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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

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

Stability analysis on a supersonic natural laminar flow (NLF) wing designed by JAXA was performed at both flight and wind tunnel test conditions as a joint research program between ONERA and JAXA. Both parties validated the NLF wing effect by confirming suppression of crossflow instability at design point, using an e(N) method with fixed β strategy proposed by ONERA. Although JAXA’s flight test vehicle and wind tunnel test model had almost same averaged roughness height of about 1micrometer, it was shown that surface roughness had little influence on measured transition location at both flight and wind tunnel test conditions. Furthermore, JAXA found that the pressure distribution measured in flight test did not completely coincide to the target pressure distribution (Cp (Target)) for the NLF wing design. It revealed that measured transition location at outer wing region was not delayed in flight test, compared with the transition location predicted with the Cp (Target) which had large gain to delay the transition. Influence of Reynolds number on transition location was also studied for the target pressure distribution Cp (Target). Consequently, chordwise transition location, (x/c)Tr., rapidly moves from mid-chord location to forward location (near leading edge) for increasing Reynolds numbers, according to the change of instability of boundary layer from Tollmien-Schlichting instability to crossflow instability., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0062081000, レポート番号: JAXA-RR-12-009E

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 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, The aerodynamic design of an SST (SuperSonic Transport) configuration with some geometry requests at given flight condition in 3rd NAL-SST-CFD (Computational Fluid Dynamics) workshop competition subject 1 is conducted by following three design tools: a conventional linear theory design tool, a CFD-based optimization tool and a CFD-based inverse design tool. In order to increase a Lift to Drag ratio (L/D) of the airplane, several design concepts are adopted such as a warp wing design, linear and non-linear area-ruled fuselage designs and a supersonic NLF (Natural Laminar Flow) wing design. The non-linear area-ruled fuselage design shows that the CFD-based optimization design tool can improve the aerodynamic properties of the airplane designed by the conventional linear theory design tool. As for the NLF wing design, however, the CFD-based inverse design tool should be carefully coordinated in order to realize the target NLF pressure distributions., 資料番号: AA0063609055, 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, 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 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

Nonlinear stability of three-dimensional supersonic boundary layers on JAXA's experimental airplane called NEXST-1 is investigated for the case in which first-mode instability is dominant. Several nonlinear mode interactions are numerically examined by solving nonlinear parabolized stability equations. The existence of the standard initial amplitude is assumed, and its value is estimated from the flight data of the NEXST-1. It is shown that disturbances with a critical parameter (ω(sub 0), β(sub 0)) predicted by linear parallel stability theory are linearly amplified until transition sets in. Then the initial amplitude correlates with the measured transition location is found to be approximately 10(exp -6) regardless of nonlinear interactions presented., 形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: DS1640260002

We applied a method based on Complex-ray theory (CRT) to the stability analysis of three-dimensional boundary layers at supersonic flight test conditions of the JAXA's experimental airplane called, NEXST-1, and investigated its capability. The dispersion relation was numerically obtained by the eigenvalue equations and linear parabolized stability equations (PSE). By adapting two kinematic wave theoretical models, we found out which wave type was dominant at measured transition locations of the flight test in 2005. As a result, it has been found that the dominant wave type on the outer wing section of the NEXST-1 was wedge-shaped disturbance, on the other hand, it was two dimensional wave-packets on the inner wing section. From the view point of transition prediction using an e to the Nth power method, there was no difference of transition N factors (N(sub tr)) between the results by the eigenvalue solutions with and without CRT method. Only a use of PSE with CRT, however, was able to reduce the discrepancy of N(sub tr). among cross sections, owing to the nonparallel effect of flows., 形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: DS1540529011

