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The coking phenomenon within low- density ablative material exposed to aerodynamic heating is studied. The char density of the tested material is measured by slicing the material thinly, and by weighing each of the sliced test pieces. Measured density profiles of the tested materials are compared with the calculated results obtained by a thermal response code. The surface and internal state of both pre- and post- materials are observed by using Scanning Electron Microscope (SEM) and X-ray Computed Tomography (CT). The results show that the measured density profiles is inversely increased to the heated surface. The density increase is likely due to the deposition of a solid carbon in the char region during aerodynamic heating., 資料番号: AA0063742010, レポート番号: JAXA-SP-07-016

資料番号: AA0064112006, レポート番号: JAXA-SP-07-023

M-Vロケットは、それ以前のM-3SIIロケットと形状が大きく異なるため、ほとんど全ての搭載アンテナについて新規開発を行った。その後、打上げ回数を重ねることで明るみになった問題点を、飛翔結果をもとに改修し、M-Vロケット搭載用アンテナとして最適な設計に至った。本報告では、開発経緯とその性能について述べる。, 資料番号: AA0064112009, レポート番号: JAXA-SP-07-023

SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design 航空宇宙技術研究所 19980316-19980317 東京 日本, SST-CFD Workshop. International CFD Workshops for Super-sonic Transport Design National Aerospace Laboratory 19980316-19980317 Tokyo Japan, A brief review of the current capabilities and outstanding issues for the efficient analysis, and design of high-speed flight vehicles using unstructured grids is given. It is concluded that all major areas required in the analysis cycle - grid generation, flow solvers and visualization - have seen major advances in recent years. This trend is expected to continue, leading us to believe that turnaround in a matter of hours or minutes for complex geometries and/or physics will become a reality sometime in the next decade., 資料番号: AA0063609021, 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, NAL(航空宇宙技術研究所)においては、NASDA(宇宙開発事業団)との共同研究として、大気圏宇宙輸送機の空気力学研究のため数値シミュレーションを行ってきた。計算機ハードウェアシステムの進歩に伴ってCFD(計算流体力学)技術はその応用範囲を広げた。そして今やCFDは強力な空力設計ツールとして発展し、打上げから再突入フェーズまでの宇宙輸送機の飛行範囲を取り扱えるようになった。ここではNALにおける数値シミュレーションの最近の進歩を紹介した。, Numerical simulation for aerodynamic study of atmospheric space transport vehicles has been conducted at NAL (National Aerospace Laboratory) as the cooperative research work with NASDA (National Space Development Agency). CFD (Computational Fluid Dynamics) technology enlarges its application area with the progress of the computer hardware systems and now CFD has been developed as the strong aerodynamic design tool which covers flight range of space transport vehicle from the launching to the re-entry phase. Here, recent progress of numerical simulations at NAL are introduced., 資料番号: AA0028638024, レポート番号: NAL SP-49T

In recent years, development of re-usable space transportation system is promoted. However, in order to realize it, development of the base technology is indispensable, particularly, aerodynamic heating and characteristics with strong real gas phenomena should be comprehended more precisely. Under such circumstances, acquisition of detailed experimental data like drag and heat transfer over hypersonic bodies is indispensable for the development of reliable space vehicles and thermal protection system (TPS), and also for the validation and progress of numerical estimation. In the present study, based on the numerical results a blunted double cone model (heat model) was designed and manufactured as the aerodynamic configuration design of re-entry capsules. The aim of this experiment is to obtain detailed data of heat flux and pressure distribution as well as the flow structures over the heat model. Coaxial thermocouples and pressure sensors are mounted at the frontal and base region thoroughly, and surface temperature and wall pressure are measured., 資料番号: AA0049500019, レポート番号: JAXA-SP-06-006

