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機体騒音を計測する風洞試験で，ケブラー布を測定部の風洞壁に用いる方法が提案されている．壁のように流れを維持することが可能で，音は風洞の外に設けられた無響音室で計測することができる．JAXAでは，ケブラー壁を2m×2m低速風洞に導入した．ケブラー壁による壁干渉を補正するため，ケブラー布の透過速度の計測をした．Devenportらが提案している近似式を用いて，圧力差と透過流速の関係を精度良くモデル化できることが分かった．織り方の異なる2種類のケブラー布に対し，それぞれ無次元定数を求め，透過流速をモデル化することができた．さらに，風洞では強く引っ張られた状態で使用されているため，引張力の影響も調査した．その結果，引張力はあまり影響しないことが分かった．, In order to measure aeroacoustic noise in wind tunnel, a Kevlar wall technique is proposed. The Kevlar wall can keep the flow in the test section and it is acoustically transparent. JAXA installed the Kevlar wall in the test section of 2m×2m low speed wind tunnel. We measured transpiration velocities through Kevlar clothes for wall interference correction. It is found that the relationship between differential pressure and transpiration velocity can be modeled with an equation proposed by Devenport et al. Non-dimensional constants of the equation were obtained for two types of Kevlar clothes, and the transpiration velocities were modeled. In addition, the Kevlar cloth is stretched when it is used in wind tunnel, and therefore effects of the tension on the transpiration velocity were also investigated. However, we found that the tension does not affect the transpiration velocities., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1430003000, レポート番号: JAXA-RM-14-001

平成23年度宇宙航行の力学シンポジウム（2011年12月19日-20日. 宇宙航空研究開発機構宇宙科学研究所）, 相模原市, 神奈川県, Symposium on Flight Mechanics and Astrodynamics 2011 (December 19-20, 2011. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000006005

宇宙航行の力学シンポジウム 平成22年度 (2010年12月16-17日. 宇宙航空研究開発機構宇宙科学研究所), 相模原市, 神奈川県, Symposium on Flight Mechanics and Astrodynamics, 2010 (December 16-17, 2010. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: AA0065106020

平成21年度宇宙航行の力学シンポジウム（2009年12月10日-11日. 宇宙航空研究開発機構宇宙科学研究本部）, 相模原市, 神奈川県, Symposium on Flight Mechanics and Astrodynamics 2008 (December 10-11, 2009. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000007027

宇宙航行の力学シンポジウム(平成20年度)(2008年12月4～5日、宇宙航空研究開発機構宇宙科学研究本部), Symposium on Flight Mechanics and Astrodynamics 2008 (December 4-5, 2008. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)), Sagamihara, Kanagawa Japan, 資料番号: AA0064729037

It is still challenging to predict surface heat-transfer rate in hypersonic flow computations. In this paper, we first performed numerical experiments by changing numerical flux functions and meshes for a hypersonic flow around a hemisphere. Results show that AUSM+ flux function by Liou (1996) on a carefully refined mesh can lead to a shock stable solution with an accurate aerodynamic heating on the surface. Then, a numerical simulation on a hypersonic flow around two bodies involving a shock/shock interaction and a boundary-layer separation has been conducted by using the same method. Although this is a rather difficult problem with a complicated flowfield, comparisons show good agreement with the corresponding experimental data, including the surface heat-transfer rate profile. Therefore, we can say that we have established a numerical method to accurately predict surface heat-transfer in hypersonic shock interacting flows. Finally, detailed analysis of the computed flowfield has been made.

To explore future RLVs, a hypersonic flowfield around a Two-Stage-To-Orbit (TSTO) configuration is analyzed in this paper. This study also demonstrates how CFD, as a powerful tool, can be applied to investigate such a complex flowfield involving a shock/shock interaction and a boundary-layer separation. First, the hypersonic flow around two bodies of the TSTO model has been numerically simulated, and then, the results are validated by comparing with experimental data taken at the UT-Kashiwa Hypersonic Tunnel. Finally, the detailed, computed flowfield is shown to have pairs of streamwise vortices (including horseshoe vortices) with alternating signs of rotation around the TSTO body surfaces.

