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NEXST-1プロジェクトは, 1997年から2005年にかけて実施した小型超音速実験機プロジェクトで, 次世代超音速旅客機（SST）の抵抗低減に関する先進技術の飛行実証を目的として実施された. 本プロジェクトでは, 圧力抵抗低減に寄与する主翼平面形とWarp形状による揚力依存抵抗の低減効果と, 翼胴干渉を考慮したエリアルール胴体設計による体積依存造波抵抗の低減効果を確認すると同時に, 世界初となる亜音速前縁型主翼の自然層流翼設計技術による摩擦抵抗の低減効果も飛行実証された. 本飛行実験データは, 実験機の開発過程で行った風洞試験結果, 並びに飛行実験後に行った実験機形状に対するCFD解析結果との比較を通して、将来の次世代SST研究における貴重なデータと考えられるため, 我が国において広く多くの研究者の利用に供することが重要との認識から, それらをデータベース（DB）化して2008年7月より公開した. その後, 約10年の運用を経て, 現在はサーバシステム及びネットワーク環境の変更に伴いDBシステムの改修が不可能となったため公開を断念せざる得なくなった. そこで, 利用者の利便性の観点から主要なデータ（NEXST-1実験機単体に関するデータ）については資料化を通して公開する方針に変更した. 本資料はそれらをまとめたものである., JAXA developed the database (NEXST-DB) for aerodynamic design of a future supersonic transport (SST). It is based on lots of aerodynamic results of JAXA's scaled supersonic experimental airplane project (NEXST-1) which consists of principal flight test results, wind tunnel tests data, and CFD analysis results. Then, this database was exhibited for Japanese researchers by using a web system in July, 2008. After replacing the web system in 2016, unfortunately the situation that the web system does not work has been continuing because the NEXST-DB was not able to be replaced at that time. Therefore, JAXA planned to publish the main data of the NEXST-DB as a technical report in order to solve this situation., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1930023000, レポート番号: JAXA-RM-19-010

Boundary layer transition along the leeward symmetry plane of axisymmetric bodies at nonzero angle of incidence in supersonic flow was investigated numerically as part of joint research between the Japan Aerospace Exploration Agency (JAXA) and the National Aeronautics and Space Administration (NASA). Stability of the boundary layer over five axisymmetric bodies (namely, the Sears-Haack body, the semi-Sears-Haack body, two straight cones and the flared cone) was analyzed in order to investigate the effects of axial pressure gradients, freestream Mach number and angle of incidence on the boundary layer transition. Moreover, the transition location over four bodies was detected experimentally. The strong effects of axial pressure gradients on the boundary layer profiles along the leeward ray, including an earlier transition under adverse axial pressure gradients, were indicated in both numerical predictions and experimental measurements. The destabilizing effect of the pressure gradient on the boundary layer flow within the leeward symmetry plane is shown to be related to the three-dimensional dynamics involving an increasing build-up of secondary flow along the leeward symmetry plane under an adverse axial pressure gradient. A detailed description of the mean flow computation, which forms the basis for the present linear stability analysis, is provided in an accompanying report that forms part 1 of this document., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1730005000, レポート番号: JAXA-RR-17-003E

