青山 剛史

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

ロケットで打ち上げられる宇宙機には、ロケットとの間のインターフェイスを通じて機械振動が加わる。さらに、広い周波数成分をもつ音圧がフェアリングを透過して宇宙機表面に加わることにより振動する。太陽電池パドルやアンテナなど軽量で大きな面積をもつ構造や、比較的高い固有振動数をもつコンポーネントは、音響荷重に影響されやすい。しかし、衛星のコンポーネントやシステムの音響試験は行なわれているが、フェアリングまでを含めた衛星の十分な音響試験及び解析は行なわれていないのが現状である。そこで、解析によって、事前に全系の検討が十分に行われることが期待されている。JAXA では、フェアリングまでを含めた衛星の音響解析手法の確立を目指し、今までに2次元WBM をフェアリング内部の衛星の音響連成問題に適用し、その応用性について検討してきた。また、解析コードの検証を目的として簡易フェアリング部材を用いた音響透過実験を、名古屋大学と

平成21年6月18-19日. 宇宙航空研究開発機構調布航空宇宙センター形態: カラー図版あり資料番号: AA0064672023レポート番号: JAXA-SP-09-011

Numerical studies were carried out to investigate the effects of turbulent boundary layers on panel flutter at supersonic speeds. In this study, Reynolds-averaged Navier-Stokes equations were solved to take into account the turbulent boundary layer and its viscous effects. First, the fluid-structure coupling code was validated. The computed flutter boundaries agreed well with experimental data. Moreover, the results showed that the viscous effects were important and should be taken into account for flutter computation. Second, the boundary-layer effects were investigated in the Mach number range of 1.0-2.4. We compared the Reynolds-averaged Navier-Stokes computation with the inviscid computation and discussed the differences between them. We found that the boundary layer has not only a stabilizing effect but also a destabilizing effect, depending on the Mach number. The most important finding is that the flutter dynamic pressure slowly increases due to the boundary layer as the Mach number increases. In addition, the design boundary methodology was reviewed in terms of the turbulent boundary-layer effect, which will be helpful for the development of a new boundary-layer correction for the design boundary. Copyright © 2009 by the authors.

Numerical studies were carried out to investigate the effects of turbulent boundary layers on panel flutter at supersonic speeds. In this study, Reynolds-averaged Navier-Stokes equations were solved to take into account the turbulent boundary layer and its viscous effects. First, the fluid-structure coupling code was validated. The computed flutter boundaries agreed well with experimental data. Moreover, the results showed that the viscous effects were important and should be taken into account for flutter computation. Second, the boundary-layer effects were investigated in the Mach number range of 1.0-2.4. We compared the Reynolds-averaged Navier-Stokes computation with the inviscid computation and discussed the differences between them. We found that the boundary layer has not only a stabilizing effect but also a destabilizing effect, depending on the Mach number. The most important finding is that the flutter dynamic pressure slowly increases due to the boundary layer as the Mach number increases. In addition, the design boundary methodology was reviewed in terms of the turbulent boundary-layer effect, which will be helpful for the development of a new boundary-layer correction for the design boundary. Copyright © 2009 by the authors.

For a two dimensional analysis of sound transmission to structural model, it is considered that one sine sound pulse incidents on a curved plate in the Cartesian coordinate. Because of using the square mesh, the curved surfaces of plate are modeled by steps, which are small enough to the wave length of the incident sound wave. From the result of sound transmission calculation using the FDTD (Finite Difference Time Domain) method, it can be seen that the transmitted waves from the curved plate are generated by lateral and bending waves in solid. The results of FFT analyses of transmitted waves make the characteristic frequencies clear. These frequencies almost coincide with the symmetrical mode frequencies obtained by MSC NASTRAN. Therefore, the FDTD analysis can simulate that the transmission waves through the curved plate are produced by its mode shapes.

