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Boiling Water Reactors (BWR) and Advanced Pressurized Water Reactors (APWR) often use spargers to release steam from the primary coolant system into a pool with subcooled water to prevent containment overpressure. Direct contact Condensation (DCC) creates sources of mass, heat, and momentum determined by the condensation regimes in a subcooled pool. Thermal stratification can develop in the pool if buoyancy forces created by the heat source dominate the momentum source. Only part of the stratified pool volume can be used as the heat sink which is a safety concern. Modeling of steam injection and its effect on a large pool is computationally expensive due to the considerable spatial and temporal scale differences between the steam-water interface dynamics and global pool circulation. To enable the prediction of realistic plant transients, Effective Heat Source (EHS) and Effective Momentum Source (EMS) models were proposed. These models aim to calculate the timeaveraged integral effects of the steam injection on the pool without resolving the dynamics of the interface and DCC phenomena. This work aims to develop the empirical correlations that predict the time-averaged effective momentum induced by steam injection into a subcooled pool. The experimental data were collected in a Separate Effect Facility (SEF-POOL) at LUT, Finland. The force acting on the injection pipe was measured in SEF-POOL to estimate the effective momentum rate created by steam injection. The condensation regime coefficient C, defined as the ratio of effective momentum rate to the theoretical momentum rate of injected steam, is presented as a function of Jakob and Mach numbers. We found that the pitch-to-diameter ratios (P/D) had a significant effect on both the effective (time averaged) momentum rate and instantaneous forces measured in the low Jakob number region, which might be attributed to the interactions between neighboring nozzles., QC 20240813

Equations of motion for compressible point vortices in the plane are obtained in the limit of small Mach number, M, using a Rayleigh-Jansen expansion and the method of Matched Asymptotic Expansions. The solution in the region between vortices is matched to solutions around each vortex core. The motion of the vortices is modified over long time scales [Formula: see text] and [Formula: see text]. Examples are given for co-rotating and co-propagating vortex pairs. The former show a correction to the rotation rate and, in general, to the centre and radius of rotation, while the latter recover the known result that the steady propagation velocity is unchanged. For unsteady configurations, the vortex solution matches to a far field in which acoustic waves are radiated. This article is part of the theme issue 'Mathematical problems in physical fluid dynamics (part 2)'.

Equations of motion for compressible point vortices in the plane are obtained in the limit of small Mach number, M, using a Rayleigh-Jansen expansion and the method of Matched Asymptotic Expansions. The solution in the region between vortices is matched to solutions around each vortex core. The motion of the vortices is modified over long time scales [Formula: see text] and [Formula: see text]. Examples are given for co-rotating and co-propagating vortex pairs. The former show a correction to the rotation rate and, in general, to the centre and radius of rotation, while the latter recover the known result that the steady propagation velocity is unchanged. For unsteady configurations, the vortex solution matches to a far field in which acoustic waves are radiated. This article is part of the theme issue 'Mathematical problems in physical fluid dynamics (part 2)'.

The effect of high-level excitation on the acoustic properties of perforates, and formulation of the non-linear part of the impedance is under scientific discussion. Analytical models including the non-linear properties, as well as various experimental studies give varying results for the acoustic impedance. This paper aims to provide detailed results obtained under high-level excitation with different acoustic wave incidence and flow configurations. Contrary to the well-established two-port configuration, here, a three-port measurement technique is used to observe the acoustic impedance of the perforated plate using excitation from the three different directions. Plane wave propagation is considered and physical quantities such as in-hole particle velocity, which is calculated at the perforate sample, are used as the controlling parameters. This paper is an attempt to study the effect of high-level excitation on the acoustic behavior of the perforates with grazing and normal acoustic incidence. Moreover, the non-linear behavior of the perforate determined in presence of an external grazing flow is also discussed., QC 20221005

Pumping in vacuum chambers is part of the field of environmental electron microscopy. These chambers are separated from each other by a small-diameter aperture that creates a critical flow in the supersonic flow regime. The distribution of pressure and shock waves in the path of the primary electron beam passing through the differentially pumped chamber has a large influence on the quality of the resulting microscope image. As part of this research, an experimental chamber was constructed to map supersonic flow at low pressures. The shape of this chamber was designed using mathematical–physical analyses, which served not only as a basis for the design of its geometry, but especially for the correct choice of absolute and differential pressure sensors with respect to the cryogenic temperature generated in the supersonic flow. The mathematical and physical analyses presented here map the nature of the supersonic flow with large gradients of state variables at low pressures at the continuum mechanics boundary near the region of free molecule motion in which the Environmental Electron Microscope and its differentially pumped chamber operate, which has a significant impact on the resulting sharpness of the final image obtained by the microscope. The results of this work map the flow in and behind the Laval nozzle in the experimental chamber and are the initial basis that enabled the optimization of the design of the chamber based on Prandtl’s theory for the possibility of fitting it with pressure probes in such a way that they can map the flow in and behind the Laval nozzle.

Pumping in vacuum chambers is part of the field of environmental electron microscopy. These chambers are separated from each other by a small-diameter aperture that creates a critical flow in the supersonic flow regime. The distribution of pressure and shock waves in the path of the primary electron beam passing through the differentially pumped chamber has a large influence on the quality of the resulting microscope image. As part of this research, an experimental chamber was constructed to map supersonic flow at low pressures. The shape of this chamber was designed using mathematical–physical analyses, which served not only as a basis for the design of its geometry, but especially for the correct choice of absolute and differential pressure sensors with respect to the cryogenic temperature generated in the supersonic flow. The mathematical and physical analyses presented here map the nature of the supersonic flow with large gradients of state variables at low pressures at the continuum mechanics boundary near the region of free molecule motion in which the Environmental Electron Microscope and its differentially pumped chamber operate, which has a significant impact on the resulting sharpness of the final image obtained by the microscope. The results of this work map the flow in and behind the Laval nozzle in the experimental chamber and are the initial basis that enabled the optimization of the design of the chamber based on Prandtl’s theory for the possibility of fitting it with pressure probes in such a way that they can map the flow in and behind the Laval nozzle.

