When an issued liquid jet breaks up into droplets at a short distance from the nozzle exit, its surface deforms significantly over its whole length. The one-dimensional model developed in the first report is refined so as to describe the deformation of the liquid jet tip and it is applied to analyze strong interactions between the capillary deformation of the jet and the nozzle exit, revealing the highly stochastic nature of the breakup phenomena. It is found that the Rayleigh type of disintegration of a liquid jet takes place only under the condition that the liquid Weber number is less than 1.5 and the liquid jet disintegrates relatively far apart from the nozzle exit. Otherwise, the regular disintegration of the jet can not occur without the help of aerodynamic action.

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A new thermal control material named the Smart Radiation Device (SRD) has shown improvement in development. The SRD can be used as a variable emittance radiator that controls the heat radiated into deep space without assistances of any electrical instruments or mechanical parts. Its total hemispherical emittance changes from low to high as the temperature increases. This new device reduces the energy consumption of the on-board heater, and decreases the weight and the cost of the thermal control system (TCS). Space environmental simulation tests on the ground were performed, and the first generation of the SRD has been demonstrating success on the MUSES-C ‘HAYABUSA’ spacecraft that was launched in May 2003. During its cruise on the orbit, the distance from the spacecraft to the sun varied from 0.86AU to 1.70AU. As the spacecraft experienced solar intensity variation by a factor 4, it was effective to use the variable emittance radiator for decreasing the heater power. In-orbit temperature indicated that the SRD had successfully minimized component temperature variation and saved heater power, as expected. With the opportunity to validate the SRD in space, this lightweight and low cost thermal control device offers a possibility for flexible thermal control on future spacecrafts.

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This paper addressed a design method for cable network systems that primarily define characteristics of large mesh antenna reflectors. In the design process of the cable network systems, many parameters and conditions are required to be determined such as network topologies, geometries and cable maximum tensions. However, these parameters are defined by trial and error against the design constraints and there is no established methodology. In this paper, the requirements of the mesh antenna reflectors were clarified and a design methodology of the cable network systems was presented. By using of proposed method, a desirable cable network system, which is content with the requirements and constraints for an antenna reflector, is obtained.

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In case of a polar orbiting satellite, insulative surfaces located on the wake side of solar array paddle can be charged negatively by aurora electrons. We carried out laboratory experiments to evaluate risks of charging and arcing on thermal control film surfaces attached on a solar array paddle backside. We used two test coupons of different design. Both of them used the same materials of flight hardware as the Advanced Land Observing Satellite (ALOS) and the Optical Inter-orbit Communications Engineering Test Satellite (OICETS), which are domestic polar orbiting satellites. When the thermal control film surface was insulated and a silver layer was not connected to coupon ground, the film surface was charged to a negative value 1–2kV lower than the electron beam energy. Many primary arcs and flashover were observed at the edge of silver layer and film surface, respectively. Some areas of silver layer were destroyed by primary arcs. On the other hand, the film with conductive coating connected to a substrate with resistance of 10⁶–10⁹Ω suppressed surface charging and flashover on the film under the electron flux of 1×10¹⁵–1.4×10¹⁷m^{- 2}s^{- 1}, effectively. We made a new coupon which had three features to give improved performance of arcing and charging mitigation. (1) Connecting the silver layer to the substrate by conductive adhesive to suppress charging and arcing at the edge of silver layer. (2) Film surface coating by conductive material and connected with conductive adhesive to avoid charging. (3) Hiding all the edges of silver layer by conductive adhesive. This coupon suppressed any arcing and charging on the film surface up to 20keV electron beam irradiation.

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It has been recognized that damage resistance and strength properties of CFRP laminates can be improved by using thin-ply prepregs. This study investigates the damage behaviors and compressive strength of CFRP laminates using thin-ply and standard prepregs subjected to out-of-plane impact loadings. CFRP laminates used for the evaluation are prepared using the standard prepregs, thin-ply prepregs, and combinations of the both. Weight-drop impact test and post-impact compression test of quasi-isotropic laminates are performed. It is shown that the damage behaviors are different between the thin-ply and the standard laminates, and the compression-after-impact strength is improved by using thin-ply prepregs. Effects of the use of thin-ply prepregs and the layout of thin-ply layers on the damage behaviors and compression-after-impact properties are discussed based on the experimental results.

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An analysis model for numerical estimation of eigenvalues of “a cable-towed glider” system is presented in this paper. This paper takes up longitudinal dynamics of the system. A cable dynamics is formulated as a distributed parameter system with flexibility extensibility and curvature of cable. For a specific example of a horizontal towing validity of the estimated eigenvalues is verified by the results of an identification of dominant modes found in the transient response of the system and stability of the system is analyzed based on root-loci regarding the relative height of a glider to a tug-plane. Modes dominantly affected by curvature of cable are also specified.

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The paper reports design process of a space boom to be installed in a 50kg-class micro-satellite, tentatively called as SOHLA-1. The total mass of the boom is 483g, while the length of the boom is 416mm. A spring-string type pin-puller was proposed. Vibration test with a bread-board model (BBM) of the boom leaded to an engineering model (EM) of the boom accommodated with a two-pin type pin-puller. Vibration test with the boom EM suggested some modifications to be made for a proto-type model (PM) of the boom. Vibration test, deployment test and thermal test on the boom PM confirmed feasibility of the flight model (FM) of the boom.

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