In the present paper, two blunt cone models with 15 degrees of half angle were tested in a high enthalpy shock tunnel to investigate fundamental real gas effect on high enthalpy aerodynamic characteristics. The curvature radius of the tip of model are R = 20 mm and R = 50 mm. The tests were carried out at 4 MJ/kg-14 MJ/kg of stagnation enthalpy and the axial forces were measured by the force balance with acceleration compensation. The present results were compared with the results obtained by calculation of CFD (Computerized Fluid Dynamics). From the comparison between the experimental results and the results of CFD the real gas effect on axial force was discussed., 資料番号: AA0049500021, レポート番号: JAXA-SP-06-006

An integrated numerical method to obtain thermal response of ablator under aerodynamic heating environment is briefly described. In this method, a two-dimensional thermal response analysis code, namely SCMA2, is loosely coupled with a thermochemical nonequilibrium CFD code. The method is applied to reproduce the thermal response of ablator obtained by the heating tests conducted in the 1 MW arc-jet wind tunnel facility at ISAS/JAXA. Surface temperature, heat flux variation and surface shape change of the ablative test piece under arc-jet flow condition are calculated by the integrated code. These obtained results are compared with those given by one-dimensional SCMA code, and also with the experimental data for the purpose of validation. As to the comparison of calculated surface temperature with experimental data, quantitative agreement is yet to be accomplished because of the uncertainties in freestream conditions and material surface conditions. However, obtained results show that the present coupled approach is still available to predict the thickness of ablator required for space vehicles in future missions., 資料番号: AA0049212014, レポート番号: JAXA-SP-05-017

In the present paper, results of force and heat flux measurement tests conducted in the JAXA 1.27 m Hypersonic Wind Tunnel (HWT) using the HB-2 standard model are presented. The tests were conducted at a nominal Mach number of 10, a stagnation enthalpy of 1 MJ/kg, and stagnation pressures from 1 to 6 MPa. Both for the force and heat tests, good data repeatability were confirmed. The heat transfer coefficient obtained from the IR thermography agreed well with the one from the thermocouple measurement. On the other hand, a conventional uncertainty estimation method which estimates the bias and random error components was applied to the force test data, whereas Monte Carlo approach was utilized to numerically estimate the uncertainty in the data reduction process of the heat flux data. The present experiment was conducted as a series of comparison test campaign between two hypersonic facilities in JAXA. The present test data is also expected to serve as a set of database for the validation of hypersonic computational fluid dynamics (CFD) codes., 資料番号: AA0049211008, レポート番号: JAXA-SP-05-016

柔構造エアロシェルは再突入体や惑星突入する衛星への応用が提案されているが、柔構造体の高速気流中での空力特性に関しては、完全に理解されているとは言えない。そこで、その高速飛行性能を実証するためにJAXAの大気球を用いた柔構造機体の飛行試験が計画され、その第1回目の試験が2003年9月に行われた。本論文は、カプセル形状の本体にフレア型の柔構造エアロシェルを取り付けた機体のフライト試験の概要、機体開発、および、実験結果についての報告を行う。フライト試験では、気球のゴンドラとフライト機体の切り離し機構の不具合によって機体単体での飛行は実現しなかったが、1)風洞実験などの事前試験によって、フレア型の柔構造エアロシェルに関して、その変形形状や空力特性などについて知見を得た、2)直径1.35mサイズの柔構造エアロシェルの収納展開を地上試験で実証した、3)膜展開構造を含むフライト機体とカプセル内部に搭載された機器は飛行中健全に機能していたことが確認された、などの今後の開発につながる様々な成果が得られた。, Membrane aeroshell has a potential for application to re-entry capsules and planetary-entry probes in the future due to its low mass and low storage volume. Its characteristics, however, is not well understood particularly in high speed flow regions. The flight test using the scientific balloon is conducted to demonstrate its high speed flight capability on September 1st, 2003. This paper describes the outline of this project, the development report of the flight model with the flare-type membrane aeroshell, and the results of the flight tests. Despite the flight model dropped with a gondola because of the trouble in the separation device, there are a lot of accomplishments in this project; 1) to obtain the knowledge about the deformation, stability, and aerodynamic characteristics of the flare-type membrane aeroshell by pre-flight wind tunnel test, 2) to demonstrate that 1.35 meter diameter membrane aeroshell can be stowed in a backside of the 0.5 meter capsule and deployed automatically in the ground test, and 3) to verify that the almost all of the sensors installed in the capsule worked normally at the high altitude and the deployed membrane aeroshell can withstand the actual flight environment., 資料番号: AA0048133004, レポート番号: JAXA-RR-04-015

