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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

エアロキャプチャは、従来の化学推進による手段と比較して、軌道投入に要する推進剤を減らすことが可能であると知られている。 突入飛行経路角のウィンドウや空力加熱環境などの厳しい条件から、これまで実現されたことはなかった。しかし、軽量かつ大面積の膜面を利用した低弾道エアロキャプチャならば、そのようなリスクを避け、実現可能性がある。JAXA将来ミッションとして計画されているソーラセイルミッションでは、母船から分離される小型オービタの木星周回軌道への投入も計画されている。そこで、ソーラセイルに使用された母船の膜面の1部を小型オービタの軌道投入に再利用するとした仮定のもと、エアロキャプチャの木星における実現可能性について検討を行った。, The aerocapture is known to be an orbit insertion technique around planetary body at a fraction of the propellant of the traditional chemical propulsive technique. Severe aerocapture conditions such as the very narrow entry corridor and the aerodynamic heating environment have prevented us from putting it into practical use in a planetary exploration mission. However, the low-ballistic-coefficient aerocapture where an orbiter is connected to the light and wide membrane decelerator has the potential to reduce the risk of the aerocapture. In the Solar Sail mission which is planed as one of the future missions of JAXA, the orbit insertion around Jupiter of the small orbiter which is separated from the mother spacecraft is included. In this paper, we investigated the feasibility of the aerocapture at Jupiter, assuming that a fraction of the Solar Sail membrane of the mother spacecraft was reused in order to insert the separated orbiter into the orbit around Jupiter., 資料番号: AA0063481062

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

Experimentally examined is the effect of applied magnetic field on a weakly ionized Argon plasma supersonic flow around a body around which the magnetic field is generated. To analyze the flow, applied is the absorption spectroscopic technique which enables determination of the translational temperature and the flow velocity. Based on the translational temperature distribution, combined with the electron excitation temperature distribution previously determined, the effect of the applied magnetic field on the behavior of the shock layer is discussed and it is confirmed that the applied magnetic field of which interaction parameter is 0.34 can affect the shock layer behavior and enhance the shock stand-off distance., 資料番号: AA0048089012

資料番号: AA0048089022

資料番号: AA0048089011

When a space vehicle returns into the earth atmosphere, the gas temperature around the vehicle becomes very high owing to mainly shock compression. To protect the vehicle from such high temperature environment, the problem on the aerodynamic heating should be analyzed precisely. In the present numerical computations, the high temperature air-flow with the real gas effects such as dissociation and ionization is studied to simulate hypersonic flows produced in a conical nozzle of the gun-tunnel placed at the Aoyama Gakuin University. The 2-D Euler code is used together with TVD scheme for shock discontinuity. To simulate the high temperature air-flow produced in a small hypersonic gun-tunnel located at Aoyama Gakuin University, numerical computations are performed by applying the 2-D Euler code with TVD scheme. The real gas effects of high temperature air with dissociation and ionization are taken into account., 資料番号: AA0048089026

資料番号: AA0048089013

When a space vehicle reenters onto the atmosphere with high velocity, radiative heating from the shocked air ahead of the vehicle plays an important role on the heat flux to the wall surface as well as convective heating. In the laboratory, the spectroscopic system has been developed for temperature measurement of radiation behind strong shock waves. In this paper, CARS method (Coherent Anti-Stokes Raman Spectroscopy) for radiative flow behind strong shock waves of high Mach number has been developed. The CARS measurement system consists of a YAG laser, a dye laser, optical system, and high-sensitivity I.I.-CCD camera systems with spectroscopy. The preliminary tests have been performed for the purpose of detecting CARS signal from strong shock waves with velocity range over 5 km/s. Finally, both vibrational and rotational temperatures could be estimated by fitting of the CARS spectra obtained from experiment and analytical calculation over Mach 15., 資料番号: AA0048089018