Boundary layer transition along the leeward symmetry plane of axisymmetric bodies at zero and non-zero angles of incidence in supersonic flow was investigated numerically as part of joint research between the Japan Aerospace Exploration Agency (JAXA) and National Aeronautics and Space Administration (NASA). Mean flow over five axisymmetric bodies (namely, a Sears-Haack body, a semi-Sears-Haack body, two straight cones and a flared cone) was analyzed to investigate the effects of axial pressure gradient, freestream Mach number, and angle of incidence on boundary layer transition. Computations revealed the strong effects of axial pressure gradient on boundary layer profile in the vicinity of the leeward symmetry plane, highlighting the three-dimensional dynamics associated with increasing build-up of secondary flow under an adverse axial pressure gradient. Independent flow solutions obtained using different flow solvers and different grids at JAXA and NASA, respectively, were in good agreement with each other. Slight differences between the two sets of solutions are attributed to a combined effect of the differences between respective thermal wall boundary conditions, numerical grids, and flow solvers. The difference due to the thermal boundary condition is confirmed to be physical and was observed for all flow conditions, as expected. However, the other differences were rather minor, and were noticeable only for the straight cone and flared cone configurations. The conditions under which these minor differences are observed and the magnitudes of these differences remain an open question. Despite being coarser than the NASA grids, the JAXA grids are shown to be sufficient for providing basic state definition for the linear stability analysis. Specifically, the results demonstrate that appropriate grid spacing had been used to obtain accurate boundary layer profiles. The present report represents part 1 of a two-part document based on the joint computational effort. Part 1 is devoted to the results of mean flow computations and the results of linear stability analyses and the corresponding experiments are described in part 2., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1730004000, レポート番号: JAXA-RR-17-002E

摩擦抵抗を低減するため自然層流翼設計が小型超音速旅客機QSSTに対して適用された。設計対象は, マッハ数1.6, 50人乗りを想定したJAXAの概念設計の参照機体である。自然層流設計システムのいくつかの部分が, 高い空力性能を達成するために改善された。下面の目標圧力分布を修正した結果, 全抵抗, すなわち摩擦抵抗だけでなく圧力抵抗も改善された。, Natural laminar flow wing design was applied to the small supersonic transport QSST by using a computational fluid dynamics based supersonic inverse design method, aimed to reduce friction drag. The design target is a small supersonic passenger aircraft (Mach 1.6, 36-50 passengers), which is a reference airframe for JAXA's conceptual design. In this study, some parts of the natural laminar flow design system were improved in order to achieve high aerodynamic performance. The total drag, i.e. not only the friction drag but also the pressure drag, was improved by this modification of the target pressure distribution of the lower surface., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1630029000, レポート番号: JAXA-RR-16-001

第89回風洞研究会議 (2013年5月23日-24日), Wind Tunnel Technology Association 89th Meeting (May 23-24, 2013), JAXA2mx2m 遷音速風洞における遷音速および超音速領域のParticle Image Velocimetry (PIV) を用いた境界層計測について紹介する．PIV の要素であるトレーサ粒子，撮像系，レーザシート系，処理手法を工夫することで，遷音速領域では約10mm，超音速領域では約1mm の境界層内部の速度分布を取得した．PIV の各要素の工夫点を紹介するとともに，遷音速領域のロケットフェアリングと超音速領域のフレア型円錐の計測結果を紹介する．, 形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: AA1530024009, レポート番号: JAXA-SP-14-011

基準形状に正弦波状，あるいは局在する波状の変形を加えることで，超音速自然層流機首を設計することを試みた．変形のパラメータだけでなく，一様流に関するパラメータも変化させ，最も自然層流効果の大きくなる形状を調査した．その結果，変形によって周方向と軸方向の圧力勾配比が減少すると，側面の境界層遷移を支配する横流れ不安定を抑制し自然層流効果をもたらすことがわかった．また，この圧力勾配比を小さくするには，圧力分布のピーク位置，すなわち圧力係数が負圧側で極大となる位置をトップラインとボトムラインで適切に配置することが重要であることがわかった．そして，そのような自然層流効果をもたらす変形形状の設計指針を獲得し，自然層流機首となる形状例が得られた．正弦波変形における自然層流効果は，実験的にも検証できた．一方，全抵抗の増加を抑制するには，局在波変形の方が有効であることが示された．, The transition characteristics were examined numerically on candidate shapes in order to design a natural laminar flow nose for a supersonic transport. Two types of deformation from a baseline shape were examined. One is sinusoidal type, and another is localized type. Not only parameters of geometry which define the deformation, but also the conditions of uniform flow were varied to obtain the largest effect of natural laminar flow. In results, the crossflow instability, which dominates the boundary-layer transition at the side area, was found to be suppressed from the small ratio of pressure gradients in azimuthal and axial directions at the side area. Moreover it was found that the relative location of bump in pressure distribution along leeward and windward rays is a key to yield small ratio of pressure gradient in azimuthal direction. The effect of natural laminar flow for sinusoidal deformation was confirmed experimentally. On the other hand, the localized deformation was found to be more effective in order to suppress increasing of total drag than the sinusoidal deformation., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0062064000, レポート番号: JAXA-RR-13-008