In this paper, a study on sound propagation of H-2A rocket plume duct with deflector is analyzed using a hybrid code of an Euler flow solver and a Linearized Euler Equations solver. Several cases of calculation for modeled jet configurations of H-2A main engine, 202, and 204 rockets have been conducted. The results of these calculations indicate that sound waves from the modeled duct have different characteristic frequency according to jet configuration., 資料番号: AA0063742034, レポート番号: JAXA-SP-07-016

Acoustic loads are the principal source of structural vibration and internal noise during launch or static-firing operations. The loads cause various problems, so it is important to predict the acoustic loads on space vehicles such as a rocket. Conventionally, the prediction has been made by empirical methods. These methods have some limitations, since shielding and reflection are difficult to be dealt with. To avoid these difficulties, we are planning to use our LEE (Linearized Euler Equation) code and evaluate the acoustic field to predict the acoustic loads on a rocket. As a first step, we report the results of some benchmark tests for validating our code's precision., 資料番号: AA0063742017, レポート番号: JAXA-SP-07-016

As base drag accounts for most of the total drag for spacecraft with a large base area, reduction in the base drag improves its aerodynamic performance. In the present study, a method using a pair of tabs normal to the base surface is proposed to reduce the base drag, and its effects on drag reduction are experimentally examined for a two-dimensional, simple bluff body with a large base area, the cross-section of which consists of a circular forebody and a rectangular afterbody, by changing the tab’s position and length. In addition, the flow fields around the model with/without tabs were visualized by the schlieren method with a high speed video camera, and CFD analysis was also made to supplement the experimental results. It was found that the tab method can remarkably increase the base pressure, which leads to a large decrease in the base drag. In the case where a pair of tabs with a non-dimensional length of 0.38 are installed at the base edges, the total drag is reduced by about 20.6% at a Mach number of 0.6 compared with that in the baseline case without tabs. Even when short tabs with a non-dimensional length of 0.07 are employed, about 10.0% reduction in the base drag is achieved. Thus, it was made clear from this study that the tab method proposed here is effective to decrease base drag in the transonic flow.

Supersonic flow fields around Two-Stage-To-Orbit (TSTO) models with different configurations have been experimentally examined in this paper. Four configurations for the orbiter have been considered: A) a hemisphere-cylinder, B) a hemisphere-cylinder with a flat bottom, C) an obliquely truncated circular cylinder, and D) a cone-cylinder. All the flow fields around these models showed complicated shock/shock and shock/boundary-layer interactions, which can be categorized into three patterns, depending on the extent to which the separation shock wave contributes to these interactions. The models B, C and D were proposed to suppress the pressure rise due to the interactions observed in the model A. As a result, the model B showed almost the same interactions as the model A, while in the model C they did not present. In the model D, a large pressure rise was seen in the case with no clearance, whereas the model undergoes the least aerodynamic interaction at a rather large clearance. It is concluded from these results that the model C is less affected by aerodynamic interactions due to the clearance than the other models.

資料番号: AA0063603005, レポート番号: JAXA-CR-07-001

平成18年度宇宙輸送シンポジウム（2007年1月18日-19日. 宇宙航空研究開発機構宇宙科学研究本部）, 相模原市, 神奈川県, Space Transportation Symposium FY2006 (January 18-19, 2007. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000005015

平成18年度宇宙航行の力学シンポジウム（2006年12月14日-15日. 宇宙航空研究開発機構宇宙科学研究本部）, 相模原市, 神奈川県, Symposium on Flight Mechanics and Astrodynamics 2006 (December 14-15, 2006. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000012026

A computational program using a CFD-CSD coupled method has been developed to study the problem of low supersonic curved panel flutter. So far the supersonic aerodynamic model has been commonly employed to compute flutter boundaries. However, the present CFD-CSD coupled method makes it possible to analyze flutter boundaries at all speeds including transonic speeds. The computed flutter dynamic pressure for M∞=3 agrees with others’ results. A curved panel flutter were analyzed at a low supersonic speed, M∞=1.2. The flutter for small curvature panels is first-mode flutter, which is similar to flat panel flutter, whereas it becomes higher-mode flutter for larger curvature plates. In this study, flutter boundaries for an access panel have also been simulated. Its critical dynamic pressure becomes significantly lower compared with the case of a panel with its all edges simply supported.