資料番号: AA0048089027

資料番号: AA0048089007

資料番号: AA0048089021

Detailed magneto-hydrodynamic (MHD) simulations of plasma flows in an inductively coupled plasma (ICP) heater are performed with air as a working gas. The plasma flows are numerically solved by the axisymmetric Navier-Stokes equations coupled with three-temperature models which extended Park's two-temperature models, and Dunn-Kang's chemical kinetic models to take into account the thermal and chemical non-equilibrium, while induction heating in plasmas is incorporated by solving the time-averaged MHD induction equations. Numerical results are compared with those of spectroscopic measurement. Good agreement is obtained in the downstream region where the flow is in local thermal equilibrium, however, qualitative discrepancy of electronic temperature between the theory and the experiment is found in the discharge chamber where Joule heating is significant., 資料番号: AA0048089023

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

資料番号: AA0048089008

近年、柔構造体が軽量で収納可能であるという特性から宇宙開発の分野で注目を集めている。その応用の1つとして、大気突入飛行体のエアロシェルとして利用することが提案されている。本研究では、フレア型の柔構造エアロシェルを再使用型宇宙輸送システムへ応用することを想定して、その利点や実現性について検討した。その結果、将来の宇宙輸送システムとして有用であることを示した。また、柔構造エアロシェルの実際のミッションヘの利用への第1歩として、2003年と2004年に宇宙研の大気球を利用して高度40kmからの落下試験を行った。本実験より、柔構造エアロシェルを有するカプセル型の機体が低亜音速から遷音速領域で安定して飛行することが実証された。最後に、さらに高速の飛行性能を実証するために観測ロケットを用いて飛行試験の提案を行う。, The flexible structures are promising in the field of the space development due to its flexibility and low mass. One of their applications is a membrane aeroshell for atmospheric entry vehicles. This study focuses on the flare-type membrane aeroshell. At first, a system study on the case that a membrane aeroshell is applied to a reusable space transportation system is conducted to clarify its merits and to discuss its feasibility. The group carried out the drop tests of the capsule with the membrane aeroshell from an altitude of 40 km using the large scientific balloons in 2003 and 2004. The stable flight in the subsonic to transonic regime is demonstrated in this test. The flight test using sounding rocket is proposed to investigate the flight performance in hypersonic regime and survivability in high aerodynamic heating environment., 資料番号: AA0047897024

資料番号: AA0048089028

資料番号: AA0048089037

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

ソーラーセイルミッションのサブミッションとして、木星の極磁場観測を実現させる小型カプセルのシステムを検討し、開発見通しを示した。直径約45cm、重量27kg、その重量の90％が耐熱材料という小型カプセルは、木星の長楕円周回軌道から木星突入を行う。, 資料番号: AA0047899047

資料番号: AA0047120005

資料番号: AA0047120024

資料番号: AA0047120040

資料番号: AA0047120006

資料番号: AA0047120041

資料番号: AA0047118034

資料番号: AA0047118045

地球大気への再突入を想定した, 高層の希薄な環境下における半球周りの熱・化学的に非平衡な極超音速流の構造を, ボルツマン方程式と等価な分子シミュレーションであるDSMC法により明らかにした。解析にあたっては, 解離反応モデルに既存の3種と, 著者の提案するVEDDモデルを適用した。本解析により, 反応モデルの性質によって流れ場の構造は大きく変化し, 衝撃波離脱距離に影響を与えることが確認された。解析結果より算出された衝撃波離脱距離と同一条件で実施された実験との比較においては, VEDDモデルによる衝撃波離脱距離およびその速度依存性が実験値と良く一致することが確認され, VEDDモデルの有効性が検証された。また, 淀み点への空力加熱量については, VEDDモデルによる値は, 実験との比較から妥当な推算値といえるが, その速度依存性からは, より詳細な解析の必要性が示唆される。

資料番号: AA0046651020

Hypersonic aerodynamic heating problem has been investigated for the shock-shock interactions of two hemi-sphere cylinder problem for future TSTO space transport systems. Two hemisphere models with a diameter of 3 cm and 6 cm are placed parallel or with the inclined angles and hypersonic wind tunnel tests have been made by NAL Mach 10 HWT. Severe local peak heating is generated on the small hemisphere nose by strong shock-shock interactions of Type 3 at the inclined angle of 15 deg. In the parallel setting cases at high angles of attack. Type 1 and Type 2 shock interaction patterns are observed. CFD computations can predict local peak heating accurately in the interaction area and the structure of shock-shock, shock-boundary layer interaction can be resolved in detail., 資料番号: AA0045948024, レポート番号: NAL SP-57

資料番号: AA0045913029