In this research, optimal design, evaluation, and aerodynamic problems for future space transportation systems are conducted. A conceptual design tool for space transportation systems called SEAT (Systems Evaluation and Analysis Tool) is under developing. The design criteria are also investigated and several types of the concept for the future space transportation systems are designed optimally and are evaluated. In the last fiscal year, a prototype of the SEAT was developed. The catalysis is dealt as aerodynamic problems for the future space transportation systems. In the last fiscal year, experimental data of various flow conditions were obtained in new ICP-heated wind tunnel., 資料番号: AA0049056006, レポート番号: JAXA-SP-05-010

REV was the first reentry capsule developed through pure domestic technology and successfully recovered. REV is thermally protected against the severe aerodynamic heating by carbon-phenolic ablator material. The present paper describes the feature of the heating test in the arc wind tunnel and the ablator characteristics clarified during the research and development process of the REV heat shield. Because the heat flux anticipated on the stagnation surface is about 2 MW/sq m, the dominant surface recession mechanism is identified to be in the reaction control region. The thermochemical reaction data have been acquired through three arc-heater facilities with different enthalpy level for calibration and tuning of the ablation analysis code. The thermomechanical behavior of the ablator under the high heat flux environment such as delamination, or spallation also has been investigated for functional safety of the heat shield., 資料番号: AA0048710005, レポート番号: JAXA-RR-04-045

The Unmanned Space Experiment Recovery System (USERS) was launched on September 10th, 2002 from Tanegashima Space Center with the H-2A Launch vehicle. On May 30th 2003, the re-entry and recovery operations were successfully conducted, bringing experiment samples and beneficial information back to Earth. USERS became the first national project to succeed in recovery from orbit. In this paper, USERS mission outline is introduced first. Followed by are discussion on the key characteristics of Reentry Module (REM), result of design and development, evaluation of re-entry and recovery operation and flight data. Finally, the obtained reentry system technologies through this project are summarized., 資料番号: AA0048710001, レポート番号: JAXA-RR-04-045

A computer code for charring ablation and thermal response analysis has been developed for simulation of one-dimensional transient thermal behavior of multi-layer stack of isotropic charring ablation materials used for re-entry capsules. A mathematical model of charring ablation including governing equations is summarized. Analytical solutions of several thermal response problems with or without ablation agreed very well with numerical results that are obtained with the charring ablation analysis code. Arc-heated tests of CFRP ablator models were carried out and measured results including temperature response data were compared with the simulation results of the ablation analysis code. The test results and the simulation results agreed very well. Measurements of the successfully recovered USERS/REV ablator were carried out. The measured results of the in-depth temperature history in the ablator, the char depth, the surface recession, etc. agreed well with the simulation results using the ablation analysis code. These results show that the ablation analysis code based on the present mathematical model are available for the thermal response prediction of CFRP ablator used in the re-entry capsules such as USERS/REV., 資料番号: AA0048710004, レポート番号: JAXA-RR-04-045

