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Through advanced detail simulation tests, adjusting model mass and control factors made successful measurement of aerodynamic force on a sphere at Reynolds numbers around critical one at the JAXA 60 cm magnetic suspension and balance system. The obtained averaged aerodynamic force for 8 seconds shows that the minimum drag coefficient after the critical Reynolds number reaches 0.062 and that any hysteresis phenomenon on drag coefficient with respect to Reynolds number was not observed around the critical one. Through the obtained fluctuations of aerodynamic force, the shrunk force fluctuating around the sphere axis parallel to main flow and some force swirling the sphere about the axis for a short time were observed at Reynolds numbers larger than the critical one. And some oscillating force on a plane including the axis was observed in the test of a sphere with double tapes for boundary layer trip in large Reynolds number region around the critical one., 資料番号: AA0063742005, レポート番号: JAXA-SP-07-016

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 study of the aerodynamic design for the reduction of sonic-boom is conducted by combining a three-dimensional Euler CFD (Computational Fluid Dynamics) code with a least-square optimization technique. This method can consider a nonlinear and a three-dimensional effects of aircraft which cannot be considered by previous low-boom design methods based on linear theory. The fuselage geometry of a low-boom aircraft designed by linear theory is modified by this design method in order to minimize the pressure discrepancies between a target low-boom pressure signature and a calculated signature. It is shown that the sonic-boom intensity of the aircraft designed by this method is smaller than that designed by linear theory. The results of the study indicate that this method is a useful tool for low-boom design., 資料番号: AA0063609016, レポート番号: JAXA-SP-06-029E

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

Retirement of the space shuttle in 2010 was already decided and there will be some difficulties in transportation from ISS (the International Space Station) from 2010 until the next generation new transportation system CEV (Crew Exploration Vehicle) will be operative. JAXA started concept studies of re-entry systems that will be loaded in HTV (H-2 Transfer Vehicle) and will be ejected after its deorbit. A minimum but affordable capsule with the capability of carrying 50 kg payload was selected from various kinds of concept. System study was conducted and it showed that the capsule is feasible but has some technical problems to be solved., 資料番号: AA0049056018, レポート番号: JAXA-SP-05-010

航空宇宙技術研究所 8 Jun. 1995 東京 日本, National Aerospace Laboratory 8 Jun. 1995 Tokyo Japan, NALでは1995年6月高エンタルピー流れのワークショップを開催し、次の課題を検討した。高温衝撃風洞の駆動気体汚染、単純な物体の周りの高エンタルピー流計算、OREX(軌道再突入実験)の空力、空気熱力学データとOREX飛行実験、衝撃波誘起非定常空力加熱現象、CFD(計算流体力学)と極超音速試験、球体および鈍頭円錐モデルの高エンタルピー空気流、OREX形状、鈍頭体回りの熱化学非平衡極超音速流、OREX回りの極超音速流れ場、OREXおよび双曲面フレア回りの極超音速流、双曲面フレア体回りの極超音速流、など。, The following topics were discussed at the symposium which held in NAL on June, 1995: driver gas contamination in a shock tunnel, high enthalpy flow computation around the simple bodies, aerodynamics of OREX (Orbital Reentry Experiment), aerothermodynamic results and OREX flight experiments, shock wave induced unsteady aerodynamic heating phenomena, CFD (Computational Fluid Dynamics) and hypersonic testing, high enthalpy air flow computation with a sphere and a blunted cone models, OREX configuration, thermochemical nonequilibrium hypersonic flow around blunt body, the hypersonic flow fields around OREX, hypersonic flow around OREX and hyperboloid flare, etc., 資料番号: AA0000110000, レポート番号: NAL SP-29

航空宇宙技術研究所 8 Jun. 1995 東京 日本, National Aerospace Laboratory 8 Jun. 1995 Tokyo Japan, 軸対象ナビエ・ストークス方程式を解いてOREX(軌道再突入実験)および双曲面フレア回りの極超音速流解析を行った。この流れの解法は、有限体積陰的全変動減少(TVD)風上スキームに基づいている。対流束は、“カーバンクル現象”を除くためにAUSM(移動上流分割法)DVスキームで評価した。これは鈍頭体回りの極超音速流れの解析によく見かける非物理的解法である。また平衡空気の熱力学的および輸送特性を推定するのに曲線あてはめを採用した。OREXの流れの状態は、高度56.9kmでの飛行条件に対応しており、そこでは実在気体効果が無視されることはない。それ故に、この流れは完全気体であるばかりでなく平衡気体であると想定した。双曲面フレアの場合、流れの状態は低温気体を使用するGottingenのLudwigチューブでの実験に対応している。この課題の数値解析結果は、衝撃波と境界層間の相互作用によって引き起こされる分離を含む極超音速流の空気力学および空気熱力学特徴を正確に予測するために、流れの非定常性を考慮しなければならないことを示唆している。, Hypersonic flow analysis around OREX (Orbiter Reentry Experiment) and hyperboloid flare have been carried out by solving axisymmetric Navier-Stokes equations. The flow solver is based on a finite volume implicit TVD (Total Variation Diminishing) upwind scheme. Convective fluxes are evaluated by AUSM (Advection Upstream Splitting Method) DV scheme in order to remove carbuncle phenomena, that are unphysical solution often appeared on hypersonic flow analysis around a blunt body. A curve fitting were also adopted to estimate thermodynamic and transport properties of an equilibrium air. Flow condition for OREX corresponds to a flight condition at altitude 56.9 km, where real gas effect should not be ignored. Therefore, it was assumed that this flow was not only to be a perfect gas, but also to be an equilibrium air. In the case of hyperboloid flare, flow condition corresponds to experiment by Gottingen Ludwig tube used cold gas. Numerical result of this problem suggests that the flow unsteadiness must be taken into account, in order to predict accurately the aerodynamic and aerothermodynamic characteristics of a hypersonic flow including separations induced by the interaction between a shock wave and boundary layer., 資料番号: AA0000110012, レポート番号: NAL SP-29

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