野中 聡

Film boiling heat transfer coefficients in liquid hydrogen were measured for the heater surface superheats to 300 K under pressures from 0.4 to 1.1 MPa, liquid subcoolings to 11 K and flow velocities to 8 m/s. Two test wires were both 1.2 mm in diameter, 120 mm and 200 mm in lengths and were made of PtCo alloy. The test wires were located on the center of 8 mm and 5 mm diameter conduits of FRP (Fiber Reinforced Plastics). Furthermore film boiling heat transfer coefficients in liquid nitrogen were measured only for the 200 mm long wire. The film boiling heat transfer coefficients are higher for higher pressure, higher subcooling, and higher flow velocity. The experimental data were compared with a conventional equation for forced flow film boiling in a wide channel. The data for the 8 mm diameter conduit were about 1.7 times and those for the 5 mm conduit were about 1.9 times higher than the predicted values by the equation. A new equation was presented modifying the conventional equation based on the liquid hydrogen and liquid nitrogen data. The experimental data were expressed well by the equation.

Forced flow heat transfer of hydrogen from a round wire in a vertically-mounted pipe was measured at pressure of 1.5 MPa and temperature of 21 K by applying electrical current to give an exponential heat input (Q=Q0exp(t/τ),τ=10 s) to the round wire. Two round wire heaters, which were made of Pt-Co alloy, with a diameter of 1.2 mm and lengths of 54.5 and 120 mm were set on the central axis of a flow channel made of FRP with inner diameter of 5.7 and 8.0 mm, respectively. Supercritical hydrogen flowed upward in the channel. Flow velocities were varied from 1 to 12.5 m/s. The heat transfer coefficients of supercritical hydrogen were compared with the conventional correlation presented by Shiotsu et al. It was confirmed that the heat transfer coefficients for a round wire were expressed well by the correlation using the hydraulic equivalent diameter.

Transient heat transfers from Pt-Co wire heaters inserted into vertically-mounted pipes, through which forced flow subcooled liquid hydrogen was passed, were measured by increasing the exponential heat input with various time periods at a pressure of 0.7 MPa and inlet temperature of 21 K. The flow velocities ranged from 0.3 to 7 m/s. The Pt-Co wire heaters had a diameter of 1.2 mm and lengths of 60 mm, 120 mm and 200 mm and were inserted into the pipes with diameters of 5.7mm, 8.0 mm, and 5.0 mm, respectively, which were made of Fiber reinforced plastic due to thermal insulation. With increase in the heat flux to the onset of nucleate boiling, surface temperature increased along the curve predicted by the Dittus-Boelter correlation for longer period, where it can be almost regarded as steady-state. For shorter period, the heat transfer became higher than the Dittus-Boelter correlation. In nucleate boiling regime, the heat flux steeply increased to the transient CHF (critical heat flux) heat flux, which became higher for shorter period. Effect of flow velocity, period, and heated geometry on the transient CHF heat flux was clarified.

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In the present experiment, by using the sounding rocket’s sub-orbital ballistic flight, realized the gravitational environment similar to that of liquid-fueled rockets during its coasting flight. In the low-gravity state, the cryogenic test fluid, liquid nitrogen, was introduced into the test sections which had similar shapes to the flow channels in the cryogenic propulsion systems. The boiling of liquid nitrogen inside the test-sections and the transition of flow regimes from gas/liquid two-phase flow to liquid mono-phase flow were successfully visualized. The temperatures, pressures and void fractions in each channel were measured as well. The mechanisms enhancing heat transfer were discussed based on the visualization. In the present case, compared with the corresponding ground test, it was confirmed that the two-phase flow in the complex channel could wet the heat transfer surfaces more easily due to the absence of gravity, and that more uniform chill-down effect could been obtained.