Normally, reentry bodies, even if they are to be recovered, are not equipped with onboard instruments such as transponders, which are useful for real-time position monitoring of the reentry bodies during the flight. This is because the surrounding plasma generated by the aerodynamic heating makes the transponder positioning nonsense as the blackout phenomena. If the trajectory during the reentry phase could be known, even if it's not monitored in real-time, the information will be a great help for the rapid recovery. The radiation from the highly-heated reentry body during the aerodynamic heating has enough intense radiation in visible region for the detection on the ground. This paper investigates a method of optical capturing of the reentry bodies by means of the visible radiation from the bodies during the aerodynamic heating. And a result of the optical tracking actually carried out on the occasion of the reentry of the REV capsule is also presented., 資料番号: AA0048710008, レポート番号: JAXA-RR-04-045

The re-entry capsule USERS/REV was successfully recovered in May 2003. REV has a heat shield system to protect inner equipments against the severe heating environment during the re-entry. The heat shield system is mainly made of three kinds of layers, i.e., CFRP ablator, felt type insulation material with thermal anchor and an aluminum shell structure. In the design of the heat shield system, contradictory requirements of heat ejection in orbit and heat protection during the re-entry had to be satisfied. Thermal protection performance of the heat shield system was calculated using an ablation analysis code. Thermal stress analysis of the ablator during the reentry was carried out using a large deformation FEM that considers materials non-linearity. For the development of the heat shield system, various kinds of tests such as the arc-heated tests of the ablator and the strain survey test of the heat shield system were carried out. After the recovery of REV, evaluation of the recovered heat shield system was carried out. The temperature time histories in the ablator obtained during the re-entry and the damage of the heat shield system measured after the recovery were compared with the results of ablation analysis and structural analysis, from which the heating environment during the re-entry and the cause of the damage were estimated. The evaluation of the heat shield system design based on the flight results is also described., 資料番号: AA0048710006, レポート番号: JAXA-RR-04-045

In the READ experiment, the in-fight measurement of the radiation, emanating from the shock layer generated in front of the USERS reentry module during the reentry flight, was successfully performed. The radiation generated at the shock layer was introduced into the onboard spectrometer in the USERS reentry module through the optical fiber. During the flight experiment, the spectrum of the radiation was recorded on the on-board data recorder and was successfully recovered after retrieval of the vehicle. The measured data shows that the spectrum was contaminated by the radiation other than the radiation emanating from the shock layer, and its intensity was stronger than was expected. The spectrum at the altitude of around 90 km shows a strong influence of the ablation gas while the spectrum at the lower altitude shows an influence of the albedo of Earth surface., 資料番号: AA0048710009, レポート番号: JAXA-RR-04-045

HOPE研究が盛んであった1990年代初期、大気圏再突入時の流れを地上で再現して実在気体熱空力特性を観測することを目的として大型高温衝撃風洞の建設計画が進められ、スクラムジェットエンジン試験への利用にも積極的であったJAXA宇宙推進技術共同センターにおいて1997年度末にHIESTが完成した。以来、HOPEその他の熱空力試験および超高速域スクラムジェットエンジン試験など現在まで1,000回を超える試験を行ってきた。ここではこれまでHIESTにおいて行ってきた試験の主要な成果について述べる。, 資料番号: AA0048471001, レポート番号: JAXA-SP-04-014

A series of wind tunnel tests were conducted to develop a boundary-layer tripping technique and also a detection technique for the boundary-layer state in transonic wind tunnels in ISTA/JAXA. A commonly used disk-type roughness element was used in the test and the conventional criterion for the effective roughness height was confirmed valid in the present case, also. A detection methodology for the boundary-layer state based on the quantitative measurement of the aerodynamic heating was applied to this experiment in a blow-down transonic facility, using both the Temperature Sensitive Paint (TSP) technique and the Infra-Red Thermograpy (IR) technique. The results indicate the methods are both useful in transonic experiments. Transition detection in a closed-circuit transonic wind tunnel was also achieved by varying the reservoir temperature, but it is found that the technique is not necessarily good for the boundary-layer state determination without any additional roughness for the reference purpose., 資料番号: AA0048470003, レポート番号: JAXA-SP-04-013