A supersonic jet into complex geometry and its accompanying acoustic field have been investigated by using Euler Equation/Linearized Euler Equation (LEE) hybrid methodology here. As a first step, we have examined its performance by computing traveling sinusoidal waves. Then, we found that the hybrid methodology achieves higher accuracy than the Euler computation when combined with a fifth-order WENO (Weighted Essentially Non-Oscillatory) scheme. Based on this knowledge, we have conducted both of fully third-order Euler/LEE and third-order-Euler/fifth-order-LEE hybrid computations. According to the results, the computed acoustic field is mainly composed of three waves having distinct frequencies: a high frequency wave (0.5 Hz) which seems to be caused by interferences from the wall, a medium frequency wave (2.2 Hz) corresponding to first-mode eigenfrequency of the jet exit geometry, and a low frequency wave (10 Hz). The computed pressure values are rather different between the two results, although these frequencies are not so much. This fact implies that the more we extend the region of fifth-order computation, the more the computed pressure values approach to physically correct ones in quantitative view., 資料番号: AA0063154016, レポート番号: JAXA-SP-06-010

This paper provides the results of numerical analysis on the sound radiation and frequency characteristics of a rocket frame deflector. In general, a deflector is built in a rocket Launch Pad (LP) in order to reduce the influences of exhaust plume on a rocket body. The rocket frame deflector is expected to be a low frequency sound source because of its large dimension. Since the low frequency sound wave is dangerous for rocket body at its natural frequency, it is necessary to make the frequency characteristics of deflector clear. The characteristic frequency analysis of a modeled rocket frame deflector is conducted using a FEM (Finite Element Method) solver. It can be seen from the FEM analysis that the strong resonance occurs at low frequency in the modeled deflector. Three dimensional numerical analysis of sound radiation from deflector including rocket plume inflow is also conducted on a modeled LP using a hybrid code of an Euler flow solver and a Linearized Euler solver. From the sound radiation analysis, the characteristic sound wave from a modeled deflector has low frequency., 資料番号: AA0063154020, レポート番号: JAXA-SP-06-010

平成17年度宇宙航行の力学シンポジウム（2005年12月12日-13日. 宇宙航空研究開発機構宇宙科学研究本部）, 相模原市, 神奈川県, Symposium on Flight Mechanics and Astrodynamics 2005 (December 12-13, 2005. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000009011

TSTO (Two Stage To Orbit) is a candidate for advanced space transportation system, which is required to fly over a wide region from subsonic to hypersonic speed with an orbiter connected with a booster. The hypersonic flow field with aerodynamic interaction around a TSTO model, which consists of a delta wing as a booster and a hemisphere-cylinder as an orbiter, has been investigated. The temperature distribution is measured by the Temperature Sensitive Paint (TSP) method, and the heat flux with thermo-couples. As a result, the flow field has been significantly affected by the clearance between the two bodies. In the case of a small clearance, heat flux on the head of the hemisphere-cylinder becomes about 3.1 times as large as the stagnation heat flux for the hemisphere-cylinder alone. On the other hand, at a large clearance, heat flux on the delta wing becomes about 1.1 times as large as the stagnation heat flux. In addition, a new thermal protection system (TPS) to reduce the aerodynamic heating due to shock/shock interaction and shock/boundary-layer interaction is also investigated in this paper. A triangular prism installed on the delta wing can change the aerodynamic heating, depending on its location. It was found that a small device installed upstream is effective, where the heat flux has been reduced by about 48%.

