野中 聡

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.

The third-generation vehicle RVT#3 equipped with a pressure-fed engine, which had upgraded in terms of durability enhancement and a LH₂ tank of composite material, successfully performed in repeated flight operation tests; and the vehicle reached its maximum flying altitude of 42m in October 2003. The next step for demonstrating entire sequence of full-scale operation is to put a turbopump-fed system into propulsion system. From a result of primary system analysis, we decided to build an expander-cycle engine by diverting a pair of turbopumps, which had built for another research program, to the present study. A combustion chamber with long cylindrical portion adapted to the engine cycle was also newly made. Two captive firing tests have been conducted with two different thrust control methods, following the component tests of combustor and turbopumps separately conducted. A considerable technical issues recognized in the tests were the robustness enhancement of shaft seal design, the adjustment of shaft stiffness, and start-up operation adapted to the specific engine system. Experimental study of GOX/GH₂ RCS thrusters have also been started as a part of a conceptual study of the integration of the propulsion system associated with simplification and reliability improvement of the vehicle system.

資料番号: SA0200020000