Multiphase, compressible and viscous flows are of crucial importance in a wide range of scientific and engineering problems. Despite the large effort paid in the last decades to develop accurate and efficient numerical techniques to address this kind of problems, current models need to be further improved to address realistic applications. In this context, we propose a numerical approach to the simulation of multiphase, viscous flows where a compressible and an incompressible phase interact in the low-Mach number regime. In this frame, acoustics are neglected but large density variations of the compressible phase can be accounted for as well as heat transfer, convection and diffusion processes. The problem is addressed in a fully Eulerian framework exploiting a low-Mach number asymptotic expansion of the Navier-Stokes equations. A Volume of Fluid approach (VOF) is used to capture the liquid-gas interface, built on top of a massive parallel solver, second order accurate both in time and space. The second-order-pressure term is treated implicitly and the resulting pressure equation is solved with the eigenexpansion method employing a robust and novel formulation. We provide a detailed and complete description of the theoretical approach together with information about the numerical technique and implementation details. Results of benchmarking tests are provided for five different test cases., QC 20210409

Pumping in vacuum chambers is part of the field of environmental electron microscopy. These chambers are separated from each other by a small-diameter aperture that creates a critical flow in the supersonic flow regime. The distribution of pressure and shock waves in the path of the primary electron beam passing through the differentially pumped chamber has a large influence on the quality of the resulting microscope image. As part of this research, an experimental chamber was constructed to map supersonic flow at low pressures. The shape of this chamber was designed using mathematical–physical analyses, which served not only as a basis for the design of its geometry, but especially for the correct choice of absolute and differential pressure sensors with respect to the cryogenic temperature generated in the supersonic flow. The mathematical and physical analyses presented here map the nature of the supersonic flow with large gradients of state variables at low pressures at the continuum mechanics boundary near the region of free molecule motion in which the Environmental Electron Microscope and its differentially pumped chamber operate, which has a significant impact on the resulting sharpness of the final image obtained by the microscope. The results of this work map the flow in and behind the Laval nozzle in the experimental chamber and are the initial basis that enabled the optimization of the design of the chamber based on Prandtl’s theory for the possibility of fitting it with pressure probes in such a way that they can map the flow in and behind the Laval nozzle.

Pumping in vacuum chambers is part of the field of environmental electron microscopy. These chambers are separated from each other by a small-diameter aperture that creates a critical flow in the supersonic flow regime. The distribution of pressure and shock waves in the path of the primary electron beam passing through the differentially pumped chamber has a large influence on the quality of the resulting microscope image. As part of this research, an experimental chamber was constructed to map supersonic flow at low pressures. The shape of this chamber was designed using mathematical–physical analyses, which served not only as a basis for the design of its geometry, but especially for the correct choice of absolute and differential pressure sensors with respect to the cryogenic temperature generated in the supersonic flow. The mathematical and physical analyses presented here map the nature of the supersonic flow with large gradients of state variables at low pressures at the continuum mechanics boundary near the region of free molecule motion in which the Environmental Electron Microscope and its differentially pumped chamber operate, which has a significant impact on the resulting sharpness of the final image obtained by the microscope. The results of this work map the flow in and behind the Laval nozzle in the experimental chamber and are the initial basis that enabled the optimization of the design of the chamber based on Prandtl’s theory for the possibility of fitting it with pressure probes in such a way that they can map the flow in and behind the Laval nozzle.

In the first half of this paper, we present a fresh perspective toward the Wind Triangle Problem in aerial navigation by deriving necessary and sufficient conditions, which we call "go/no-go conditions", for the existence/non-existence of a solution of the problem. Although our derivation is based on simple trigonometry and basic properties of quadratic functions, it is mathematically rigorous. We also offer examples to demonstrate how easy it is to check these conditions graphically. In the second half of this paper, we use function theory to re-examine another problem in aerial navigation, namely, that of computing true airspeed — even in supersonic flight — from only three instrument readings obtainable from a basic flight instrument panel: calibrated airspeed, pressure altitude, and total air temperature. We present the first known mathematically rigorous analysis of the use of fixed-point iteration to compute true airspeed and Mach number from the Rayleigh Supersonic Pitot Equation. Our analysis comes with error estimates that are not found anywhere in the literature.

The present study focuses on the numerical investigation of the Mach 1.86 supersonic jet from the 2-D convergent-divergent nozzle for NPRs 4 to 6 with a step size of one, covering the overexpanded and the near correct expansion conditions. The computational model is constructed into design modular of ANSYS 16.0 workbench. The structured grids are generated into the ICEM module of the ANSYS workbench and the fluid flow analysis, on the modelled part is done by FLUENT solver. The extent of jet mixing is identified by plotting non-dimensional total pressure, along the jet centerline, with respect to the non-dimensional distance downstream of the nozzle exit and along the jet centerline. Also the Mach contours are plotted for different NPRs indicating, the number of shock cells and the variation of the shock strength with respect the change in the NPR due to the prevailing adverse pressure gradient at the nozzle exit.

[EN] In this work, the effect of high injection pressures and ambient gas properties on diesel spray penetration and spreading angle are studied. To this end a multi-hole piezoelectric injector was used, and MIE-scattering optical technique to visualize the spray. Injection pressures up to 270 MPa were used throughout the experiments. Additionally, the spray behavior going from subsonic to supersonic state was analyzed by controlling the ambient gas speed of sound, promoting in this way supersonic jets. For this purpose, measurements were done using three different ambient gases (SF6, CO2, and N2) at isothermal conditions. The results showed that sprays near transonic or in supersonic state had a higher penetration rate than those in subsonic state. Furthermore, among the sprays near transonic or in supersonic state, those with higher Mach number had faster penetration. Differently, within the sprays at subsonic state, no significant variations in spray penetration rate were found, regardless of the difference in the Mach number. Shock waves appearances were pointed out as a possible explanation for the spray penetration variations observed. Finally, a statistical analysis is presented for the spray penetration under isothermal conditions and for each ambient gas.

Aeroacoustics has received great attention in the past decade, owing to the ever stricter noise regulations. Despite the stochastic nature of most aeroacoustic systems, non-deterministic investigations in regards to computational aeroacoustics (CAA) are limited. In this paper, uncertainty quantification has been achieved for the noise propagation stage of hybrid CAA, and also on the noise prediction of a non-lifting helicopter rotor in hover. Analytical and computational fluid dynamics test cases have been analyzed, with uncertainties propagated through these systems using non-intrusive polynomial chaos methods. It is shown here that the source of the uncertainty in the noise is dominated by the major characteristic properties of the simulations, such as the mean flow Mach number and blade tip Mach number. Only at a low tip Mach number uncertainties in the blade thickness may contribute significantly to the noise uncertainty. Apart from this, it is seen to be unlikely that small uncertainties in the geometry, ambient conditions and observer properties will contribute significantly to the noise uncertainty. A peak pressure uncertainty of up to 20% is seen in the hovering helicopter test case, from small, realistic uncertainties. This highlights the importance of considering uncertainties in CAA investigations. © 2018 Elsevier Ltd

