall the turbulent boundary layer is laminarized; however, re-transition occurred with the naturally occurring turbulent spots under a favorable pressure gradient. With an increase in the mean velocity in the test section, the location of occurrence of spots moves forward into a more accelerated region. As the turbulent spots grow downstream even in the accelerated flow, the formation of a turbulent boundary layer is unavoidable on the test section wall at the higher Reynolds number., 資料番号: AA0046106000, レポート番号: NAL TR-1462

小型超音速ロケット実験機の自然層流翼の遷移特性をホットフィルムと赤外線カメラを用いて計測した。高後退翼の遷移を支配する横流れ不安定の発達を抑制するために、自然層流翼では前縁付近の急加速領域を狭くすることで横流れ速度成分を抑えている。本翼の遷移Reynolds数は単位Reynolds数に依存せず、設計迎角(α=2deg)において70％半翼幅位置における遷移Reynolds数は0.8 x 10(exp 6)であることがわかった。線形安定解析結果から本翼を支配する不安定性はTollmien-Schlichting不安定性であり、横流れ不安定は設計通り抑制されていることがわかった。表面粗さと主流乱れの影響を考慮すると、飛行試験 M=2、α=2deg (設計迎角)において70％半翼幅位置の遷移位置は約50％翼拡張と推定された。, Transition characteristics on the natural laminar flow wing of a nonpowered scaled supersonic experimental airplane was measured by hot film sensors and infrared camera in two supersonic wind tunnels at Mach 2. In order to suppress the growth of crossflow instability which dominates the transition on highly swept wings, the pressure falls very rapidly in the vicinity of leading edge of the wing, so as to minimize the crossflow velocity components. From test results, the unit Reynolds number had no effect on the transition Reynolds number of the wing. The transition Reynolds number was 0.8 million at 70 percent semispanwise location at Mach 2 at the angle of attack of 2 deg. Tollmien-Schlichting instability dominates the transition and crossflow instability was found to be suppressed on the wing. Considering the influence of surface roughness and free-stream turbulence, the transition at 70 percent semi-spanwise position at Mach at the angle of attack of 2 deg in flight is estimated to be located at 50 percent chordwise position., 資料番号: AA0032998000, レポート番号: NAL TR-1437

航空宇宙技術研究所 28-29 Sep. 1992. 25-26 Apr.1993 東京 日本, National Aerospace Laboratory 28-29 Sep. 1992. 25-26 Apr.1993 Tokyo Japan, 主流に対して斜めに置かれた円柱に沿う3次元境界層の横流れ不安定について実験的に調べた。これまで長い間、根源的な擾乱と考えられていた定常的な渦の代わりに、はっきりした振動数をもつ時間進行波が熱線風速計による不安定流れ場の計測のうちに観察された。熱線風速計の配列を用いて詳細な計測を行い、この非定常擾乱が、線形安定理論によって最も不安定な擾乱と予言されている横流れ進行波と同一のものであることを確認した。しかしながら、理論予測と実験結果の間にはなお一致しない点が残っている。, Experimental investigations are made on crossflow instability of the three dimensional boundary layer along a yawed circular cylinder. Unsteady disturbances with definite frequencies instead of stationary vortices, which have long been considered as generic disturbances, are observed in hot-wire surveys of the unstable flow field. Detailed measurements with the aid of hot-wire arrays have confirmed that the unsteady disturbances are identical with the so-called crossflow traveling waves predicted by linear stability theory as the most unstable disturbances, although some discrepancies still remain between experimental results and theoretical predictions., 資料番号: AA0004170010, レポート番号: NAL SP-21

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combinedsystem termed as Rocket Based Combined Cycle engine) are expected to be the most effective propulsion system forBooster stage of space launch vehicles. At hypersonic regime, it will be operated at rather high rocket engine output forfinal acceleration with some Isp gains due to air-breathing effects. In this regime, attaining thrust at this high-speedregime becomes very difficult, so that parallel injection of the fuel for scramjet combustion is favorable as the momentumof the injection can contribute to the thrust production. Thus, embedded rocket chamber was supposed to the operated asfuel rich gas generator at very high output. This configuration was tested at simulated flight Mach number of 11 at HighEnthalpy Shock Tunnel (HIEST) with detonation tube as the source of the simulated rocket exhaust. However,combustion of the residual fuel in the rocket exhaust with airflow could not be attained in the past study, so thatmodification was carried out. Combustion was attained within the combustor, however, insufficient mixing was found toresponsible for the rather low combustion efficiency, finding being from complementary mixing test at a blow-down typetunnel., 資料番号: PA1410085000