宇宙航空研究開発機構総合技術研究本部(旧称:独立行政法人航空宇宙技術研究所)において、2次元軸対称有限要素アブレーション解析コードFISCAMが開発された。本コードは、航空宇宙技術研究所において開発された熱伝導解析FEMプログラムTHAPを元に、通常の熱伝導に加えてアブレーション挙動を再現できるよう改造を施したものである。アブレーション材料の挙動は大きく3つ(炭化領域／熱分解領域／母材領域)に分かれている。3領域とも通常の熱伝導方程式が適用されるが、熱分解領域では母材が熱分解により炭化を起こし、それに伴い熱分解気体が発生し、その結果潜熱が奪われ、発生した気体は炭化領域から外部に向けて流出する。炭化領域についてはこの分解ガスの質量流量は一定と仮定する。また、表面においては外部からの空力加熱、酸化による発熱、放射による熱流出、昇華による熱吸収が考慮され、空力加熱については表面からの熱分解ガスの流出による熱流入遮断効果を考慮している。このFISCAM開発の目的は、再突入物体やロケットノズルに代表される2次元軸対称物体が強力な熱荷重を受けた際に材料が熱分解を起こし、炭化しつつ熱分解ガスを放出、同時に表面炭化層が酸化により失われ後退する現象であるアブレーション挙動を解析することであった。従来より開発、使用していたコードは1次元であり、FISCAMのような軸対称物体に適用できるコードにより、肉厚が大きくかつ曲率の大きい物体の挙動に対し、1次元解析よりも精度よく再現する事が可能になる。加えてFISCAMは基礎方程式の離散化に有限要素法を用いている事も特色の1つであり、FISCAMと同様の目的で開発されたと考えられるNASAエイムズ研究所で開発されたコードなどに対する差別化も図っている。, The Finite-Element based Identification System of Charring Ablative Materials (FISCAM) Program is presented. The program was developed for the analysis and identification of heat conduction problem which includes thermal decomposition of material, pyrolysis gas generation, and surface recession due to combustion of the charring material. The governing equation of two-dimensional axisymmetric heat transfer problem was discretized using a finite-element mathematical model. An identification problem of heated cylinder is presented in this paper., 資料番号: AA0048469009, レポート番号: JAXA-SP-04-012

A trajectory-based heating analysis of the Galileo probe entry flowfield is attempted to reproduce the heatshield recession data obtained during the entry flight. In the present calculation, the mass conservation equations for the freestream gas (hydrogen-helium gas mixture) and the ablation product gas are solved with an assumption of thermochemical equilibrium. The ablation process is assumed to be quasi-steady and is coupled with the flowfield calculation. The radiative energy transfer calculation is tightly coupled with the flowfield calculation, where the absorption coefficients of the gas mixture are given by the multiband radiation model having 4,781 wavelength points for wavelength range from 750 to 15,000 A. The injection-induced turbulence model proposed by Park is employed to account for the enhanced turbulence effect due to the ablation product gas. It is shown that the final recession profile of the flight data at the frustum region can be closely reproduced if the injection-induced turbulence model is employed, although that at the stagnation region is overestimated. The cause of the enhanced radiative heating that occurs at the frustum region is given in connection with the enhanced turbulence effect in the shock layer., 資料番号: AA0048469010, レポート番号: JAXA-SP-04-012

資料番号: AA0048140004, レポート番号: JAXA-SP-04-011

An experimental investigation for an efficient tripping method of the hypersonic boundary layer is conducted with a 5 deg half angle sharp cone in ISTA/JAXA phi0.5 m hypersonic wind tunnel. Aerodynamic heating distributions and fluctuations of wall pressure are measured both with and without roughness elements. A set of transition data and wall pressure fluctuation measurements in the wind tunnel are obtained. A two-dimensional wavy roughness of a wave length approximately as two times much as the estimated boundary layer thickness doesn't show any effects on the location of the boundary layer transition, while in the case where the boundary layer thickness is approximately 1.25 times of the half wavelength, a clear delay in transition is observed. Effects of conventional three dimensional roughness elements on tripping effectiveness are also examined., 資料番号: AA0047776012, レポート番号: JAXA-SP-04-002