The generation mechanism and possible counter measures for fluid driven whistles in low Mach number flow duct networks are discussed. The vortex sound model, where unstable shear layers interact with the acoustic field and act as amplifiers under certain boundary conditions, is shown to capture the physics well. Further, for the system to actually whistle an acoustic feedback to the amplifying shear layer is also needed. The demonstration example in this study is a generalized resonator configuration with annular volumes attached to a straight flow duct via a number of small holes, perforations, around the duct's circumference. At each hole a shear layer is formed and the acoustic reflections from the resonator volumes and the up and downstream sides provides a possible feedback to them. Not only the Helmholtz mode but also ring modes in the annular volumes provide a feedback to sustain whistles. The attenuation properties as well as the whistling frequencies at varying inlet mean flow velocities for this system are studied both numerically and experimentally showing that good quality predictive simulations are possible using the vortex sound theory. Finally a few countermeasures against whistling are tested. Both the feedback and the shear layers are manipulated. Best effect was found disturbing the shear layers by covering the holes with a coarse mesh., QC 20220620

The present study focuses on the numerical investigation of the Mach 1.86 supersonic jet from the 2-D convergent-divergent nozzle for NPRs 4 to 6 with a step size of one, covering the overexpanded and the near correct expansion conditions. The computational model is constructed into design modular of ANSYS 16.0 workbench. The structured grids are generated into the ICEM module of the ANSYS workbench and the fluid flow analysis, on the modelled part is done by FLUENT solver. The extent of jet mixing is identified by plotting non-dimensional total pressure, along the jet centerline, with respect to the non-dimensional distance downstream of the nozzle exit and along the jet centerline. Also the Mach contours are plotted for different NPRs indicating, the number of shock cells and the variation of the shock strength with respect the change in the NPR due to the prevailing adverse pressure gradient at the nozzle exit.

The flow-generated sound from an air inlet at a low Mach number is studied using detached eddy simulation and direct numerical beamforming. The flow is solved through RANS and compressible detached eddy simulation for various velocity ratios, defined as the ratio between the flow speed at the duct entrance and in the free stream. Results for the flow show very good agreement with experimental data from the National Advisory Committee for Aeronautics (NACA). Pressure data is extracted at the position of a virtual microphone array inside the simulation domain, outside of the unsteady flow region. The array data is beamformed on a source test grid over the inlet using standard beamforming and Linear Programming deconvolution approaches. As beamforming is performed from simulated data directly in the acoustically resolved domain, the method is referred to as Direct Numerical Beamforming (DNB). The results give information about the location, amplitude and frequency content of the sound sources around the inlet. The resulting incoherent sound sources can be used as a model for the sound radiated from the inlet. Using a simple Green's function, the corresponding far-field spectrum is computed for two velocity ratios, showing a very good agreement with the spectra obtained from a Ffowcs Williams and Hawkings integral., QC 20161125

A multi-port method is used to extract complex mode amplitudes from experimental data of in-duct orifice plates for frequencies with cut-ons of several higher order modes. The two-port method is therefore extended to resolve higher order modes, i.e. spinning modes and radial modes. The work presented in Sack et all.1 is continued and a new test- rig design is presented which enables more stable measurements and cleaner data. The scattering and source data for higher order modes of a circular sharp-edged orifice plate with low Mach-number mean flow is presented and used for advanced post-processing. The gained data is used to predict the scattering of a tandem orifice configuration and the prediction is compared to measurement results., QC 20161125

Heat exchanger tube bundles, consisting of tube arrays in cross flows, are vital in the efficient working of power generation systems. If sound propagates through these bundles, it can lead to resonance or acoustic attenuation, and thereby affecting the working of the power generation unit. Therefore, it is important to study the aeroacoustics of tube rows. The aim of the present work is to experimentally validate the quasi-steady compressible model developed to study the aeroa-coustic response of an array of tubes with bias flow. In order to accomplish this, the array of tubes is approximated by a geometry consisting of two half cylinders separated by a gap and having a bias flow through the gap. Firstly, the case with no flow is considered and the experimental results for the reflection and transmission coefficients are compared against the analytical expressions developed by Huang and Heckl (Huang and Heckl, 1993, Acustica 78, 191-200). Then the cases with flow are considered. A quasi-steady subsonic compressible model is developed to predict the reflection and transmission coefficients, valid for low Strouhal numbers, with the additional assumption of small Helmholtz number and low Mach numbers. This model is validated against the experimental results for the transmission and reflection coefficients. A two-port multi-microphone measurement technique is used to obtain the pressure data and a subsequent wave decomposition is utilised to extract the transmission and reflection coefficients. The results show good agreement with theory for various Mach numbers, in the low Strouhal number regime., Funding Details: FP7-PEOPLE-ITN-2012, EC, European CommissionQC 20161125

More efficient and physically smaller axial turbine designs are promoted to lower emissions and increase revenue. The physical size and the weight of the axial turbine can be minimised by adjusting the distance between successive stator and rotor rows. The influence of changing stator-rotor axial clearance can usually have either a positive or a negative influence on the turbine performance, and the reasons for this varying behaviour are not currently fully understood. A previous study revealed several design parameters that correlate with the efficiency curve shape. However, the effects of two parameters, namely the stator outlet Mach number and the reduced blade passing frequency, still remained unclear. Therefore, a novel approach is taken to analyse the correlations between the two design parameters and the axial clearance efficiency curve shape. Several different turbines are analysed using data available in the literature, and also new data are presented. The study suggests that the stator outlet Mach number correlates reasonably well with the efficiency curve shape, and it was further linked to five loss mechanisms and the rotational speed. Although the unsteady interaction plays an important role in the loss share, the level of unsteadiness did not correlate with the efficiency curve shape. © 2015 Institution of Mechanical Engineers.

The thesis presented here details the design, construction and initial testing of rig for use with high temperature film cooling testing. The film cooling rig was supplied with hot mainstream gas from a well-stirred reactor operating on a propane/air mixture capable of multiple equivalence ratios. The Hastelloy test plates contained an internal cooling channel to allow for overall effectiveness measurements. Thermocouples on both the freestream and internal surfaces of the test plate provided temperature differences for heat flux calculations. The test plates had a quarter circle leading edge with a tapered trailing edge to provide surface curvature for the film cooling studies. The height of the test channel could be adjusted for multiple Mach numbers to be set for the mainstream flow. Using a plate containing five rows of holes in trenches and two rows of showerhead holes, methodology was developed for collecting and analyzing the necessary data to obtain net heat flux reduction and overall effectiveness results. This methodology was then applied to the same plate to gather comparative results for reacting versus non-reacting film cooling., The original document contains color images.