Digital Wind Tunnel (DWT) is being developed in a project entitled Digital/Analog Hybrid Wind Tunnel in JAXA. The main objective of this project is to sophisticate the transonic wind tunnel in JAXA by applying IT technologies. We have several hurdles to establish DWT such as the developments of an automatic grid generation tool and a high performance CFD solver. In this paper, the prediction accuracy of aircraft loads by using the automatic grid generation tool and the modeling of porous wall for the CFD calculations of whole wind tunnel including an aircraft model, support instruments, and walls are discussed.

A simulation method for full helicopter configuration is constructed by combining an unsteady Euler code and an aeroacoustic code based on the Ffowcs-Williams and Hawkings formulation. The flow field and helicopter noise are calculated using a moving overlapped grid system, and the mutual effect of main rotor and tail rotor are studied for the helicopter in hover or forward flight. In the hovering flight calculation, the tip vortex of the tail rotor is dragged by the induced flow of the main rotor, and the detailed phenomena of the flow pattern are captured well. In a forward-flight calculation, noises from the main rotor and tail rotor are predicted to understand the tail rotor noise for both self noise and the interaction noise with the main-rotor wake. Comparison of noise magnitude shows that the relative importance of tail rotor noise with respect to the main rotor noise according to the flight conditions.

In transient flight, rotor wakes and tip vortex generated by unsteady blade air-loads and blade motions are fully unsteady and 3-dimensionally-aperiodic, giving rise to significant complicity in accurate analysis compared to steady flight. We propose a hybrid approach by splitting the motions of a maneuvering helicopter into translation and rotation. Translation is simulated using a non-inertial moving (translating) coordinate for which new governing equations are derived, and rotations are simulated by moving each grid in the frame. A flow simulation (CFD) code is constructed by using the hybrid approach, then two simple cases (accelerating/decelerating flight and right-turn flight) for maneuvering helicopter are calculated using the moving overlapped grid method, which is now one of the most advanced techniques for tip-vortex capture. The vortex bundling phenomena, which is a main characteristic of right-turn flight, is well captured by the simulation code. The results of the present study provide better understanding of the characteristics for maneuvering rotorcraft, which can be valuable in full helicopter design.

Predictions of helicopter BVI noise using three-dimensional Euler code with a single blade grid are conducted under three different conditions: BVI noise caused by (1) interaction between rotating blades and vortex shed from fixed wing vortex generator, (2) interaction between rotating blades and tip vortices shed from preceding blades, and (3) interaction between rotating blades with elastic deformation and shed tip vortices. In the CFD calculation, the Field Velocity Approach (FVA) and Scully's vortex model are used to import the wake information into the calculation grid and to determine the induced velocity made by tip vortices, respectively (cases 1-3). Beddoes generalized wake model is used to prescribe the tip vortices position in the wake (cases 2 and 3). Information about blade elastic deformation is imported from HART II project experimental data into the calculation (case 3). Acoustic analyses based on Ffowcs-Williams and Hawkings (FW-H) equation are conducted subsequently in each case. The results front the calculations show good agreement with experiments in all three cases, indicating that application of FVA, Scully's model, and Beddoes generalized wake model is effective for BVI noise prediction in this study, which is intended for low calculation cost using a single blade grid. Also, use of blade elastic deformation data in the calculation shows marked improvement in calculation precision. Consequently, the method used in this study can predict BVI noise under various conditions of wake or blade deformation with acceptable precision and low calculation cost.

In transient flight, rotor wakes and tip vortex generated by unsteady blade air-loads and blade motions are fully unsteady and 3-dimensionally-aperiodic, giving rise to significant complicity in accurate analysis compared to steady flight. We propose a hybrid approach by splitting the motions of a maneuvering helicopter into translation and rotation. Translation is simulated using a non-inertial moving (translating) coordinate for which new governing equations are derived, and rotations are simulated by moving each grid in the frame. A flow simulation (CFD) code is constructed by using the hybrid approach, then two simple cases (accelerating/decelerating flight and right-turn flight) for maneuvering helicopter are calculated using the moving overlapped grid method, which is now one of the most advanced techniques for tip-vortex capture. The vortex bundling phenomena, which is a main characteristic of right-turn flight, is well captured by the simulation code. The results of the present study provide better understanding of the characteristics for maneuvering rotorcraft, which can be valuable in full helicopter design.