In order to make the access to space for researches by the sounding rocket much easier and make the opportunities of the rocket launches much frequent, a fully reusable sounding rocket is proposed, and its preliminary system design study is made progress in ISAS/JAXA. Technical demonstrations and system/subsystem designs for a reusable sounding rocket are also underway. The technical key issues to develop an operational system of reusable sounding rocket are 1) a reusable engine, 2) aerodynamics and returning flight mechanics, 3) cryogenic propellant management, 4) health management system, 5) repeated operations. In order to verify these important technologies, 1) repeated engine tests, 2) a model flight demonstration, 3) a reusable insulation development, 4) sloshing tests of cryogenic liquid propellant, 5) a landing gear development, 6) a construction of health management system for hydrogen gas leak, have been completed or are planned now. Progress of these recent activities for development of reusable sounding rocket is summarized in this paper.

© 2014 AIP Publishing LLC. Liquid hydrogen (LH2) is a key issue in a carbon-free energy infrastructure at the energy storage and transportation stage. The typical features of LH2 are low viscosity, large latent heat and small density, compared with other general liquids. It is necessary to measure a mass flow of liquid hydrogen with a simple and compact method, especially in a two phase separate flow condition. We have proposed applying a hot-wire type flow meter, which is usually used a for gas flow meter, to LH2 flow due to the quite low viscosity and density. A test model of a compact LH2 hot-wire flow meter to measure local flow velocities near and around an inside perimeter of a horizontal tube by resistance thermometry was designed and made. The model flow meter consists of two thin heater wires made of manganin fixed in a 10 mm-diameter and 40 mm-length tube flow path made of GFRP. Each rigid heater wire was set twisted by 90 degrees from the inlet to the outlet along the inner wall. In other words, the wires were aslant with regard to the LH2 stream line. The heated wire was cooled by flowing LH2, and the flow velocity was obtained by means of the difference of the cooling characteristic in response to the flow velocity. In this report, we show results on the basic experiments with the model LH2 hot-wire flow meter. First, the heat transfer characteristics of the two heater wires for several LH2 flow velocities were measured. Second, the heating current was controlled to keep the wire temperature constant for various flow velocities. The relations between the flow velocity and the heating current were measured. The feasibility of the proposed model was confirmed.

Copyright © 2014 by the International Astronautical Federation. At ISAS/JAXA, we have various space flight tools such as scientific satellites and planetary probe, various launchers including sounding rockets and balloons. In relatively small size of the flight program such as sounding rockets and balloons, we have been conducting both studies using these flight opportunities such as scientific observations, utilization of micro-G environment and so on. In order to make the access to space for researches by the sounding rocket much easier and make the opportunities of the rocket launches much frequent, a fully reusable sounding rocket is proposed, and its preliminary system design study is underway in ISAS/JAXA. In the phase-A, we are progressing the system design and technical demonstrations which are indispensable for the development of the operational system. The technical key issues are 1) a reusable engine, 2) aerodynamics and returning flight mechanics, 3) cryogenic propellant management, 4) health management system, 5) repeated operations. Present status of the reusable sounding rocket activities is summarized in this paper.

To make the opportunities of space scientific researches by using sounding rockets much frequent, a fully reusable sounding rocket is proposed, and system/subsystem designs such as aerodynamics, propulsion systems, structures, etc are progressed in ISAS/JAXA. In the phase- A, we are conducting the technical demonstrations which are indispensable for the development of the operational system. The technical key issues are 1) a reusable engine, 2) aerodynamics and returning flight mechanics, 3) cryogenic propellant management, 4) health management system, 5) repeated operations. The present status of system design and the demonstrations of these technologies planed in phase A are introduced in this paper. Copyright© (2013) by the International Astronautical Federation.

We have analyzed the asymmetric separation flow over a slender body at high angle of attack by numerical simulations aiming a control of the asymmetric vortices using a dielectric barrier discharge (DBD) plasma actuator. The Reynolds Averaged Navier Stokes/Large-Eddy Simulation hybrid method (RANS/LES) was adopted with the high-order compact spatial difference scheme. First, for investigating the characteristics of the asymmetric separation flow, the simulation of the flow field over the slender body was conducted for various angle of attack and bump height. Note that the bump is added near the body apex to simulate the symmetry-breaking imperfection. When the angle of attack or the bump becomes higher, the asymmetricity of vorticities becomes stronger. The side force has nonlinearity with the angle of attack or the bump height. Next, numerical simulations of the flow control using a plasma actuator were conducted. The side force coefficient can be continuously controlled in response to output power of the actuator within about ±1.0 on an average by the actuator's actuation at the aft body. However, the flow control effect is totally difference between starboard-side actuator's actuation and port-side actuator actuation. In addition, it is strongly influenced by the angle of attack.