小型超音速実験機（ロケット実験機）飛行実験データ解析完了報告会(平成20年7月10日. 東京大学山上会館), Transition location of an unmanned and scaled supersonic experimental airplane (NEXST-1) is measured by the flight test. The wing of NEXST-1 is designed with our original CFD-based inverse design method to apply the natural laminar flow concept. It is the first challenge to apply the natural laminar flow wing concept to a supersonic vehicle. The concept is tried to validate by measuring the surface pressure and the transition location. In this paper, the flight data analysis to detect transition location by four kinds of sensors and the results are summarized. Also the development of numerical transition prediction and its results are summarized. The transition location at the design condition, which is detected experimentally, delays in contrast with that at the off-design point. Therefore the natural laminar flow effect is confirmed., 形態: カラー図版あり, 資料番号: AA0064239007, レポート番号: JAXA-SP-08-008

小型超音速実験機（ロケット実験機）飛行実験データ解析完了報告会(平成20年7月10日. 東京大学山上会館), An aerodynamic database system named the NEXST-DB was constructed by JAXA to contribute on the aerospace research area in Japan. A number of aerodynamic data on the supersonic experimental airplane (NEXST-1) can be accumulated by the NEXST project. Three kinds of data of the pressure distribution, aerodynamic forces and boundary layer transitions were carried on the NEXST-DB. These aerodynamic data were obtained from the flight test, CFD analysis and wind tunnel tests. Furthermore, the geometry data of the NEXST-1 and the grid data for CFD analysis were also carried on the NEXST-DB. Those data can download on the website., 形態: カラー図版あり, 資料番号: AA0064239008, レポート番号: JAXA-SP-08-008

Linear stability analysis of compressible three-dimensional boundary layers on a supersonic natural laminar flow wing, which has strong crossflow and wall curvature effect extremely near the leading edge, is conducted at the flight test condition by solving parabolized stability equations (PSE) as well as the eigenvalue problem in linear stability theory (LST). The effects of the non-parallelism and the wall curvature which the LST approach excludes, and the normalization condition of the PSE method on the stability are investigated, focusing on both stationary and traveling modes subject to crossflow instability and the first unstable mode. Subsequently, the so-called e to the nth power method is used as a transition prediction method coupled with the fixed β technique. Finally each N factor map on the wing based on the PSE approach is compared with the result by LST. It is found that stationary modes are very sensitive to the wall curvature compared with traveling modes, and that some shortcomings related to the pathological problem, which existed on the results by the LST approach, can be solved by the current PSE method., 形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: DS1540301001

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

遷移計測の高度化高効率化を念頭に、最適な計測手法選定と関連技術開発を目的として、開発試験において実施することが予想される非断熱性の輸送機形態模型を対象に遷移計測を実施した。試験は宇宙航空研究開発機構風洞技術開発センター6.5m×5.5m低速風洞(LWT1)において、ONERA低速標準模型を対象に行った。計測手法としては、定量的点計測法としてプレストン管およびホットフィルムセンサー、また、可視化的手法として感温液晶フィルムおよび赤外線カメラを使用した。試験の結果、プレストン管およびホットフィルムによる定量計測は、計測点の移動に時間を要するものの詳細かつ正確なデータ取得に適していることが実証された。また感温液晶と赤外線カメラを用いた計測およびその結果から、非断熱性の模型上であっても模型の強制冷却を行うことにより模型表面温度分布からの遷移点判定が可能であることが実証された。総合的には、本試験のような条件下においては、赤外線カメラ計測結果から得たマクロな情報を元に、プレストン管あるいはホットフィルムにより重要なポイントの詳細計測を行う手法が最適であるという結論を得た。, This study developed procedures to measure the boundary layer transition on a metallic aircraft model in a typical large-scale low-speed wind tunnel. At the JAXA 6.5 m x 5.5 m low-speed wind tunnel (LWT1), the boundary layer on the left wing of the ONERA model was investigated using two quantitative methods (Preston Tube and Hotfilm Sensor) and also two visualizing methods (temperature-sensitive Liquid Crystal, Infrared Camera). In the course of the experiments a technique for investigating transition points was developed, which is presented here., 資料番号: AA0047949000, レポート番号: JAXA-RR-04-018

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

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