Predictions of helicopter BVI noise using three-dimensional Euler code with a single blade grid are conducted under three different conditions: BVI noise caused by (1) interaction between rotating blades and vortex shed from fixed wing vortex generator, (2) interaction between rotating blades and tip vortices shed from preceding blades, and (3) interaction between rotating blades with elastic deformation and shed tip vortices. In the CFD calculation, the Field Velocity Approach (FVA) and Scully's vortex model are used to import the wake information into the calculation grid and to determine the induced velocity made by tip vortices, respectively (cases 1-3). Beddoes generalized wake model is used to prescribe the tip vortices position in the wake (cases 2 and 3). Information about blade elastic deformation is imported from HART II project experimental data into the calculation (case 3). Acoustic analyses based on Ffowcs-Williams and Hawkings (FW-H) equation are conducted subsequently in each case. The results front the calculations show good agreement with experiments in all three cases, indicating that application of FVA, Scully's model, and Beddoes generalized wake model is effective for BVI noise prediction in this study, which is intended for low calculation cost using a single blade grid. Also, use of blade elastic deformation data in the calculation shows marked improvement in calculation precision. Consequently, the method used in this study can predict BVI noise under various conditions of wake or blade deformation with acceptable precision and low calculation cost.

In this study, the relationship between the characteristics of sound transmission and acoustically induced vibration has been experimentally investigated. It was found that sound transmission is affected by the proper oscillation of a material. In particular, at the natural frequencies with axisymmetircal modes, lower transmission peaks have been observed.

An unsteady two-dimensional inviscid blade-vortex interaction has been calculated using the finite difference lattice Boltzmann method of the compressible Enter model. The perturbed discrete Boltzmann equation based on a prescribed vortex approach has been proposed to prevent a vortex from diffusing by numerical dissipation. The discretization of the governing equation is based on a second-order-accurate explicit Runge-Kutta time integration and a fifth-order-accurate upwind scheme that includes additional terms to capture the shock waves clearly. Subsonic and transonic flows around an airfoil were simulated to validate the perturbed discrete Boltzmann equation system. The numerical results were compared with other numerical data, and good agreement has been obtained. In the simulation of the transonic blade-vortex interaction, an instantaneous pressure coefficient, a time history of a lift coefficient, and patterns of acoustic waves were compared with other numerical results, and were found to agree with them very well. We have also investigated a generation mechanism of acoustic wave caused by the blade-vortex interaction, the effect or the flow Mach number, and the influence of the vortex miss distance.

An unsteady two-dimensional inviscid blade-vortex interaction has been calculated using the finite difference lattice Boltzmann method of the compressible Enter model. The perturbed discrete Boltzmann equation based on a prescribed vortex approach has been proposed to prevent a vortex from diffusing by numerical dissipation. The discretization of the governing equation is based on a second-order-accurate explicit Runge-Kutta time integration and a fifth-order-accurate upwind scheme that includes additional terms to capture the shock waves clearly. Subsonic and transonic flows around an airfoil were simulated to validate the perturbed discrete Boltzmann equation system. The numerical results were compared with other numerical data, and good agreement has been obtained. In the simulation of the transonic blade-vortex interaction, an instantaneous pressure coefficient, a time history of a lift coefficient, and patterns of acoustic waves were compared with other numerical results, and were found to agree with them very well. We have also investigated a generation mechanism of acoustic wave caused by the blade-vortex interaction, the effect or the flow Mach number, and the influence of the vortex miss distance.

A simulation method for full helicopter configuration is constructed by combining an unsteady Euler code and an aero-acoustic code based on the Ffowcs-Williams and Hawkings formulation. The flow field and helicopter noise are calculated using a moving overlapped grid system, and the mutual effect of main rotor and tail rotor are studied for the helicopter in hover or forward flight. In the hovering flight calculation, the tip vortex of the tail rotor is dragged by the induced flow of the main rotor, and the detailed phenomena of the flow pattern are captured well. In the forward-flight calculation, noises from the main rotor and tail rotor are predicted to show tail rotor noise for both self noise and the interaction noise with the main-rotor wake. Comparison of noise magnitude shows the relative importance of tail rotor noise according to flight conditions.