Motivated by the recent development in fabricating graphene/carbon nanotube-based surface compliant loud speakers, the effectiveness of thermoacoustic actuators that locally introduce high-intensity acoustic waves for active flow control is examined by performing high-fidelity large eddy simulation for compressible flow over a wall-mounted hump at a Reynolds number of 0:5 x 1000000 and Mach number of 0.25. Based on performance characteristics of the grephene-based thermoacoustic actuators, high-frequency actuation around Helmholtz number of 3.0 is considered. We observe that the actuation is able to introduce small-scale perturbations to the shear layer in the separated ow and attenuate the formation of large scale spanwise vortices. This ow control technique elongates the recirculation zone and shifts the low-pressure region downstream of the hump. As a consequence, the drag on the hump is reduced by approximately 4.31% and 6.33% for two and three-dimensional simulation, respectively., Presented at the AIAA Aerospace Sciences Meeting (53rd) held in Kissimmee, FL on 5-9 Jan 2015.

We derive equations for the mean entropy and the mean internal energy in low-Mach-number temperature stratified turbulence (i.e. for turbulent convection or stably stratified turbulence), and show that turbulent flux of entropy is given by Fs=σus, where σ is the mean fluid density, s is fluctuation of entropy and overbars denote averaging over an ensemble of turbulent velocity fields, u. We demonstrate that the turbulent flux of entropy is different from the turbulent convective flux, Fc=T σ us , of the fluid internal energy, where T is the mean fluid temperature. This turbulent convective flux is well-known in the astrophysical and geophysical literature, and it cannot be used as a turbulent flux in the equation for the mean entropy. This result is exact for low-Mach-number temperature stratified turbulence and is independent of the model used. We also derive equations for the velocity-entropy correlation, us, in the limits of small and large Péclet numbers, using the quasi-linear approach and the spectral τ approximation, respectively. This study is important in view of different applications to astrophysical and geophysical temperature stratified turbulence., QC 20160212

The main component for the reduction of the exhaust noise from internal combustion engine (IC-engine) is the properly designed silencer. In addition, the acoustic performance of the after treatment devices (ATD) has gained interest as noise reduction components. Typically an ATD for diesel engines is assembled from components such as selective catalytic reducers (SCR), diesel oxidation catalysts (DOC) and diesel particulate filters (DPF). One new alternative to the conventional DPF is the particle oxidation catalyst (POCR ). The POC R substrate studied in this paper is of type POC-X, which consists of fine, corrugated metallic wire mesh screens piled askew and rolled into a cylindrical shape. According to earlier studies the acoustic behaviour of the filter with negligible flow can be estimated with reasonable accuracy starting from the Kirchhoff solution. It was also concluded earlier that treating the filter as an acoustic resistance gives good results in the plane wave frequency range with moderate Mach-number flow. In this paper, the experiments on the acoustic performance of the POC-X in the non-plane wave frequency range with low Mach-number flow (M = 0.01) are introduced. Starting from the earlier findings, procedure to obtain an estimate for the acoustic performance of the filter in the entire audio frequency range is proposed. Using the computationally effective semi-empirical model presented in this paper is especially reasonable when optimizing the complete exhaust system of an IC-engine for the maximum noise reduction., QC 20160304

Using CFD the computational investigation of two axisymmetric and axiradial cones (diffusers) typical for the exhaust branch pipes of steam turbines was carried out in the wide range of the pressure variation at the entrance of the cones with available tip-clearance leakage. As the pressure decreased the Mach number at the cone entrance increased. It has been established that in subsonic modes the cone with continuous external by-pass without the circular blade had the flow without separation and considerably lower values of loss coefficients in comparison with the diffuser that has a projection of the internal by-pass and the circular blade. At trans-sonic rates the coefficients of total losses of both studied cones are increased reaching the values that exceed 1 and have similar values; however the coefficients of internal losses and exit velocity losses differ significantly. The obtained results allow us to extend somewhat the database for the designers that are involved in the solution of problems related to the design of outlet branch pipes for flow rates approaching the sound velocity.

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.

Navier-Stokes simulations are carried out for the investigation of a body-flap effect on the subsonic to supersonic base-entry aerodynamic characteristics for the conical SSTO-rocket configuration. It is clarified that the body-flap effect has strong dependence on the angle of attack and Mach number. The body-flap effect is small at subsonic and transonic conditions. At the subsonic and transonic conditions, the windside pressure level over the body-flap is not so high as compared to the leeside due to the influence of the flow which is accelerated over the base surface. It is also clarified that the pressure level behind the body-flap does not become lower than that in the recirculating flow behind the body in all the subsonic to supersonic conditions.

For the anti-symmetric flutter speeds of a tip-fin-shaped reentry space vehicle model, numerical predictions using Doublet-Point Method (DPM) have been different from experimental results in a part of lower supersonic region. The cause of the discrepancy is suspected to be that the local flow over the wing differs from the free stream. Instead of free stream Mach numbers, measured local Mach numbers are adopted for flutter boundary analyses, and the corrected numerical results show good agreement with experimental ones. Therefore, it is made clear that the correction by local Mach numbers improves the accuracy of the flutter boundary., JAXA Special Publication, 宇宙航空研究開発機構特別資料

TSTOを2020年までに実現する計画がJAXAにて検討されている。その機体構成は、1段が空気吸込み式エンジン、2段が再使用ロケットエンジンの構成としたものである。一方、ATREXエンジンの改良を続け、エンジンによる基盤技術研究として、過去10年以上に渡り、能代多目的実験場にて試験データを蓄積してきた。ATREXエンジンのメリットとしては、高推力、簡易エンジン構造、低圧損なコアエンジン部などがあるが、一方デメリットとして低比推力および熱交換器用耐熱材の課題がある。特に、ATREXエンジンの高性能を引き出すためには、1,200K以上で作動可能な熱交換器として、加工性、耐熱性などの課題がクリア出来る材料の登場を待つ必要がある。これらの問題をカバーすべく、現状の技術で開発可能な、予冷ターボ式(PCTJ)エンジンを、TSTO用空気吸込み式エンジンの候補として選定検討した。その検討過程および試験計画などについて報告する。

To improve combustion efficiency and the inlet start condition, a scramjet model with a ramp combined with a single strut was tested under Mach 8 flight conditions at the Ramjet Engine Test Facility of the Ramjet Propulsion Research Center. The attached ramp shielded some of the fuel injectors. The fuel flow rate from the open injectors to the flow path was designated as the effective fuel flow rate. In the tests, a combustion efficiency of 90 percent was attained with vertical injection of hydrogen fuel. The thrust increase was 590 N at the effective equivalence ratio of 1.3. However, because the engine geometry was not optimized, a sufficient increase in thrust was not attained. The high temperature and high pressure necessary for ignition and combustion were achieved by the ramp. When the pressure in the isolator was 160 times that of the free-stream air due to combustion, the inlet was in the started condition despite the high pressure. This improved started condition was attained by using the ramp to increase the pressure in the inlet., ラムジェットエンジン試験設備(RJTF)におけるマッハ8条件での始動性および燃焼性能を改善するために、ストラットに加えてランプを取り付けたスクラムジェットエンジン模型の試験を行った。燃料を垂直噴射した時に、エンジン出口での燃焼効率は約90％であった。当量比1.3の条件時に、590Nの推力増加を測定した。しかしながら最適化されていないエンジン形状であったために、十分な推力増加は得られなかった。着火および燃焼に必要な高い温度および圧力が、ランプの装着によって達成された。燃焼器内圧が、エンジンに流入する気流圧力の160倍に達した場合にも、インレットは始動状態を保つことができた。これはランプを用いてインレット内圧力を上昇させた効果による。, Technical Report of National Aerospace Laboratory, 航空宇宙技術研究所報告

