野中 聡

A slender-bodied vehicle with asymmetrically arranged protuberance generates strong side force due to asymmetric vortices, even at a low angle-of-attack. We investigated effects of the well-known RANS turbulence models [SA-R (C-rot = 0.0, 1.0, and 2.0), SST, and SST-2003] by comparing the numerically obtained side force values on supersonic slender-body, along with the flow structure. As a result, all the SA-R models showed good agreement with the experiment regardless of the C-rot (which controls the degree of modification from the original SA model), although a separation point on the protuberance side slightly changed depending on the C-rot value. On the other hand, as for the SST models, when the vorticity was used to evaluate eddy viscosity (original SST) the side force exhibit 44% deviation from the experiment, whereas SST-2003, in which the strain rate was employed instead, significantly reduced the discrepancy to 0.7%.

Typical launch vehicles feature several surface devices, such as ducts, cameras, and side jet motors. When these protuberances are asymmetrically distributed, an asymmetric vortex is formed around the vehicle, generating a side force. Furthermore, the side force induced by one protuberance may be reduced or increased by another, depending on their arrangement. Therefore, in this study, the side force characteristics and associated flowfield of a slender body with two protuberances were investigated for the Mach number 1.5. One of the protuberances was mounted leeward of the front portion of the vehicle (front protuberance), while the other was mounted at various azimuthal positions in the middle portion (middle protuberance), and both were on the port side. In particular, when the middle protuberance was on the windward side, its wake vortex was in proximity to the vehicle, reducing the side force. By contrast, when the middle protuberance was in the vehicle's horizontal plane, the near-wall flow was entrained by the wake vortex of the middle protuberance, causing the flow deceleration. This induced high surface pressure on the port side, yielding an extremely high side force.

To accurately obtain a drag value for slender bodies (e.g., a rocket used in wind tunnel testing), base pressure and cavity pressure are required to be corrected to remove the effect of a sting. As the drag value significantly depends on the correction of the base/cavity pressure, attention should be paid, in spite of the absence of clear guidelines, to decide these two pressure values. In this study, to resolve this issue, wind tunnel test and numerical calculation results are used. As a result, first, we found that it is appropriate to measure the pressure near the base surface at four points and average them to estimate the base pressure. Second, the cavity pressure can be estimated the same as the uniform-flow static pressure even when shock waves are generated around the vehicle body. Based on these corrections, the total drag coefficient, C-D, showed an effective agreement between the wind tunnel test and computational fluid dynamics. Furthermore, in absence of the sting, the drag value increased by 8%.

“Nose-first entry” flight has been proposed as one of the methods of return flight of a vertical take-off and vertical landing reusable rocket using engine thrust for vertical landing. In this flight method, the engine exhaust jet opposes the free-stream during the turnover maneuver and landing, causing concern that the aerodynamic force acting on the vehicle changes due to a complicated flow field made by the interaction between the exhaust plume and free-stream. A slender-body model that can eject a supersonic jet was studied in a low-speed wind tunnel to characterize the flow. The aerodynamic forces were measured, and the flow pattern on the surface of the model was visualized using the oil-flow technique. The results indicate that the axial force decreases in the low angle-of-attack region, and the rate of change in the axial force is much smaller compared with previous studies in which the jet is ejected from a blunt configuration. The surface flow pattern is also changed by the jet ejection. However, the normal force and pitching moment do not change. Therefore, the influence of the jet strongly depends on the vehicle shape, and the aerodynamic characteristics are restrictive in slender-body shape rockets.

We focused on the liquid hydrogen (LH2) cooled field winding superconducting generator and conducted a test to develop a liquid hydrogen cooling system for a rotor of superconducting generator. The LH2 supply and exhaust system for the rotating tank which simulated the generator rotor was designed and fabricated. The rotation test was actually performed at a speed of 1800 rpm successfully and safely in a state where LH2 was stored and continuously kept liquid level in the rotating tank. The rotating tank equipped with eight thermometers, two MgB2 superconducting liquid level meters to know the state of liquid hydrogen in the tank. A rotation speed variation test was performed, and the behavior of the temperature and the liquid level in the rotating tank was observed. The vibrations of the rotor shaft throughout the test were within the allowable value. The experiments demonstrated the LH2 cooling system of the superconducting generator rotor.

© 2020 Elsevier Ltd Heat transfer from inner surface of a vertical heated pipe to subcooled or saturated liquid hydrogen flowing upward was measured for quasi-steadily increasing heat input up to fully developed film boiling regime and decreasing the heat input through the film boiling regime. The experiments were carried out at the inlet pressures of 400, 700 and 1100 kPa and subcoolings from 0 K to 11 K. Three test pipe heaters made of SS310S with inner diameters of 6 and 8 mm and lengths of 100 and 200 mm were used. Experimental data from non-boiling to developed film boiling and developed film boiling down to minimum film boiling was obtained with the record of mass flow rate by continuously increasing and decreasing the heat input. It was observed that though the mass flow rate decreases with the increase of the heat generation rate, the heat transfer coefficient increases. Discussions on heat and mass transfer in inverted annular flow, dispersed droplet flow, single-phase vapor flow regimes and their changing conditions from one by one were carried out to clarify the phenomena. A calculation code of heat transfer characteristics was developed based on the discussions. The calculated results are in good agreement with the experimental results.