A simulation method for full helicopter configuration is constructed by combining an unsteady Euler code and an aero-acoustic code based on the Ffowcs-Williams and Hawkings formulation. The flow field and helicopter noise are calculated using a moving overlapped grid system, and the mutual effect of main rotor and tail rotor are studied for the helicopter in hover or forward flight. In the hovering flight calculation, the tip vortex of the tail rotor is dragged by the induced flow of the main rotor, and the detailed phenomena of the flow pattern are captured well. In the forward-flight calculation, noises from the main rotor and tail rotor are predicted to show tail rotor noise for both self noise and the interaction noise with the main-rotor wake. Comparison of noise magnitude shows the relative importance of tail rotor noise according to flight conditions.

資料番号: AA0063742022レポート番号: JAXA-SP-07-016

打上げ時の宇宙機には、ロケットの間のインターフェイスを通じて機械振動が加わる。さらに、広い周波数成分をもつ音圧が宇宙機表面に加わることにより振動する。太陽電池パドルやアンテナなど軽量で大きな面積をもつ構造や、比較的高い固有振動数をもつコンポーネントは、音響荷重に影響されやすい。この音響振動の数値予測手法に着目すると、既存の解析手法としては、低周波側では有限要素法(FEM)等の決定論的手法、高周波側では統計的エネルギー法(SEA)等の確率統計的手法が適用されている。しかし、一般に両手法では信頼性の高い解析のできない中間周波数領域が存在することが知れている。そこで、本論文では、まず、低周波領域における解析に留まっているFEM の性質について議論するため、FEM の基本的なモデング手法や解析手法等について整理し、単純なモデルのランダム応答解析を実施した。さらに、補強材を含む単純な衛星主構造モデルを用いて、FEM による衛星の音響振動解析を実施した。これらの解析結果から、数値分散誤差の影響や音響と構造の連成解析の重要性について明らかにした。次に、このような要素ベース手法の欠点を克服するために新たに提案された波動ベース法(WBM)について議論した。まず、本研究において開発した2 次元WBM 解析コードと商用FEM ソフトウェアを用いて、いくつかの単純な音響構造連成解析と非連成音響解析を行い、WBM コードの検証を行うとともに基本的なWBM の性質を把握した。そして、この結果を踏まえて、フェアリング内の衛星の音響振動連成解析を実施した。以上の結果から、WBM の音響振動連成解析における高精度な決定論的性質が確かめられた。つまり、支配方程式を厳密に満たす波動関数の性質により数値分散誤差を伴わないことから、既存の解析手法で精度の高い解析ができない中間周波数領域を含む広い周波数領域における解析が可能であることが分かった。さらに、空間離散化(メッシュ)を必要としないため、モデル化が非常に容易で計算負荷の小さい実用的な手法であることが分かった。

ヘリコプタ全機周りの流れ場を解析する3次元非定常オイラー・コードをベースに,ブレード/渦干渉(Blade-Vortex Interaction:BVI)騒音低減用のアクティブ・フラップ制御(Active Flap Control:AFC)を解析するコードを開発し,Ffowcs Williams and Hawkings(FW-H)の式に基づく音響解析コードと組み合わせることで,AFCの重要な2つのパラメータであるフラップの位相角と取り付け位置が騒音に及ぼす影響を解析した。ここで,位相角の影響を容易に把握するための簡易的なモデルを提案し,そこから得られる知見を適用することで,実験値の予測を格段に改善できることを示した。また,ここで用いた特定の条件では,位相角60°のとき5.62dBの騒音低減効果が得られた。フラップ位置の影響については,より翼端に近い位置に配置されたフラップを適正な位相角で作動させることが騒音低減に有効であることを示した。また,フラップを翼端に近い位置に配置した場合,フラップの外端から発生するフラップ渦がブレードの翼端渦と融合する現象が捉えられた上,それが翼端渦の強度に影響を及ぼしてBVIの起こる瞬間にブレード上圧力変動を増加させることを確認した。さらに,フラップ位置を内側から外側に移動させると,ある位置を境に騒音低減のメカニズムが劇的に変わることが分かったが,そのメカニズム解明については今後の課題である。