航空宇宙技術研究所1.27m極超音速風洞では、新しい試験技術の1つとして光学的可視化法による極超音速流場の速度分布測定を試みている。本実験ではマッハ数10の極超音速流における平板模型と楔模型からの衝撃波干渉領域の速度場を測定するため、炭酸ガスのドライアイス微粒子によるレーレ散乱光を1台のICCDカメラで観測した。得られたデータは衝撃波干渉領域内の微小な速度差を識別できることを示し、レーレ散乱を利用した計測方法が有効であることを確認した。, JAXA Special Publication, 宇宙航空研究開発機構特別資料

In order to validate real gas CFD (Computational Fluid Dynamics) code and construct appropriate real gas models, analysis of HEK (High Enthalpy shock tunnel at Kakuda Research Center), high enthalpy shock tunnel experiment is made by using one temperature chemically nonequilibrium CFD code. Aerothermodynamic heating distributions on the windward surface of HYFLEX (Hypersonic Flight Experiment) six percent model are compared in detail. Four test conditions in HEK are selected to investigate enthalpy and binary scaling effects on heating characteristics and good agreements are obtained. To extrapolate data from ground to flight, real gas computations are also made for HYFLEX flight conditions at catalytic and non-catalytic wall assumptions. On the other hands, at high altitude and high mach numbers like HOPE-X (H-2 Orbiting Plane Experimental) reentry flight, trim capability of high angle attack is an important design factor. To investigate real gas effects of these aerodynamic characteristics, calculations are made around space shuttle configuration along its trajectory from mach number five to 26. Real gas effects on center of pressure locations are compared well with the shuttle OADB data., 実在気体CFD(数値流体力学)コードを確認して適切な実在気体モデルを構築するために、1温度で化学的に非平衡のCFDコードを使ってHEK(角田宇宙推進技術研究センター高エンタルピー衝撃風洞)実験を解析した。HYFLEX(極超音速飛行実験)6%モデルの風上表面における空力加熱分布を詳細に比較した。HEKについて4テスト条件を選択して、加熱特性に対するエンタルピーと2成分スケーリングの効果を研究し、良い一致を得た。データを地上から飛行へ補外するために、触媒および非触媒壁条件におけるHYFLEX飛行条件について実在気体計算を行った。一方、HOPE-X(宇宙往還技術試験機)再突入飛行のような高高度高マッハ数において、高角度アタックのトリム能力が重要な設計指標となる。これらの空力特性に対する実在気体の効果を研究するために、マッハ数が5から26までのスペースシャトルの軌道に沿った配置周囲で計算を行った。圧力位置の中心に対する実在気体の効果をシャトルのOADBデータと良く対比した。

The high speed flight demonstration phase 2 was performed under cooperation of NAL, NASDA and CNES at Esrange in Sweden on July 1st, 2003. The purpose of this paper is to describe evaluation for the guidance and control system of the vehicle. The evaluation was executed by two approaches, a normal analysis and a special analysis. Although some cases of the Monte-Carlo simulation that is a part of the normal analysis could not satisfy mission requirements, detailed analysis of them says the system would lead the vehicle to successful flight experiment. The guidance and control system is graded by actual flight data, and it is cleared that the system satisfied all of the mission requirements. Based on the results of the preflight and the post-flight evaluation, this paper concludes that the design and evaluation for the guidance and control system of the vehicle were properly carried out.

Main characteristics, outlines of facilities, design methodologies and details of the equipments used are described for the NAL (National Aerospace Laboratory)'s new hypersonic wind tunnel system which is now under construction. This system, which is scheduled to be completed around the end of FY 1994, consists of the existing phi 50 cm tunnel and the new phi 1.27 m leg with a nozzle of a Mach number of 10. The new leg is being constructed parallel to the existing tunnel which has nozzles of Mach numbers of 5, 7, 9 and 11; not only the tunnel scale, but also daily operational cycles become approximately 2.5 times those of the existing tunnel. The operational conditions of the new leg are as follows: stagnation conditions of To = 600-1,200 K and Po = 1-10 MPa, Mach number = 10, Re = 3 x 10(exp 5)-4.5 x 10(exp 6)/m and a viscous parameter of bar V = M/square root of Re = 0.005-0.014. The operational conditions of the NAL's hypersonic wind tunnel system could cover the flight conditions of a winged space vehicle such as the Space Shuttle and the Japanese future unmanned winged reentry vehicle HOPE in M-Re and M-M/square root of Re maps in its Mach number range of M = 5-11., 1991年度から4年計画で建設が進められている当研究所の極超音速風洞システムについて、その基本計画、基本設計および基本特性に関する概要ならびに風洞の本体、諸設備、制御･計測装置、建屋･電源およびシステム運用に関する詳細について述べる。極超音速風洞の大型化整備としては、既設極超音速風洞に平行にマッハ数10,ノズル出口直径1.27mの大型極超音速風洞を併設し、マッハ数範囲の広い(5-11)既設風洞を極超音速空力研究目的に、大型風洞を開発用詳細データ取得目的に使い分けるものとする。また近年の極超音速風洞利用の需要増大に鑑み、風洞データ生産性向上のため、通風能力を既設風洞の2.5倍を目途とした諸設備の増強を行った。これら両風洞と容量･機能増強を図った諸設備の全体を、極超音速風洞システムとする。これにより当研究所は、米国空軍所属の世界最大の空気作動型極超音速風洞と同等のスケールの極超音速風洞を保持することとなる。本報告は、極超音速風洞大型化構想とそれに基づいた基本設計、基本特性ならびにそれを具体化した風洞システム･運用について述べることにより、諸制約の下での極超音速風洞の設計の手順と、風洞各部の詳細構造を明らかにするものである。

Reflection and refraction of linear disturbance waves by an oblique shock wave is studied by a linear analysis. Several different cases are considered, when the incident plane wave is a fast acoustic, slow acoustic, entropy, or vorticity wave. Results show that 1) a critical angle of the wave incidence exists, beyond which the regular wave solution of the linear problem can not be realized, 2) strongest sound waves are generated behind the shock wave, if the incident wave strikes the shock wave at the critical angle, and 3) at large Mach numbers of the upstream flow, the sound amplification factor asymptotically increases as the Mach number squared, while at the critical incidence, it is increased as the Mach number cubed., JAXA Special Publication, 宇宙航空研究開発機構特別資料

In this study, the flowfield around future TSTO Launch Vehicle was investigated by utilizing an electric discharge method. The experiments were carried out under the condition that the angles of attack of the model are -10, 0, and 10 deg. Shock shapes around a model were visualized. These results confirmed that discontinuous planes of density change existed in the shock-layer. The structures of shock waves and discontinuous planes of density change were shown in this paper. When angles of attack of the model became large, there was a tendency for the thickness of each layer to become thin.