© Published under licence by IOP Publishing Ltd. Our research group has been researched for developing MgB2 superconducting energy apparatus in liquid hydrogen immersion cooling, such as superconducting generators. In this study, two-pole MgB2 race-Track coil, which was a model for a few tens kVA superconducting generator field magnet, was designed and made. This coil consists of two pieces of a similar MgB2 race-Track coil which has 529 turn with straight section of 150 mm, bending diameter in the end of section of 100 mm and thickness of 34 mm. We carried out excitation tests of the coil immersed in liquid hydrogen at the temperature of 21 K to 32 K, and the load line of the coil was obtained. The critical current under self-field was obtained for various temperatures. The normal zone propagation behavior of the coil at the quench was also investigated using several potential taps installed in the coil.

© 2019 Hydrogen Energy Publications LLC This paper presents a hydrogen ignition experiment conducted to establish safety standards for high-pressure hydrogen handled at the hydrogen stations for fuel cell vehicles (FCV). In the experiment, cryogenic hydrogen pressurized to over 80 MPa was leaked from a pinhole nozzle, and the blast pressure at the ignition and the flame length during steady combustion were measured. The hydrogen supply equipment used in the experiment has a maximum flow rate of 100 kg/h, a maximum discharge pressure of 90 MPa, and a temperature adjustment range of 50 K–300 K. Four types of pinhole nozzles with different outlet diameters, viz. 0.2 mm, 0.4 mm, 0.7 mm, and 1.0 mm were used to leak the hydrogen. In the experiment, the effects of the pinhole nozzle diameter, hydrogen supply pressure and temperature, and an igniter location on the blast pressure and flame length were evaluated. The igniter being appropriately positioned, once a steady flame was formed, combustion continued even if the ignition source was turned off, which necessitated the stopping of hydrogen supply to extinguish the fire. As a result of the experiment, it was found that the blast pressure and the flame length can be expressed as the correlation equations of the hydrogen leakage flow rate. However, even if the leakage flow rate was the same, we found that the flame length increases with decreasing the hydrogen supply temperature. We presented a correlation equation for the cryo-compressed hydrogen flame length that is about 30% longer than the previously presented equations for 300 K hydrogen flame.

© 2020 American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. To overcome the shortcomings of the schlieren visualization method, which is the difficulties in capturing the three-dimensional flow structures, the application of the computed tomography has been studied. This study aims to investigate the usefulness of this method for complicated flow structures in the wind tunnel. By applying this method to the flow field around a rocket-like slender body with a medium angle of attack in a supersonic flow, the following two issues were examined. The one is the influence of the shock waves reflected at the wind tunnel wall (especially the visualization windows). The results show that the artifacts caused by these shock waves can be reduced by using the buffer region in reconstruction. The other is the fine structures around the body. The results show that this method has enough resolution to capture the fine structures such as the vortices separated from the body, shock waves diffracted by the vortices, shock waves caused by the vortices, and complicated shock structures around protuberances. In particular, the reconstructed results show the asymmetry of the vortices due to the protuberances, suggesting that this method is useful for comparison with the numerical simulations.

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Modal analyses are data-driven techniques for coherent-structure extraction that have been widely used for understanding and modeling jet aeroacoustics. For the same goal, another approach is the conditional sampling analysis that is an event-driven technique to extract turbulent structures and related acoustic waves, which are inherently nonlinear and intermittent phenomena. This study aims to examine the consistency and differences between the data-driven modal analyses and event-driven conditional sampling analysis for the wavepacket structures of the Mach waves of a Mach 1.8 ideally expanded jet, by applying these techniques to the schlieren visualization movies. The results of the conditional sampling analysis (i.e., conditional averages) show convected wavepackets that correlate with intermittent loud events in far-field in an aft-angle direction of the jet. These convected wavepackets are then compared with the wavepackets extracted using the modal analyses. Among three modal analyses examined, the spectral proper orthogonal decomposition captures wavepackets that have almost the same features to those of the convected wavepackets, such as the peak location of the amplitude envelope, propagation direction and speed of the waves, and the convection speed in the downstream direction. On the other hand, possibly due to the intermittent and convected nature, multiple-peak wavepackets are.

<p>Conventional rockets are faced with several problems such as high launching cost. Therefore, in Japan, a reusable vertical-takeoff-and-vertical-landing (VTVL) rocket vehicle is being developed. This vehicle utilizes nose entry as the return flight system including the attitude change (turnover) due to aerodynamic forces. To safely achieve turnover, it is necessary to reduce the difference between the maximum value and minimum value of C_m (i.e., pitching-moment coefficient). In this study, a delta-wing with vortex flaps (developed for the aircraft industry) is attached to the aft of the vehicle with the expectation of improving the C_m characteristics during the turnover process. Consequently, when the flap deflection angle is 0°, the nose-up C_m can be reduced at forward angles (i.e., AOA 0° - 90°) because vortices generated by the fins result in a nose-down C_m and cancel the nose-up C_m. Moreover, when the flap deflection angle is -30°, the nose-down C_m is enhanced at the backward angles (i.e., AOA 90° - 180°) because the flaps reduce the vortices generated by fins. Hence, setting the flap deflection angle at 0° for the forward angles and -30° for the backward angles reduced the difference between the maximum and minimum values of C_m (i.e., 12% smaller than a conventional model).</p>