ロータの回転や6自由度の機体運動等の複雑な動きのため、メイン/テールロータ及び胴体からなるヘリコプタの数値シミュレーションにおいて、正確で速い補間アルゴリズムの重要性が高まってきている。本報告では、直交格子と曲線格子で構成される移動重合格子を利用したCFD コードに対して、より正確で早い補間法を提案した。新しい補間法では、(1)直交格子の特性、(2)ヘリコプタ・ブレードの特殊な幾何学的配置、及び(3)並列計算時の計算負荷バランスなどを十分有効に利用できるアルゴリズムが考案されている。第一章では、Alternating Index Searching (AIS) アルゴリズムを提案し、従来のLinear Searching アルゴリズムに対し、2 次元の簡単なケースと実際のヘリコプタを模擬した3 次元計算のケースで補間計算の速度を比べた。第二章では、並列計算における各計算ノードの負荷バランスを考慮したReverse Index Searching (RIS) アルゴリズムを提案した。この補間法を利用することによって、ヘリコプタのより効果的な大型計算を実現することができた。

In this study, numerical calculations of sound propagation for a simply modeled rocket plume duct with different jet configurations of H-IIA202 and 204 rockets have been conducted. The Euler flow calculation option of a hybrid Euler/LEE (Linearized Euler Equation) code is used. The results of these calculations indicate that the sound waves from the modeled rocket plume duct have different frequency characteristics according to the jet configurations.

The limitations of the finite element method (FEM) mainly on numerical dispersion errors and on model dimension indicate that an alternative deterministic approach is necessary for the coupled vibro-acoustic analysis in the higher frequency range. Therefore, a novel approach called the wave base method (WBM) is discussed since it requires no meshes, and subsequently provides no dispersion errors and small model dimension. A 2-dimensional WBM code is then developed, and some examples are solved to show some advantageous features of the WBM. Moreover, a simple vibro-acoustic simulation on spacecraft is demonstrated, and local responses can be obtained due to its deterministic characteristic. It is also shown that the WBM has high potential for the vibro-acoustic analysis with the wide frequency range.

The importance of accurate and fast interpolation algorithm is growing up for helicopter simulation with multi-body configuration, because of complex movement of helicopter including rotor-rotation and flight motion. In this paper, interpolation algorithms are implemented for two different grids, Cartesian grid and curvilinear grid, of which the overlapped grid system consists. New searching algorithms are proposed to make full use of (1) the characteristics of Cartesian grid, (2) special geometric configuration of helicopter, and (3) load balancing in parallel computation. By applying these searching algorithms, efficient massive computation can be achieved for the helicopter configuration.

Blade Vortex Interaction (BVI) noise of helicopter is predicted by using CFD analysis based on three dimensional Euler equations and subsequent acoustic analysis. In the CFD analysis, Beddose generalized wake model is used to prescribe the geometrical position of tip vortices in the wake, and Lamb-Oseen model is used to consider the vortex dissipation in order to avoid overestimation of BVI noise which sometimes occurs when the dissipation is not considered in the model. Lamb-Oseen model simulates the vortex dissipation by enlarging the core radius and attenuating the swirl velocity of vortex. The calculated sound pressure shows adequate decrease of BVI noise peaks as well as the reduction of small fluctuations compared to the no dissipation result. In consequence, the use of vortex dissipation model like Lamb-Oseen model is effective to improve the quality of calculation by avoiding the overestimation of BVI noise and by reducing the unrealistic fluctuations in the sound pressure calculation.