The generation of tonal noise by airfoils has been a continuing problem of which the overall mechanism still remains yet to be uncovered. Results of an experimental investigation on tonal noise generation are presented. Quantitative measurements were obtained by use of hotwire and microphone. Furthermore, a flow visualization of the phenomenon was carried out by use of a smoke wire technique. Based on these experiments, a detailed description of the flow close to the trailing edge is obtained. The sound generation process takes place on the pressure side in the near-trailing edge region, where a thin layer of separated flow exists, which is essential to the mechanism. A Kelvin-Helmholtz instability appears to be the cause of the vortex generation. The interaction of shedding eddies with the trailing edge, resulting in the emission of pressure waves, represents the noise source. Events on the suction side are found to be of no significant influence to the sound generation mechanism. The role of TS (Tollmien Schlichting) waves in the sound generation process is re-evaluated and a novel self-excited feedback loop is proposed to explain the occurrence of tonal noise generation., 資料番号: AA0063908008, レポート番号: JAXA-SP-07-026E

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

A series of experimental investigations have been carried out to examine the influence of free-stream turbulence and acoustic noise on laminar-to-turbulent transition in rectangular wing flow. The quantitative examination deduces an empirical relation between the total growth rate N for the transition prediction obtained from the e(sup N) method and both residual vortical and non-vortical fluctuations. Also, we propose a modified relation, which takes account for both external fluctuations with effective contribution to transition., 資料番号: AA0049208026, レポート番号: JAXA-SP-05-012

Measurement system of supersonic boundary-layer transition for the SST experimental airplane (NEXST-1) is evaluated for successful capture of transition signature prior to the flight test. Output levels of hot-film sensors and unsteady pressure transducers at the flight test are estimated on the basis of experimental results obtained in ground facilities and are found to be significantly above electric noise level in transitional state. Estimation shows that transition point with the maximum values of shear-stress and pressure fluctuations in the chordwise direction is inferred to be detectable., 資料番号: AA0047776010, レポート番号: JAXA-SP-04-002

Characteristic features of a wave packet in a rotating-disk boundary layer were experimentally investigated. The point disturbance originating from a small hole on the disk forms an isolated wave packet which consists of two types of unstable waves. Comparison with the prediction by Itoh's linear stability theory enabled the identification of these waves; the streamline-curvature (S-C) modes with the smaller wave number in the circumferential direction and the higher frequency are observed prior to the appearance of cross-flow (C-F) modes with the larger wave number and lower frequency. The hot-wire analyses quantitatively show that such characteristic features of these modes are in good agreement with the theoretical results. Time trace of the edge of packet does not indicate the occurrence of absolute instability but shows that the wave packet convectively passes through the radius line corresponding to the critical Reynolds number for the absolute instability predicted by Lingwood., 資料番号: AA0045947029, レポート番号: NAL SP-56

Behavior of stationary vortices in rotating-disk boundary layer was experimentally investigated by hot-wire measurements. Sources of disturbance such as the disk vibration and free convection around the disk seem to affect the unsteadiness of the stationary vortices. The circumferential wave numbers and the position of vortices fluctuate for each rotation. The fluctuation was quantitatively evaluated. Small roughness elements have the effect of lowering the fluctuation although the nonlinear growth of the vortices becomes significant. This fact implies that the development of nonlinearity is not main cause of the unsteady behavior of the stationary vortices., 資料番号: AA0045947016, レポート番号: NAL SP-56

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, 最近の線形理論は、付着線上の流れに比較して極低レイノルズ数時に流れが不安定になる領域では多重不安定性が共存することを示している。不安定性が臨界レイノルズ数の低下に特に作用する現象を明らかにするため、付着線近傍の境界層不安定性に焦点を当てて実験的な検討を行った。これらの実験結果を理論計算と比較した。, The latest linear theory shows that multiple instabilities co-exist in the region, where the flow becomes unstable at very low Reynolds number compared to that on the attachment line. An experimental investigation has been done focusing boundary-layer instability near the attachment line to clarify the phenomena, especially on which instability devotes to the drop of the critical Reynolds number. These experimental results are compared with the calculations done by the theory., 資料番号: AA0028637009, レポート番号: NAL SP-48T