A method for computational aeroacoustics using the equations split into the incompressible flow part and the acoustic part is investigated in this paper. The equations derived based on Mach number expansion, and the method to solve them with high accuracy and stability are shown. The finite volume method using WENO scheme is applied to the acoustic field around a circular cylinder. Comparing the present method with the approximate solution based on the acoustic analogy, it is found that the acoustic field has been decayed according to the square root of the distance from the object further from about 10 times the diameter. In addition, 15 grid points per a wave length can give a sufficiently accurate solution by the present method.

As future space vehicles, Reusable Launch Vehicle (RLV) needs to be developed, where there are two kinds of RLV: Single Stage To Orbit (SSTO) and Two Stage To Orbit (TSTO). In the latter case, the shock/shock interaction and shock/boundary layer interaction play a key role. In the present study, we focus on the supersonic flow field with aerodynamic interaction between a delta wing and a hemisphere-cylinder, which imitate a TSTO, where the clearance, h, between the delta wing and hemisphere-cylinder is a key parameter. As a result, complicated flow patterns were made clear, including separation bubbles.

資料番号: AA0048089029

資料番号: AA0047120010

資料番号: AA0047118034

The aerodynamic sound generated from a protuberance has been simulated in this study. The protuberance, which is placed on a flat plate, has three kinds of shape at the leading edge. The propagation of sound waves is examined by solving the three-dimensional N-S equations to observe small pressure fluctuations both in the near and far fields. Numerical results show that vortices created from near the leading edge may make pressure waves, leading to acoustic waves. Moreover, FFT analysis shows that the frequency profile partly agrees with experimental data., 資料番号: AA0047427004, レポート番号: JAXA-SP-03-002

資料番号: AA0045426037

A numerical scheme to treat the problem of fluid-structure coupling has been developed in this study. A high order scheme based on the finite volume method (FVM) is used for fluid calculation, as well as the finite element method (FEM) for structure calculation in order to make more detailed analysis of flutter phenomena. The present method has been applied to the problem of panel flutter. For the supersonic case of M = 1.2, the flutter shows a rapid increase in amplitude into a limit cycle oscillation with a first mode. From the comparison between Euler and Navier-Stokes computation, it is clear that a boundary layer delays the growth of the flutter oscillation. For the subsonic case of M = 0.95, a static, upwardly divergent solution has been obtained, where vortices are produced due to the fluid-structure interaction in the Navier-Stokes computation., 資料番号: AA0045948020, レポート番号: NAL SP-57

Recently CFD is commonly used in design, where high efficiency is required. The cartesian grid has an advantage over other grids because it is easy to generate grid. However, many grid cells are necessary to correctly capture various flow phenomena, which spoils the merit of the cartesian grid. In the present study, an LU-SGS method is employed for time integration to overcome this problem., 資料番号: AA0045948045, レポート番号: NAL SP-57

The subsonic flow around a 45 deg sweep delta wing is numerically simulated to examine the effects of a hybrid method using both leading edge rotation and trailing edge jet on aerodynamic forces. Computation is performed at an angle of attack of 20 deg and a Reynolds number of 2.0 x 10(exp 4) based on the wing root chord. In the plain case, the flow is fully separated from the wing surface, and no suction peak due to leading edge vortices is observed. However, in the case with the hybrid method, the lift was confirmed to be increased., 資料番号: AA0045948051, レポート番号: NAL SP-57

平成14年度宇宙科学企画解情報解析シンポジウム（2003年2月10日. 宇宙科学研究所(ISAS)）, 相模原市, 神奈川県, Space Science Informatics Symposium FY2002 (February 10, 2003. Institute of Space and Astronautical Science(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000112006

衝撃風洞試験において、極超音速流中における2物体間空力干渉流れ場の測定を行い、以下のような結論を得た。(1)感圧塗料を用いた圧力分布の測定法を用いることにより、比較的少ない実験回数で2物体間の干渉流れ場を詳細に観察することができた。(2)導電性塗料とPoly-Urethaneを用いて作成した薄膜ゲージは現在のところ測定精度が不十分で、作製法や母材材料の見直しが必要である。, 資料番号: AA0033392030