ロータの回転や6自由度の機体運動等の複雑な動きのため、メイン/テールロータ及び胴体からなるヘリコプタの数値シミュレーションにおいて、正確で速い補間アルゴリズムの重要性が高まってきている。本報告では、直交格子と曲線格子で構成される移動重合格子を利用したCFD コードに対して、より正確で早い補間法を提案した。新しい補間法では、(1)直交格子の特性、(2)ヘリコプタ・ブレードの特殊な幾何学的配置、及び(3)並列計算時の計算負荷バランスなどを十分有効に利用できるアルゴリズムが考案されている。第一章では、Alternating Index Searching (AIS) アルゴリズムを提案し、従来のLinear Searching アルゴリズムに対し、2 次元の簡単なケースと実際のヘリコプタを模擬した3 次元計算のケースで補間計算の速度を比べた。第二章では、並列計算における各計算ノードの負荷バランスを考慮したReverse Index Searching (RIS) アルゴリズムを提案した。この補間法を利用することによって、ヘリコプタのより効果的な大型計算を実現することができた。

In this study, numerical calculations of sound propagation for a simply modeled rocket plume duct with different jet configurations of H-IIA202 and 204 rockets have been conducted. The Euler flow calculation option of a hybrid Euler/LEE (Linearized Euler Equation) code is used. The results of these calculations indicate that the sound waves from the modeled rocket plume duct have different frequency characteristics according to the jet configurations.

The limitations of the finite element method (FEM) mainly on numerical dispersion errors and on model dimension indicate that an alternative deterministic approach is necessary for the coupled vibro-acoustic analysis in the higher frequency range. Therefore, a novel approach called the wave base method (WBM) is discussed since it requires no meshes, and subsequently provides no dispersion errors and small model dimension. A 2-dimensional WBM code is then developed, and some examples are solved to show some advantageous features of the WBM. Moreover, a simple vibro-acoustic simulation on spacecraft is demonstrated, and local responses can be obtained due to its deterministic characteristic. It is also shown that the WBM has high potential for the vibro-acoustic analysis with the wide frequency range.

The importance of accurate and fast interpolation algorithm is growing up for helicopter simulation with multi-body configuration, because of complex movement of helicopter including rotor-rotation and flight motion. In this paper, interpolation algorithms are implemented for two different grids, Cartesian grid and curvilinear grid, of which the overlapped grid system consists. New searching algorithms are proposed to make full use of (1) the characteristics of Cartesian grid, (2) special geometric configuration of helicopter, and (3) load balancing in parallel computation. By applying these searching algorithms, efficient massive computation can be achieved for the helicopter configuration.

Blade Vortex Interaction (BVI) noise of helicopter is predicted by using CFD analysis based on three dimensional Euler equations and subsequent acoustic analysis. In the CFD analysis, Beddose generalized wake model is used to prescribe the geometrical position of tip vortices in the wake, and Lamb-Oseen model is used to consider the vortex dissipation in order to avoid overestimation of BVI noise which sometimes occurs when the dissipation is not considered in the model. Lamb-Oseen model simulates the vortex dissipation by enlarging the core radius and attenuating the swirl velocity of vortex. The calculated sound pressure shows adequate decrease of BVI noise peaks as well as the reduction of small fluctuations compared to the no dissipation result. In consequence, the use of vortex dissipation model like Lamb-Oseen model is effective to improve the quality of calculation by avoiding the overestimation of BVI noise and by reducing the unrealistic fluctuations in the sound pressure calculation.

A study to establish a numerical prediction method for vibro-acoustics during rocket launch has been conducted. This method consists of some analysis elements: numerical analysis of sound generation, propagation, transmission, and vibro-acoustics of payload. In this paper, sound pressure distributions obtained by these analyses are visualized. At first, flow and acoustic fields of modeled H-IIA launch pad are shown, which is obtained by a hybrid methodology of the Euler and Linearized Euler Equations (LEE) solvers. Then, an arch-shaped transmission wave is observed from a sound transmission analysis of hollow wall model using the Finite Difference Time Domain (FDTD) method. Finally, it is clearly shown that the acoustic field is affected by structural vibrations and the coupled vibro-acoustic problem of flexible satellite model can be solved by the Wave Based Method (WBM).