Taking the opportunity of the development of STOL (Short Takeoff and Landing) experimental aircraft 'Asuka', the thrust calibration tank for a TPS (Turbine Powered Simulator) was built in the transonic wind tunnel at National Aerospace Laboratory. In this tank, the thrust and fan flow are measured with great accuracy by using a balance and a sonic venturi meter, respectively. Also, the tank itself is made of an airtight container in order to get the pressure ratio equivalent to the Mach number at a wind tunnel test. This paper describes the outline, functions and characteristics of the tank, and also the thrust calibration at the low speed simulation. The characteristics of the tank are as follows. (1) The sonic pressure ratio, and the pressure ratio equivalent to the simulating Mach number can be set up in the tank. (2) The influence on the balance can be reduced to some extent by introducing a differential pressure compensating bellows system, an interference elimination system with six degrees of freedom of piping, etc. (3) It is necessary to use the air-exhaust ventilator of the transonic wind tunnel or an apparatus with the ventilating capacity similar to the ventilator for the fixation of transonic pressure ratio., STOL(短距離着陸機)実験機「飛鳥」の開発を契機に航技研の遷音速風洞に、TPS(タービン動力模擬エンジン)の推力検定装置を作った。この装置では、TPSの推力を天秤で、またファン流量を音速流量計で精度良く測定する。また、風洞試験時のマッハ数に相当する圧力比を得るために全体が気密容器になっている。本報では、装置の概要、機能と特性、および低速模擬推力検定について説明を行った。本装置の特徴は次の通りである。(1)装置内に音速圧力比、模擬マッハ数に相当する圧力比を設定できる。(2)差圧補償ベローズ機構、6自由度配管干渉除去機構などの採用により天秤への影響をかなり少なくできる。(3)遷音速圧力比の設定には遷音速風洞の排風機かそれと同様の排風能力を持つ機器を使用する必要がある。

This paper presents comparison of EFD/CFD results submitted to the 3rd SST-WorkShop held in December 2001. Comparison among CFD results shows mostly good agreement. However, it shows big discrepancy in C(sub D) estimation. Comparison between CFD and EFD shows mostly good agreement except in several aspects. A big discrepancy of C(sub p) distributions can be seen at the leading edge on the upper surface as well as at the outer wing. The comparison also shows a small difference in C(sub L-alpha) and C(sub M-alpha), indicating good agreement in most of C(sub L) and C(sub D) estimation. Sensitivity analysis for free stream Mach number and attack angle, as well as dependency analysis for the limiter functions of numerical scheme have been conducted as a verification of CFD, indicating that the discrepancies between EFD and CFD can not be resolved by these considerations alone., Special Publication of National Aerospace Laboratory, 航空宇宙技術研究所特別資料

The Jet-Powered Supersonic Experimental Airplane (NEXST-2) being developed by the National Aerospace Laboratory (NAL) will have a high-speed propulsion system with two-dimensional external compression air-intake with variable ramps. This report describes the aerodynamic design of the NEXST-2 second configuration air-intake (C313 air-intake) and its wind tunnel test results. The C313 air-intake was designed at a Mach number of 2.0, using both a 1D empirical estimation method and 3D CFD analysis. The main objective of the wind tunnel tests was to obtain the aerodynamic characteristics required to estimate the performance of the propulsion system, and develop the air-intake control system. The wind tunnel tests were performed in Mach numbers ranging from 0.6 to 2.1 at the 2 m x 2 m Transonic Wind Tunnel and 1 m x 1 m Supersonic Wind Tunnel at NAL. The wind tunnel model is a 19.2 percent scale isolated C313 air-intake model. The aerodynamic characteristics obtained in the wind tunnel tests are in reasonable agreement with those predicted by the CFD analysis used for the aerodynamic design, so aerodynamic performance such as pressure recovery and distortion at the engine face met the design target. At higher Mach numbers (Mo greater than 1.8) however, buzz initiated by Ferri-instability was observed at a relatively high mass flow ratio in subcritical conditions. Consequently the stable margin in subcritical operations was smaller at higher Mach numbers. The test results also indicated that the state of air-intake flow could be appropriately correlated with the pressure ratio in the bleed plenum chamber to the freestream total pressure, used for the air-intake control. Side-wash upstream of the air-intake did not affect the aerodynamic performance at side-wash angles of less than 4 degrees., 航空宇宙技術研究所が開発を進めているジェットエンジンを搭載した小型超音速実験機(ジェット実験機)には、マッハ2までの飛行速度で作動する2次元外部圧縮型可変形状超音速インテークを搭載する予定としている。ジェット実験機の推進システム性能予測およびインテーク可変制御開発に必要なインテーク空力特性を取得するために、ジェット実験機第2次形状超音速インテーク(C313形状)の19.2％縮尺模型を用いた単体空力特性取得試験を航空宇宙技術研究所の遷音速風洞および超音速風洞においてマッハ0.6〜2.1の範囲で実施した。ジェット実験機の推進システム性能予測およびインテーク可変制御開発に必要なインテーク空力特性データを取得するとともに、風洞試験で得られた空力特性がCFD解析により事前に予測した特性と概ね一致し、設計目標を満足した。また、マッハ1.3以上においては低流量側でバズが発生し、特にマッハ1.8以上の高マッハ域においてはFerri型不安定によるバズ発生のためインテークの安定作動余裕が小さいことが判明したが、適切な超音速ランプ可変制御および抽気制御によりインテークの安定作動域を改善できる。さらに抽気プレナム圧力比は本試験条件の範囲においては第2ランプ角や抽気条件による影響は小さく、インテークの作動状態を適切に示すパラメタとなることを確認した。インテーク入口部における横流れの影響については横流れ偏角4度までの範囲において、試験を行った何れのマッハ数においても大きな空力性能の劣化は見られず、良好な特性を示した。, Technical Memorandum of National Aerospace Laboratory, 航空宇宙技術研究所資料