The low-speed flow for a 45 deg sweep delta wing is numerically simulated to see the effects of leading-edge rotation on aerodynamic forces. In the present study, cylindrical leading edges are employed, where the portion from x = 0.25 C to C can be rotated. The rotational speed can be varied arbitrarily. The computation is performed at an angle of attack of 20 deg and a Reynolds number of 2.0 x 10(exp 4) based on the wing root chord. The flow is fully separated from the wing surface in the case of no the leading edge rotation, so that no suction peaks due to leading edge vortices appear over the wing surface, which means that the delta wing will not have an effective vortical lift. On the other hand, in the case of rotating the leading edges, leading edge vortices can successfully by produced, resulting in a remarkable increase in lift., 資料番号: AA0032819045, レポート番号: NAL SP-53

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県, Space Transportation Symposium FY2013 (January 16-17, 2014. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000016053, レポート番号: STCP-2013-054

航空宇宙技術研究所 7-9 Jun. 2000 東京 日本, National Aerospace Laboratory 7-9 Jun. 2000 Tokyo Japan, 2次元の翼周りの流れを低速度でシミュレーションした。この研究では、定常解または準定常解を求めるために、2つの方法を試みた。マトリックスフリーのLU-SGS(Lower-Upper Symmetric Gauss Siedel)陰解法と、陰解法と陽解法を結合したハイブリッド法である。乱流モデルについては、ボールドウィン-ロマックス・モデルまたは乱流のないモデルを、圧縮型ナビエ-ストークス方程式に適用した。これらの計算結果は、実験データとよく合っていた。, The flow around a 2-D (two dimensional) airfoil is simulated at a low speed. In this study two schemes are considered: the matrix-free LU-SGS (Lower-Upper Symmetric Gauss Siedel) implicit scheme and the explicit/implicit hybrid scheme to obtain steady or quasi-steady state solutions. Regarding turbulence model, either the Baldwin-Lomax model or no turbulence model is applied to the compressible Navier-Stokes equations. Results of these calculations are well compared with experimental data., 資料番号: AA0028635024, レポート番号: NAL SP-46

The characteristics of a plate burner and a heat exchanger are experimentally investigated in this paper. The present burner consists of a stack of ceramic flaments catalyzed with Pd, which is featured by low concentrarions of NOx and CO. Furthermore, the burner surface emits light and thermal radiation in a high activity region of catalyst. Therefore, the thermal efficiency, that is, heat input to the heat exchanger with regard to combustion load, is expected to increase. In this study the combustion characteristics of this burner and the efficiency of the heat exchanger have been examined to see the efficiency of the system. Two major results were obtained ; one is that the thermal efficiency has been improved by about 5 ％, and the other is that unburned combustion gas after the heat exchanger was reduced because combustion starts on the burner surface. Thus it is confirmed that the catalytic burner is useful for energy saving and downsizing of a combustor.

航空宇宙技術研究所 16-18 Jun. 1999 東京 日本, National Aerospace Laboratory 16-18 Jun. 1999 Tokyo Japan, 3次元物体周りの非粘性流の高速な自動化解析のためのCartesian格子生成法を開発した。この方法は固体表面とCartesian格子の間の断面が簡単でない場合を考慮することができる。従って薄い物体や鋭角または複数固体領域などをセル内に含む場合を厳密に取り扱うことができる。このようなケースはしばしば超音速輸送機(SST)モデルの計算に現れてくる。計算効率を上げるために、セルの内側の物体表面を表すために擬似平面近似を行い、その結果この適用例では全物体表面の要素を約3分の1に減らすことができた。曲面のまわりの局所細分化格子について新アルゴリズムを使うことによってより高い計算効率を達成できる。, A Cartesian grid generation method has been developed for fast and automated analysis of inviscid flows around three dimensional bodies. It can take into account non-simple cases of intersection between a solid surface and a Cartesian grid, so that a cell containing a thin body, a sharp edge or multiple solid regions is properly handled. These cases appear frequently in the computation of a Supersonic Transport (SST) model. To increase the efficiency of computation, a pseudo-planar approximation is made to represent a body surface inside a cell, resulting in a reduction of total body surface elements by a factor of about three in this application. Higher efficiency is achieved using a new algorithm for local grid refinement around a curved surface., 資料番号: AA0001961042, レポート番号: NAL SP-44