A fully reusable sounding rocket is proposed and conceptually designed in ISAS/JAXA. In phase A in the proposed project, technical demonstrations for key technologies to develop the reusable sounding rocket are planed as follows; 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development and 6) health management system construction. In this paper, the present system design of the reusable sounding rocket and technical demonstrations are summarized.

The Epsilon launch vehicle, successor of the M-V rocket, which conveyed "Hayabusa" is currently under development in Japan. The Epsilon is also designed for sending scientific satellites to outer space, and its first flight is scheduled to be in 2013. In this study, by conducting both numerical simulations and wind-tunnel tests, the aerodynamic characteristics and associated flow features of the Epsilon launch vehicle are extensively investigated at Mach 1.5. The results provided are axial/normal/side forces, pitching/yawing/rolling moments, detailed threedimensional flow structure, along with effects of the Reynolds number (between wind-tunnel and flight conditions), skin stringers (small devices on the main body), and the difference from another configuration called "NextGenEpsilon". This set of data includes unavailable ones at either the experiment standalone or the actual flight. Magnitudes of computed aerodynamic coefficients are in good agreement with the experiment and within the design criteria. According to the results axial force is not affected by those above-mentioned factors, but local normal force is influenced around skin stringers by the Reynolds number, and the rolling moment can change even its sign. Detailed explanationswith regard to relations between aerodynamic characteristics and those flowstructures are also provided. © 2012 AIAA.

We have analyzed the asymmetric separation flow over a slender body at high angle of attack by numerical simulations aiming a control of the asymmetric vortices using a dielectric barrier discharge (DBD) plasma actuator. The Reynolds Averaged Navier Stokes/Large-Eddy Simulation hybrid method (RANS/LES) was adopted with the high-order compact spatial difference scheme. First, for investigating the characteristics of the asymmetric separation flow, the simulation of the flow field over the slender body was conducted for various angle of attack and bump height. Note that the bump is added near the body apex to simulate the symmetry-breaking imperfection. When the angle of attack or the bump becomes higher, the asymmetricity of vorticities becomes stronger. The side force has nonlinearity with the angle of attack or the bump height. Next, numerical simulations of the flow control using a plasma actuator were conducted. The side force coefficient can be continuously controlled in response to output power of the actuator within about +/- 1.0 on an average by the actuator's actuation at the aft body. However, the flow control effect is totally difference between starboard-side actuator's actuation and port-side actuator actuation. In addition, it is strongly influenced by the angle of attack.

A fully reusable sounding rocket is proposed and conceptually designed in ISAS/JAXA. In phase A in the proposed project, technical demonstrations for key technologies to develop the reusable sounding rocket are planed as follows; 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development and 6) health management system construction. In this paper, the present system design of the reusable sounding rocket and technical demonstrations are summarized.

Liquid hydrogen LH2 has excellent properties as a coolant, such as large latent heat, low viscosity coefficient, etc. Not only MgB2 but also other high-Tc superconductors are expected to have excellent properties when cooled by LH2. It is necessary for a stability design of a high-Tc superconductor cooled by LH2 to make an electro-magnetic characteristic clear. However, due to the handling difficulties of LH 2, there are only few papers on the properties of LH 2-cooled superconductors, especially under the external magnetic field. In this paper, an experimental set-up is described, which was designed and fabricated for the evaluation of the electro-magnetic characteristics of high-Tc superconductors cooled by LH2. The LH2 cryostat of 309 mm inner diameter was set co-axially with vacuum layer in the LHe cryostat in which the LHe-cooled superconducting magnet for external magnetic field (up to 7 T) was set. The LH2 cryostat has three power leads for feeding up to 500 A to the test high-Tc superconductors. © 2002-2011 IEEE.