宇宙航空研究開発機構では、将来の宇宙輸送システムコンセプトの絞り込み、開発すべき技術課題の抽出、および現状技術の改善目標を設定するため、概念設計ツール(Systems Evaluation and Analysis Tool: SEAT)を開発中である。本報告書では、まずSEAT開発の目的、開発シナリオをまとめ、開発した雛形ツールについて報告する。ついで、代表的なエンジンの評価を行うため、雛形ルールを用いて5種類の宇宙輸送システムの概念設計を行った。本報告書では、液体ロケットエンジン、Turbine-Based Combined Cycle(TBCC)エンジン、およびRocket-Based Combined Cycle(RBCC)エンジンを対象とした。概念設計は、全備重量が最小となるように行った。その結果、母機、軌道機共にロケットエンジンを使用する2段式宇宙往還機のコンセプトが、最も軽量となる結果が得られた、本システムはエンジン搭載性についても問題がなく、最も現実的なシステムである。TBCCエンジンやRBCCエンジンといった空気吸い込み式のエンジンを使用する場合には、必要なエンジン基数が多くなり、その搭載性に問題があるという結果が得られた。, The Japan Space Exploration Agency (JAXA) is developing a concept design tool called SEAT (Systems Evaluation and Analysis Tool) to assist in evaluating candidate space transportation system concepts and selecting promising concepts for further development. This paper describes the SEAT development objectives, development strategy, and present status. The concept design of five systems that use the rocket, the Turbine-Based Combined Cycle (TBCC), and the Rocket-Based Combined Cycle (RBCC) engine are also performed using the SEAT. The concept design is performed to minimize the gross take-off weight. The lightest system concept is the Two-Stage-To-Orbit (TSTO) using rocket propulsion for both stages in this design., 資料番号: AA0063606000, レポート番号: JAXA-RR-07-006

JAXA1.27m極超音速風洞では赤外線サーモグラフィーを用いて空力加熱率分布の測定を行っている。AGEMA社製LW900型の赤外線カメラを3台使用し、模型全体の加熱率分布を同時に計測できる技術を確立した。赤外線サーモグラフィー法の測定精度を確認するため、半球標準模型試験および平板ランプ模型試験を行った。半球標準模型試験は、半径0.05mの半球模型を用い、よどみ点の加熱率をSagnierらの推定値と比較し±5％程度の誤差で測定できた。模型の大きさによる測定誤差を調べるため、半径0.04m、0.03m、0.02m、0.01m、0.005mの半球模型と半径0.0075m、0.005m、0.0025mの鈍頭平板模型を用い、曲率半径の影響を調べた。半球模型で0.01m、鈍頭平板模型で0.005mまでは上記推定値と比較して±10％程度の誤差で測定できた。平板ランプ模型試験では同軸熱電対法と赤外線サーモグラフィー法を同形状の模型で実施し、両測定法の相違を測定した。模型先端部の曲率の変化によりランプ部の加熱率が大きく変化するが、両者による測定結果はよく一致していることを確認した。赤外線サーモグラフィー法の測定誤差要因を項目ごとに整理し誤差を見積もった。その結果、曲率半径の小さな部位や角は除き、総合的な測定精度として±10％以内の誤差で測定できることがわかった。, In the 1.27 m hypersonic wind tunnel, we have developed an Infrared (IR) thermography technology to measure heat transfer distribution on a model using three IR cameras of AGEMA LW900. In order to verify the technology we conducted two wind tunnel tests: a hemisphere standard model test and a flat plate ramp model test. In the hemisphere standard model test, we confirmed that the heat transfer at a stagnation on a hemisphere of 0.05 m in radius can be measured within an error of 5 percent by comparing the result with the Sagnier estimate. We investigated the effect of model size on measurement accuracy by testing several sizes of hemisphere and blunt-edge models, and confirmed that the models of hemisphere radius of more than 0.01 m and the blunt-edge radius of more than 0.005 m can be measured within an error of 10 percent. In the flat plate ramp model test, we compared the IR thermography method with a method of using co-axial thermocouples. The results of the two methods agreed very well. We analyzed error sources that affect the IR thermography and confirmed that the accuracy within +/- 10 percent is retained except regions of sharp corners., 資料番号: AA0063460000, レポート番号: JAXA-RR-06-026