航空宇宙技術研究所 7-9 Oct. 1998 (23th). 24-26 Mar. 1999 (24th) 東京 日本, National Aerospace Laboratory 7-9 Oct. 1998 (23th). 24-26 Mar. 1999 (24th) Tokyo Japan, 共軸環状ノズルを流れる非定常流れ場を数値計算した。正確なデータを得るために、計算は高次精度の数値計算コードを用いて行った。狭い隙間は超音波ノズルの形をしており、衝撃波が境界層との干渉の結果発生する。このために連続的に振動する流れが生成される。この現象はある種の擬似衝撃波不安定性と考えることができる。, The unsteady flow field in a coaxial annular nozzle is numerically calculated. To obtain accurate data, calculation is performed by using a high-order accuracy numerical code. The narrow clearance takes a shape of supersonic nozzle, where shock waves are formed as a result of interacting with the boundary layer. This causes the flow to oscillate continuously. This phenomenon is considered to be a kind of pseudo-shock instability., 資料番号: AA0001960032, レポート番号: NAL SP-43

航空宇宙技術研究所 6-7 JUN. 1996 東京 日本, National Aerospace Laboratory 6-7 JUN. 1996 Tokyo Japan, マッハ数1.5の場合の平板上の圧縮性境界層内に生成する擾乱の空間的発達のシミュレーションをナビエ・ストークス方程式を解いて行った。4次非振動スキームおよびRoeの近似リーマン解法を空間離散化に、3次Runge-Kuttaスキームを時間積分に用いた。上流域で与えた擾乱は流れ方向とスパン方向に広がった。壁温を変えて擾乱発達に及ぼす壁面冷却効果を調べた。擾乱は擾乱源の下流にくさび形領域を形成して伝搬することが分かった。くさび角と流れ方向速度の変動の大きさは壁冷却により増加した。横方向広がり角は実験データの角度に一致した。, The spatial development of a disturbance generated in the compressible boundary layer on the flat plate is simulated in the case of a Mach number of 1.5 by solving the Navier-Stokes equations. The fourth-order nonoscillatory scheme and Roe's approximate Riemann solver are utilized for space discretization and the third-order Runge-Kutta scheme for time integration. The disturbance given in the upstream region spreads in the streamwise and spanwise directions. The wall temperature is varied to examine the effects of wall cooling on the development of disturbance. It is found that the disturbance propagates in the wedge-shaped region downstream of the source. The wedge angle and magnitude of the streamwise velocity fluctuation are increased by wall cooling. Furthermore, the laterally-spreading angle is close to that of the experimental data., 資料番号: AA0000685044, レポート番号: NAL SP-34

自由振子運動する円柱のまわりの流れの数値シミュレーションを行った。この運動では物体に空気力が働いて運動に影響する。本件では、この力によって運動は減衰する。この問題につきナビエ･ストークス方程式を(u,v,p)定式化に対する有限差分/有限容積法により解いた。圧力に関するポアソン方程式を解くために多段階格子法も用いた。結果はポテンシャル方程式の解に比べ良好である。, A numerical simulation is presented for the flow around a circular cylinder in a free-pendulum motion. In this motion, the aerodynamic forces are exerted on the body and affect the body motion. In this case, the motion will damp out due to its forces. For this problem, the Navier-Stokes equations are solved by a finite difference-finite volume method, where the n-v-p formulation is employed. A multi-grid technique is also employed to solve the pressure Poisson equation. The results are well compared with potential flow solutions., 資料番号: AA0004174048, レポート番号: NAL SP-27