境界層の遷移機構を正しく理解した上で遷移位置を予測する精度をより高めていくために，超音速流中に置かれた軸対称物体を対象として2003年3月から2012年12月に著者らが実施した試験結果のうち，表面静圧変動計測に関する結果をまとめる．, Experimental results of surface pressure measurements on axisymmetric bodies in supersonic flows, which were obtained from 2003 to 2012, are summarized. It aims to improve numerical technique to predict laminar-turbulent transition in boundary layer with understanding of transition mechanism., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0061881000, レポート番号: JAXA-RM-12-011

超音速流中におかれた軸対称物体上に発達する境界層流れに対する遷移機構の理解を促進するとともに，数値流体解析の結果を検証することを目的として，軸対称物体の表面静圧分布計測を宇宙航空研究開発機構が所有する1m×1m超音速風洞において実施した．その結果，従来の静圧孔を用いた表面静圧分布の計測結果は，全面層流を仮定した数値流体解析結果と計測の不確かさの範囲内でよく一致していることが確認された．表面静圧は風洞測定部内の自由流マッハ数分布の影響を受けることがわかったが，その影響は境界層遷移に対しては無視できることが確認された．, Surface pressure measurement on axisymmetric bodies in supersonic flow was performed in 1m×1m supersonic wind tunnel in Japan Aerospace Exploration Agency, for the purposes of promoting understanding of transition mechanism, and of validation of computational fluid dynamics, for boundary layers on axisymmetric bodies in supersonic flow. It is confirmed that results of surface pressure measurement by use of conventional pressure tap were in good agreement with numerical results of computational fluid dynamics within the range of the assured measurement precision. Moreover, it is found that the pressure distribution is affected by non-uniformity of freestream in the test section of the wind tunnel, but the influence is negligible small for transition prediction., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0065214000, レポート番号: JAXA-RM-11-016

次世代超音速旅客機の研究開発として進められた小型超音速実験機(NEXST)プログラムにおいて蓄積された空力データを収録した超音速機空力設計データベース(NEXST-DB)を構築した．インターネット公開型のデータベースであり，Web上で図示機能やテキストデータのダウンロード機能を充実させ，正確なデータが迅速に利用できるようにした．本NEXST-DBには空力データ，機体形状データ，空力設計ツール，そして対外発表資料を掲載・公開している．空力データには，NEXST-1の飛行実験や，CFD解析，風洞試験，境界層遷移解析結果を掲載し，機体形状データはCAD形式の形状ファイルが利用できるようにした．空力設計ツールは，NEXST-1の空力設計に用いた抵抗低減設計ツールやCFD逆問題設計ツールを掲載し，利用の多様化を図った．掲載内容や機能，利用方法の詳細についてまとめた．, A database for aerodynamic design of the supersonic transport (NEXST-DB) was constructed by JAXA. Many kind of aerodynamic data uploaded on the NEXST-DB were accumulated from the NEXST program. The uploaded data on the NEXST-DB consisted by aerodynamic results, airplane geometries, aerodynamic design tools and publications. A fl ight test, CFD analysis, wind tunnel test and the boundary layer transition analysis results obtained from the NEXST-1 project were mainly uploaded as the aerodynamic results. The aerodynamic design tools were developed on the project and used for designing of the NEXST-1 airplane. All of the aerodynamic data and tools uploaded on the NEXST-DB can freely be downloaded from the website. The detail information of the publications related to the NEXST program was also introduced., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0065139000, レポート番号: JAXA-RM-11-009

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, 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

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

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

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