本報告は、JAXA高エンタルピ衝撃風洞(HIEST)において実施された、HB-2形状バリスティック標準模型を用いた空気力/空力加熱率測定試験結果に関するものである。力試験はアルミ製模型に十字1列型3分力天秤を装着して実施され、データの振動成分を除去するために加速度補正が施された。熱試験で用いられた模型はクロメル製で計28個の同軸熱電対が装着されてあり、温度時間履歴データを1次元熱伝導の仮定の下で数値的に積分することにより加熱率を求めた。力/熱試験共にコニカルノズルを使用し、力試験についてはエンタルピ4、8、11MJ/kg、熱試験については4、8MJ/kgの試験条件でそれぞれ実施された。力試験結果はエンタルピ4MJ/kg条件の場合、軸力係数に関して他風洞の試験データと良好な傾向の一致を示している一方で、8、11MJ/kgの場合は系統的なずれが観測されており、実在気体効果の影響が可能性として示唆される。一方熱試験における無次元加熱率データは、ブローダウン式コールド極超音速風洞データと模型頭部において数％程度の範囲でよい一致を示した。本試験は、JAXA極超音速関連風洞対応風試の一環として実施されたものであり、また本試験データは高エンタルピ衝撃風洞における基礎データとしても有用である。, This report presents a detailed description and results of force and heat flux measurements conducted at the High-Enthalpy Shock Tunnel (HIEST) of the Japan Aerospace Exploration Agency (JAXA). The HB-2 standard hypersonic ballistic configuration was employed as a model. The force measurement tests used an aluminum alloy model and a three-component aerodynamic balance. An acceleration compensation technique was used to remove vibration components from the original data. Heat flux measurement tests used a chromel model with a total of 28 co-axial thermocouples press-fitted onto the surface. Time histories of temperature data were numerically integrated to determine the surface heat flux by applying one-dimensional heat conduction theory. Both force and heat tests used a conical nozzle, and a total of three enthalpy levels of 4, 8, and 11 MJ/kg and two levels of 4 and 8 MJ/kg were set for the force and heat tests, respectively. Although much higher stagnation enthalpy can be attained, relatively lower enthalpy levels were selected to compare with a blow-down type hypersonic wind tunnel and to alleviate uncertainties inherent in high-enthalpy facilities. The force test results correlated well with other hypersonic wind-tunnel data for the axial force coefficient at the enthalpy level of 4 MJ/kg, while a notable discrepancy was observed at the higher enthalpy levels of 8 and 11 MJ/kg. The real gas effects on aerodynamic characteristics are examined and discussed. In the heat test, a comparison of the heat flux distribution along the surface with data obtained in a blow-down type hypersonic wind tunnel showed good agreement in the nose part, typically within several percent. The present experiment was conducted as a series of comparative test campaigns between two hypersonic facilities in JAXA. Since the tunnel stagnation conditions and the corresponding data are tabulated in detail for each shot, the present test data are believed to serve as a complete database for evaluating aerodynamic force and heat flux data obtained at large-scale high-enthalpy shock tunnels., 資料番号: AA0049482000, レポート番号: JAXA-RR-05-030E