Direct Numerical Simulations of compressible, homogeneous, decaying isotropic turbulence are conducted using a Fourier pseudo-spectral method in order to deepen the understanding of physics and mechanics of compressible turbulence and to generate statistical data bases useful for turbulence modeling. The simulations are carried out on a vector-parallel supercomputer Numerical Wind Tunnel (NWT) at National Aerospace Laboratory (NAL) for a 128(exp 3) grid. The existence of eddy shocklets is confirmed at high turbulent Mach numbers. The time evolutions of turbulence statistical are also analyzed, indicating that the compressibility effects on turbulence is significant at higher fluctuating Mach numbers., 圧縮性流体の乱れの物理に関する理解を深めるため、また、乱流モデルに有用な統計データベースを作り出すためフーリエ疑スペクトル法を用いて減衰過程の圧縮性一様等方性乱流の直接シミュレーションを行った。シミュレーションは、NALのベクトル並列処理型スーパーコンピュータ上で128{3}の格子により実行した。高い乱流マッハ数におけるeddy shockletの存在が確認された。乱流統計量の時間的発展を解析し、ゆらぎのマッハ数が高い場合には、乱流に対する圧縮性の影響が著しいことを示した。

The flight range of HOPE (H-2 Orbiting Plane) would be extended to the high altitude of several hundreds km and its speed to the Mach number of 30. In this regard, the aerodynamic characteristics must be estimated for various flight conditions such as low speed during climbs, high speed at low angles of attack, high angles of attack during re-entries, and low speed at landing approach. The aerodynamic characteristics of HOPE were to be estimated in three flight ranges, segmented by different flight Mach number ranges. Items to be estimated were mainly lift to drag ratios, static and dynamic stability in longitudinal and lateral directions, some aerodynamic characteristics particular to the HOPE configuration, jet reaction control systems, boundary layer aerodynamic heating, vortex and flow separation, shock wave interactions, and real gas effects. In respect to numerical simulations, the aerodynamic characteristics became independent of the Mach number in flight as the Mach number increases. A Newtonian flow approximation, however, still held, so that hypersonic aerodynamic characteristics could be dealt with good accuracy. Tools for estimation must be verified by numerical simulations, wind tunnel tests, and flight tests., HOPE(H-2打ち上げ型有翼回収機)の飛行領域は、数百kmの高い高度およびマッハ数30の大きな速度領域に亘っている。このため、上昇時の低速、小迎角の高速、帰還時の高速大迎角、着陸進入時の低速大迎角などの空力特性の推定が必要である。このようなHOPEの空力特性推定は、飛行マッハ数領域を3分割して行われ、また推定項目は、揚抗比、縦横の静安定と動安定、HOPE機体形状による空力特性、ジェット反動コントロール、境界層空力加熱、渦および剥離、衝撃波干渉、および実在気体効果などを含んでいる。数値シミュレーションでは、マッハ数が高くなると空力特性はマッハ数に依存しなくなるが、ニュートン流近似が成立するので、極超音速空力特性を良い精度で取扱うことができた。また、推定ツールは、数値シミュレーション、風洞試験および飛行試験による実証が必要であった。

When computational simulation for low Mach number flows is carried out with a compressible flow solver, its performance deteriorates in terms of both convergence speed and accuracy. Difficulties with low speed flow calculation are caused by the disparate particle and acoustic wave speed. Local preconditioning method for the low speed flow minimizes this wave speed range to improve convergence of an iterative method. When calculating unsteady flows, this difficulty also arises. Therefore, preconditioning method is applied to the inner iteration for the unsteady calculation. This paper reports the accuracy of the basic low speed unsteady flows calculation with preconditioning method., JAXA Special Publication, 宇宙航空研究開発機構特別資料

The National Aerospace Laboratory, Kakuda Propulsion Laboratory completed the RamJet engine Test Facility (RJTF) and has been carrying out tests using H2-fueled scramjet engines and CH4-fueled ramjet engines since 1994. A 1/5-subscale wind tunnel of the RJTF (freestream Mach numbers = 3.4, 5.4 and 6.7 at the nozzle exits, 100 mm x 100 mm) was constructed to supplement the engine tests. A special group was formed in the Ramjet Research Division to conduct these studies and to develop new techniques for evaluation of aerodynamic and combustion performances. In this report, details of the subscale wind tunnel and the studies are summarized. First, the design, calibration of the wind tunnels and transient loads during the wind tunnel start are summarized, then various supplementary studies, i.e., engine alignments, air capture ratios, oil flow on engine internal walls, drag and wall pressure measurements and simulation of engine unstart process are explained. Studies on the mixing process of fuel injected into engines and on the ramp-compression-type inlets for a subsonic combustion ramjet engine are also introduced. A series of combustor experiments using cold flow was conducted to investigate flow structure around the fuel injectors behind backward-facing steps in scramjet combustors. In the component experiments, optical measurements, such as Particle Image Velocimetry (PIV) and Pressure Sensitive Paint (PSP) methods were applied. Measurement techniques for the RJTF engine tests were also developed in these studies. Freezing criteria of chemical reactions in gas sampling probes, six-force balances and response times in low pressure measurements are reviewed. In the appendix, designs of the M3.4-, M5.4- and M6.7-two dimensional, square, and M8-conical nozzles, an air-driven ejector and a noise suppressor are summarized, as are construction of a new wind tunnel and the heat loss in storage air heaters. Finally, the know-how accumulated during RJTF operations is mention, 航空宇宙技術研究所・角田推進研究所ではラムジェットエンジン試験設備(RJTF)を完成し、水素を燃料とするスクラムジェットエンジンやメタン燃料ラムジェット(HYPR用)エンジンを試験している。このRJTFのエンジン試験を支援するためにRJTFの1/5縮尺の小型風洞を整備し、ラムジェット推進研究部内に横断的な小型風洞実験班を組織した。小型風洞実験班の目的は2つある。第1は、RJTF試験に関わる問題を解決するためのパイロット試験を実施する事である。第2の目的は、RJTFエンジン試験の補完実験を行う事である。本報告書では、この小型風洞を用いた全ての研究とノウハウについてまとめた。最初に小型風洞設備の構造について説明し、次に自由噴流型風洞の始動、風洞およびエンジン不始動時の荷重、風洞ノズル検定について述べる。そしてこれまでのエンジン試験を補完する実験、スクラムジェットエンジンにおける空気捕獲率の測定、オイルフロー実験をまとめた。また抗力測定と壁圧分布測定を実施した。これらのデータからエンジン内部流解析が可能となり、エンジン理論燃焼性能を予測した。また、燃焼によるエンジン不始動移行過程を調べる模擬実験や燃料拡散を調べるためのHe噴射実験、亜音速燃焼ラムジェット用ランプ圧縮型インレット性能も試験した。小型風洞ではエンジン全体の空力測定に加えて超音速燃焼器の要素実験を行い、燃料ジェットと主流との干渉や燃料拡散を調べた。またレーザによるPIVや感圧塗料実験も実施した。計測に関してはRJTFエンジン試験を支援するための6分力や圧力測定における応答性や、ガス採取プローブでの反応凍結を燃焼模擬空気を用いた超音速風洞で検討した。付録にはノズル設計法と形状データ、小型風洞用エジェクタと消音器の設計、第2(K)風洞の整備、蓄熱体加熱器の蓄熱特性、そして最後にRJTF設備とそのエンジン試験の中で蓄積したノウハウについてまとめた。, Technical Report of National Aerospace Laboratory, 航空宇宙技術研究所報告

