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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