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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., 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 present study is engaged in acoustic excitation of the S-C (Streamline-Curvature) mode through the annular slit. A 0.8 m diameter steel flat disk and its rotation speed were redesigned in such a way that the S-C mode can be directly excited. To introduce unsteady disturbances into the disk boundary layer through the slit, a loud speaker was installed underneath the disk. This redesign allowed the maximum local Reynolds number of 124 at the slit. A multi-purpose spectrum analyzer with the dual channel inputs was used for acquiring and analyzing all data. Several conclusions are drawn from the experimental results. (1) Unsteady disturbance due to S-C instability was successfully amplified by artificial forcing through an annular slit on a rotating disk. (2) The S-C mode had zero wavenumber in the azimuthal direction and its critical Reynolds number was approximately 120. (3) Excitation through an annular slit was not suitable for realization of the most unstable S-C mode with the critical Reynolds number of 63. (4) The range of unstable frequencies for the zero azimuthal wavenumber as well as the radial wavenumber, and the total growth rate for the most unstable annular disturbance showed excellent agreement with theoretical predictions., 本の研究は、環状スリットを通した流線曲率(S-C)モードの音響励起に関するものである。直径0.8mの鋼製の平板ディスクおよび回転速度を再検討し、S-Cモードが直接励起されるようにした。スリットを通して非定常擾乱を誘因するために、ラウドスピーカをディスクの下に置いた。この再設計によって、最大の局所レイノルズ数はスリットで124となった。データの取得と解析には、2チャンネル入力の多目的スペクトルアナライザを用いた。実験の結果、いくつかの結論が導き出された。(1)回転盤上の環状スリットを通して、S-C不安定性による非定常擾乱は人工的にうまく増幅された。(2)S-Cモードは、方位角方向での波数がゼロ、レイノルズ数は約120であった。(3)環状スリットを通した励起は、臨界レイノルズ数63の最も不安定なS-Cモードの実現には適していなかった。(4)ゼロの方位角方向波数および動径方向波数の不安定周波数の範囲、および最も不安定な環状擾乱の全成長率は、理論による予測と非常に良い一致をみた。

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., Special Publication of National Aerospace Laboratory, 航空宇宙技術研究所特別資料

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.

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., Original contains color illustrations, Meeting Information: 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), 形態: カラー図版あり, 会議: 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

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"., Meeting Information: 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), 会議: 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

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., 会議: 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）

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 JAXA-ONERA cooperative research project started in April 2000, because ONERA had great ability of analyzing transition phenomena and JAXA 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, good cross validation of both ONERA's and JAXA 's e(N) codes was obtained. There was also high correlation between the experimental results in ONERA's continuous circuit-flow type supersonic wind tunnel and both laboratories' predictions under the assumption of a critical transition N value of 6. In addition, a risk of transition due to attachment-line contamination was predicted at inner wing, using the well-known Poll's criterion., Original contains color illustrations, 形態: カラー図版あり

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, 会議: 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）