20041125-20041126, 資料番号: AA0048471000, レポート番号: JAXA-SP-04-014

航空宇宙数値シミュレーション技術シンポジウム2004 宇宙航空研究開発機構 20040624-20040625 東京 日本, Aerospace Numerical Simulation Symposium 2004 Japan Aerospace Exploration Agency 20040624-20040625 Tokyo Japan, 資料番号: AA0048469000, レポート番号: JAXA-SP-04-012

宇宙航空研究開発機構の1.27m極超音速風洞は、ノズル出口直径1.27m、固定ノズルによるマッハ数10の世界最大規模の極超音速風洞である。1995年の完成以来、6分力、圧力分布、空力加熱分布測定などの試験に利用されている。1999年には機械的再加工によるノズルの改修が完了し、気流品質は格段と高いものになっている。本風洞の気流特性確認試験として、これまでピトー／総温度レークを用いてピトー圧力分布、総温度分布を、AGARD-J(円錐形状)模型を用いて気流傾角を測定してきた。これらの気流特性確認試験の一環として、標準模型のひとつであるAGARD-E(半球形状)模型を用いて詳細な表面圧力分布測定試験を行った。試験はノズル改修前後2回に分けて実施し、改修前後の比較を行った。その結果、ノズル改修後の気流の質は、改修前に比較して大きく改善されたことを確認できた。表面圧力分布の実験値とナビエ・ストークス方程式によるCFD結果との比較を行って良く一致することが判った。シュリーレン装置により実験的に得られた離脱衝撃波の解析から気流の比熱比を求め、不完全気体の式から得られる値とほぼ一致することを確認した。, The 1.27 m hypersonic wind tunnel at JAXA (Japan Aerospace Exploration Agency), with its fixed Mach 10 nozzle, is one of the largest hypersonic wind tunnels in the world. It has been contributing to hypersonic vehicle design by supplying data on aerodynamic force and moment, pressure distribution, and aerodynamic heating. The tunnel was completed in 1995, and the nozzle was re-machined in 1998 to improve flow quality. To verify the flow characteristics, calibration tests using pitot rakes and an AGARD-J model have been performed. A series of calibration tests was performed to measure the surface pressure distributions on the AGARD-E (hemisphere) model before and after nozzle modification. This report first compares the experimental data before and after nozzle modification and evaluates the effects of the modification. Then, it compares the experimental data with the CFD results and with other AGARD-E results conducted in other wind tunnels. Finally, estimated is the ratio of specific heats gamma using schlieren photographs of the detached shock. The results indicate a significant improvement in the quality of the flow. The experimental results using the modified nozzle also show good agreement with the results of CFD and the other wind tunnels. The estimated ratio of specific heats gamma agrees well with the theoretical value based on an imperfect gas., 資料番号: AA0048135000, レポート番号: JAXA-RR-04-019

The Institute of Space Technology and Aeronautics has investigated various concepts of next generation reusable space transportation systems. Lifting-body re-entry vehicle is one of the most promising concepts among them. This type of vehicle has no wings and derives its lift solely from the shape of its body, and has the advantages of reduced structural weight, superior volumetric efficiency, and better characteristics against aerodynamic heating at hypersonic speed. On the other hand, the disadvantages of its low lift-drag ratio, poor stability and controllability at low speed makes it more difficult to land on a conventional runway than in the case of a winged vehicle like in the Automatic Landing Flight Experiment (ALFLEX), which conducted in 1996. In this report, we introduce the outline of the intended Lifting-body Flight Experiment (LIFLEX), which aims to gain the automatic landing technology using a small, low-cost system., 資料番号: AA0049500025, レポート番号: JAXA-SP-06-006

Experimental investigation of the boundary-layer transition on a 10-deg sharp cone at 2-deg incidence was conducted at Mach 2. Transition locations were obtained with small surface roughness on the order of 0.1 micron using an infrared camera. The result showed that the most forward transition occurred on the side rather than on the leeward ray as observed in the previous studies. The comparison of the transition front was made between two supersonic tunnels with similar turbulence levels., 資料番号: AA0047776011, レポート番号: JAXA-SP-04-002