Roeの近似リーマン解法とMUSCL型の外挿法を用いて、OREXのまわりの圧縮性･粘性流れの数値解法を行った。計算は、マッハ数7.1,レイノルズ数6.518×10{6}で行った。, A numerical analysis of compressible and viscous flow around the Orbital Reentry Experiment (OREX) has been carried out with the Roe's approximate Riemann solver and Monotone Upstream-centered Scheme for Conservation Law (MUSCL) type extrapolation. The calculation was performed at Mach number 7.1 and Reynolds number 6.518 x 10(exp 6)., 資料番号: AA0004173006, レポート番号: NAL SP-26

第42回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム2010 (2010年6月24日-25日. 米子コンベンションセンター BiG SHiP), 42nd Fluid Dynamics Conference / Aerospace Numerical Simulation Symposium 2010 (June 24-25, 2010. Yonago Convention Center BiG SHiP), Yonago, Tottori Japan, Although the flow rate through many holes on a transonic wind tunnel wall is very small， this flow makes a significant influence on the wind tunnel test results． However there are less data which show the relationship between the flow rate and the differential pressure through holes at this flow conditions． In this study， the relationship between the flow rate and the differential pressure through a hole on a transonic wind tunnel wall is experimentally investigated． In the case that the flow rate and the differential pressure are very small， the results show that their relationship is linear． Furthermore the slope of the line measures up to the results of CFD．, 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064967045, レポート番号: JAXA-SP-10-012

航空宇宙技術研究所 10-12 Jun. 1992 東京 日本, National Aerospace Laboratory 10-12 Jun. 1992 Tokyo Japan, 今日、航空機や宇宙機の全機形状まわりのナビエ・ストークス計算が手の届くものとなり、計算空気力学(CFD)の進歩には目を見張るものがあるが、一方では、その信頼性の確立という点では、まだ解決すべき多くのことが残されているように思われる。ある特定の課題を設定し、いくつかの異なるCFDプログラムを走らせた結果を持ち寄って、風洞試験データとの比較などを通してそれらを評価するのも、信頼性確立に向けた努力の一環と考えられる。今回、2次元翼の空力特性計算、および全機形状まわりのナビエ・ストークス計算という課題を設定して、ワークショップを行った。本文は、2次元部門の課題「Garabedian-Korn 75-06-12翼型の空気力学的特性解析」について、課題の詳細、すなわち、設計点の解析、設計点近傍の空力特性など、の課題募集要項のオーガナイザによる説明、7件の応募論文、オーガナイザによる2次元部門のまとめ、および、資料編として、応募者氏名解法情報一覧、応募ケース一覧、結果一覧表、応募者提出図などを掲載する。, In the first part of this paper, description of 2D Problem is given. The Problem A is to compute a shock-less flow of Garabedian-Korn 75-06-12 Airfoil at design condition of Mach number, 0.75, C(sub 1), 0.63, Reynolds number, 6 and 20 x 10(exp 6). The Problem B is to obtain the aerodynamic characteristics of Mach number sweep and angle of attack sweep. The reference point is the same condition in Problem A. The output data format is described. Together with given problems, this subsidiary includes a collection of nine papers which are presented by applicants as the analysis results of two dimensional aerodynamic problems., 資料番号: AA0004169001, レポート番号: NAL SP-20