Operating conditions of a rocket-ramjet combined cycle engine for a single-stage-to-orbit aerospace plane were studied. The operation modes of the engine included an ejector-jet mode, a ramjet mode, a scramjet mode and a rocket mode. Characteristics of the engine operating conditions were studied analytically. Study of the effective impulse function revealed that a higher specific impulse was preferable in supersonic flight, whereas a greater thrust coefficient was preferable in hypersonic flight. These characteristics were examined by simulating an engine operating in an aerospace plane flight. A comparison was made between several engines with different combinations of thrust and specific impulse when used for the transportation of a mass into orbit. The mass that could be carried into orbit was larger with a ramjet mode of higher specific impulse and with a scramjet mode of greater thrust. In the present study, the thrust augmentation effect of the ejector-jet mode was also found to be small at low subsonic speeds and to increase with increasing flight Mach number., 単段式宇宙機用のロケット-ラムジェット複合サイクルエンジンの作動条件について検討を行った。このエンジンはエジェクタージェットモード、ラムジェットモード、スクラムジェットモード、そしてロケットモードで構成される。エンジン作動特性について、解析的に検討を行なった。有効比推力についての検討から、超音速域では高比推力での作動が望ましく、極超音速域では推力の増強が必要であることが明らかとなった。これらの結果を、宇宙機の飛行シミュレーションによって検討した。比推力と推力の組合せを変えた幾つかのエンジンを使って、低軌道への運搬質量を比較した。その結果、高比推力のラムジェットモードおよび大推力のスクラムジェットモードによってより多くの質量を運べることがわかった。その他、低亜音速域では、エジェクタージェットモードの推力増強効果は小さく、マッハ数の増加と共にその効果は増加することもわかった。, JAXA Research and Development Report, 宇宙航空研究開発機構研究開発報告

JAXA Special Publication, 宇宙航空研究開発機構特別資料

Force tests were carried out using an AGARD-B calibration model in the JAXA 0.5 m Hypersonic Wind Tunnel (HWT1) at a nominal Mach number of 5. The aerodynamic characteristics of the AGARD-B calibration model were obtained successfully. The experimental results showed good agreement in comparison with those of other wind tunnels. The data obtained at the experiment and other tunnels were correlated well with the free stream Mach number., JAXA Special Publication, 宇宙航空研究開発機構特別資料

JAXA Special Publication, 宇宙航空研究開発機構特別資料

As a part of the NAL's scaled supersonic experimental airplane program to establish an aerodynamic design system based on Computational Fluid Dynamics (CFD), an aerodynamic configuration with no propulsion system was designed in a two-stage process. First, a baseline configuration was designed using a conventional linear theory. Then, the CFD and a supersonic inverse method were used to refine the wing geometry and to achieve a higher lift-to-drag ratio at a design point of M = 2 and CL = 0.1. A non-linear effect such as body to wind interference and a wing thickness effect were also handled in this phase. By making use of the inverse method, an upper surface of the wing was designed aiming at a natural laminar flow with a flat-type pressure distribution to reduce a friction drag. A wing warp was also adjusted to achieve an optimal load distribution designed by the linear theory to reduce a pressure drag. The designed new wing was evaluated by a Navier-Stokes analysis and an incompressible boundary layer stability code (SALLY code) and was found to have an improved lift-to-drag ratio, with a wider laminar flow regime and smaller friction drag than the initial geometry. Finally, the CFD analysis was compared with supersonic wind tunnel data and was found to be a well validated tool to be used in the design., JAXA Special Publication, 宇宙航空研究開発機構特別資料

We consider the propagation of sound through a slowly moving fluid in a 2-dimensional duct. A detailed description of a flow-acoustic model of the problem using B-spline based isogeometric analysis is given. The model couples the non-linear, steady-state, incompressible Navier-Stokes equation in the laminar regime for the flow field, to a linear, time-harmonic acoustic equation in the low Mach number regime for the sound signal. B-splines are used both to represent the duct geometry and to approximate the flow and sound fields. This facilitates an exact representation of complex duct geometries, as well as high continuity approximations of state variables. Acoustic boundary conditions on artificial truncation boundaries are treated using a mode matching formulation. We validate the model against known acoustic modes for a uniform flow through a straight duct. Improved error convergence rates are found when the acoustic pressure is approximated by higher order polynomials. Based on the model, we examine how the acoustic signal varies with sound frequency, flow speed and duct geometry. A combination of duct geometry and sound frequency is identified for which the acoustic signal is particularly sensitive to the flow speed.

The hypersonic laminar to turbulent transition problem above Mach 10 is addressed experimentally in the short duration VKI Longshot gun tunnel. Reentry conditions are partially duplicated in terms of Mach and Reynolds numbers. Pure nitrogen is used as a test gas with flow enthalpies sufficiently low to avoid its dissociation, thus approaching a perfect gas behavior. The stabilizing effects of Mach number and nosetip bluntness on the development of natural boundary layer disturbances are evaluated over a 7 degrees half-angle conical geometry without angle of attack. Emphasis is initially placed on the flow characterization of the Longshot wind tunnel where these experiments are performed. Free-stream static pressure diagnostics are implemented in order to complete existing stagnation point pressure and heat flux measurements on a hemispherical probe. An alternative method used to determine accurate free-stream flow conditions is then derived following a rigorous theoretical approach coupled to the VKI Mutation thermo-chemical library. Resulting sensitivities of free-stream quantities to the experimental inputs are determined and the corresponding uncertainties are quantified and discussed. The benefits of this different approach are underlined, revealing the severe weaknesses of traditional methods based on the measurement of reservoir conditions and the following assumptions of an isentropic and adiabatic flow through the nozzle. The operational map of the Longshot wind tunnel is redefined accordingly. The practical limits associated with the onset of nitrogen flow condensation under non-equilibrium conditions are also accounted for. Boundary layer transition experiments are then performed in this environment with free-stream Mach numbers ranging between 10-12. Instrumentation along the 800mm long conical model includes flush-mounted thermocouples and fast-response pressure sensors. Transition locations on sharp cones compare favorably with engineering c, Doctorat en Sciences de l'ingénieur, info:eu-repo/semantics/nonPublished