Pressure drops in the non-isothermal sloshing of liquid nitrogen and liquid hydrogen were experimentally investigated. Heat and mass transfer indicated by the pressure change induced by violent sloshing in the cryostat driven by the mechanical exciter were successfully obtained. Based on the visualized image of liquid motion, it was found that splash and wavy surface induced by violent sloshing of liquid should enhance heat transfer around liquid surface and result in pressure drop. It was also suggested that the instability of singlecomponent system with different temperature between the gaseous and liquid phase could make the pressure control more difficult. © 2012 by Takehiro Himeno.

A new reusable sounding rocket is proposed (1) to reduce operation costs from that of existing expendable sounding rockets and (2) to provide frequent launch opportunities. This paper describes the current status of the conceptual design of the reusable sounding rocket, including system concept, mission profile and system configuration. Some of the R&D status is also presented. Copyright © (2012) by the International Astronautical Federation.

In ISAS/JAXA, a fully reusable sounding rocket is proposed as one step for the future full-fledged reusable transportation system. This vehicle is a vertical take-off and vertical landing (VTVL) rocket vehicle. And this has the capability of ballistic flight to the altitude over 120 km and returning to the launch site. In the flight sequence, the vehicle takes off vertically and cuts main engine off about 100 seconds later, and reaches to an altitude about 120 km during the ballistic flight. After that, the vehicle turns back into a nose first attitude. During the return flight, the vehicle is guided to above the launch place. Then the vehicle makes a turnover maneuver to base first attitude form a nose first entry attitude. This makes it possible to achieve the deceleration and soft landing by its main engine thrust. It is considered that there are many technical concerns to realize this vehicle. To show the feasibility of the vehicle, technical demonstrations are under way in JAXA. One of the technical concerns is a turnover maneuver during the return flight. As for the inversion maneuver, it is considered for applying the aerodynamic turnover maneuver which is caused by the differences of pitching moment depending on the vehicle configurations. For example, vehicle configurations are changed by deploying strake, a kind of canard. To verify the turnover maneuver capability, it is considered to demonstrate the glide tests using the small sized vehicle model. In this demonstration, the technical problems for the turnover maneuver as for the vehicle dynamics and guidance control strategy will be investigated. In this paper, the turnover maneuver control for the nose first entry of the small sized vehicle is numerically simulated and the return flight guidance to a point is proposed. And turnover maneuver dynamics of the vehicle is investigated for practical use. The guidance and control method proposed here would be applied for the full scale reusable sounding rocket vehicle. Copyright © (2012) by the International Astronautical Federation.

Asymmetric separation vortices over a slender body at a high angle of attack exert a strong side force on the body and lead to the loss of attitude stability. We investigated the active control of the separation flow over a slender body and addressed the proportional control of the side force and the pitching moment. A flow control experiment was conducted in a wind tunnel using a cone-cylinder test body and a Dielectric Barrier Discharge (DBD) plasma actuator as a flow control device. The free-stream velocity was 9 m/s and the Reynolds number was approximately 42000. The side force coefficient was proportionally controlled within approximately ±1.0 using the actuator at the aft body, and the static stability angle of attack was controlled from 25 to 40 degrees and 65 to 85 degrees by controlling the pitching moment when the center of gravity was at the 55% position from the body apex. We estimated that a higher actuator output power is required for the effective control of the aerodynamics in a real flight. In addition, we confirmed that the actuator burst operation mode could reduce the required output power.

The aerodynamic characteristics of a vertical landing rocket are affected by its engine plume in the landing phase. The influences of interaction of the engine plume with the freestream around the vehicle on the aerodynamic characteristics are studied experimentally aiming to realize safe landing of the vertical landing rocket. The aerodynamic forces and surface pressure distributions are measured using a scaled model of a reusable rocket vehicle in low-speed wind tunnels. The flow field around the vehicle model is visualized using the particle image velocimetry (PIV) method. Results show that the aerodynamic characteristics, such as the drag force and pitching moment, are strongly affected by the change in the base pressure distributions and reattachment of a separation flow around the vehicle.