著者らは，定常音響構造連成解析において，従来の手法では解析が困難な中間周波数帯を解析可能な決定論的手法である波動ベース法（WBM）に着目し，フェアリング内部の宇宙機の2 次元内部定常音響振動連成解析を行ってきた．本報告では，著者らが開発し拡張した2 次元解析コードを用いて，内部音響透過問題と，解析領域の広い外部音響伝播問題に対するWBM の適用性を検討した結果を示す．音響透過問題に関しては，単純な弾性板を通じた音響透過損失の計算を行い，有限要素法(FEM) による解析結果と比較することによってWBM の解析精度の検証を行った．その際，WBM とFEM の両者に対して，弾性板の端部における不連続な境界条件に付随して解析精度が悪化する問題を指摘し，簡易的な手法で問題を回避する方法を提案した．一方，外部問題においては，ロケット打ち上げ時の射場の定常音響伝播の計算を行い，無限遠において音が反射しないことを保証する放射条件を厳密に満足して，広い解析領域および広い周波数領域における問題を扱うことができることを示した．, Acoustic environment at a launch site is affected by jets from rocket engines, as well as by geometric features of a launch pad. Then, spacecraft are also exposed to acoustic pressure with the wide frequency range transmitted through a payload fairing. In the case that vibro-acoustic response is predicted using steady-state vibroacoustic analysis, there actually exists the mid-frequency range in which there are no mature numerical methods. This report deals with the novel Wave Based Method (WBM), a deterministic approach for steady-state vibroacoustic analysis in the wide frequency range. Then the two-dimensional (2D) WBM code developed by the authors is applied to some sound transmission problems and exterior problems. In the sound transmission analysis, it is found that discontinuous impedance boundary conditions (BCs) must be fi tted by using continuous functions to obtain more accurate results using both WBM and FEM. After smoothing the BCs, numerical prediction results of our sound transmission models by the 2D WBM and FEM are compared, and the 2D WBM are verified. Furthermore, in order to verify the 2D code extended for the exterior problems, it is applied to simple launch pad models. In both propagation and transmission problems, it can be stated that the WBM has quite high potential to be applied in the higher frequency range., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064697000, レポート番号: JAXA-RR-09-008

航空宇宙技術研究所 10-12 Jun. 1992 東京 日本, National Aerospace Laboratory 10-12 Jun. 1992 Tokyo Japan, 新らしい変数qを導入しその回転が運動量方程式の対流項および圧力勾配項を表わすような非圧縮ナビエ・ストークス方程式の新らしい定式化を提案した。得られた支配方程式は、線形の力学的方程式および非線形の非同次項をもつ新変数qに対するポアソン方程式より成る。これらの方程式は発散なしの条件を自動的に満す。新変数qに対する境界条件はディリクレ型であり、ポアソン方程式は従来の点緩和法で速かに解くことができる。さらに、対応する圧力場を効率よく求めるための新らしい方法を提案した。空洞内流れおよび円柱まわりの流れに適用した結果は、本方法の優秀さを示す。, This paper proposes a new formulation for the incompressible N-S (Navier-Stokes) equations by introducing a new variable q, the rotation of which represents the convective and pressure gradient terms of the momentum equations. The derived governing equations consist of the linear dynamic equations and the Poisson equation for the new variable q with the non-linear source. These equations automatically satisfy the divergence free condition. The boundary condition for the new variable q is the Dirichlet type, the Poisson equation can be quickly solved by the conventional point relaxation method. Furthermore, a new method to economically obtain the corresponding pressure field is also proposed. Applications to a driven cavity flow and the flow around a circular cylinder show excellent performances of the new method., 資料番号: AA0004168025, レポート番号: NAL SP-19

ロケット打ち上げ時に生じる構造振動や内部騒音の主因となる音響荷重を予測することは重要である。この予測は、従来、経験的手法によって行われてきたが、そうした手法では音響伝播過程における遮蔽や反射が考慮されていない。こうした問題を解決するために、我々はEuler/LEEコードを開発し、H-IIAの打ち上げ射場の音響評価に応用してきた。本報告では、第1報に引き続き当コードのLEEオプションの使用方法を解説する。, Acoustic loads are the principal source of structural vibration and internal noise during launch. Therefore, it is important to predict the acoustic loads on spacecraft such as a rocket. Conventionally, the prediction has been made by empirical methods. In these methods shielding and reflection are not considered. To solve these problems, we have been developing an Euler/LEE(Linearized Euler Equation) hybrid code and applying it to the acoustic evaluation of H-IIA's launch pad. In this report, we explain how to use the LEE option of the hybrid code., 形態: カラー図版あり, 資料番号: AA0064260000, レポート番号: JAXA-RM-08-009