Tetsuo Hiraiwa
(JAXA, Japan)

Combination of a air-breathing engines and rocket engines (RBCC, Rocket Based Combined Cycle engine) gives an opportunity to reduce onboard oxygen consumption and to increase system weight margins for various application from cruiser to launch vehicles. For launch vehicle applications, combination of a ramjet/scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines was proposed (termed as Rocket-Ramjet Combined Cycle engine), and related R&D activities are undergoing at Japan Aerospace Exploration Agency, Kakuda Space Center. This presentation is to address the current status of the system analysis on this engine and vehicle.

Sadatake Tomioka
(JAXA, Japan)

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket Based Combined Cycle engine) are expected to be the most effective propulsion system for Booster stage of space launch vehicles. At hypersonic regime, it will be operated at rather high rocket engine output for final acceleration with some Isp gains due to air-breathing effects. In this regime, attaining thrust at this high-speed regime becomes very difficult, so that parallel injection of the fuel for scramjet combustion is favorable as the momentum of the injection can contribute to the thrust production. Thus, embedded rocket chamber was supposed to the operated as fuel rich gas generator at very high output. This configuration was tested at simulated flight Mach number of 7~11 at High Enthalpy Shock Tunnel (HIEST) with detonation tube as the source of the simulated rocket exhaust. However, combustion of the residual fuel in the rocket exhaust could not be attained. To understand the physics and to evaluate the combustion enhancement technique (auxiliary injection of fuel), direct-connect combustor tests were performed. Results of both the engine model tests at HIEST and the direct-connect tests are summarized and presented.

Sadatake Tomioka
(JAXA, Japan)

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket Based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO system or TSTO system with separation at high altitude needs final stage acceleration in space, so that the RBCC engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass's opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, and the results will also be presented.

Tetsuo Hiraiwa
(JAXA, Japan)

Combination of a air-breathing engines and rocket engines (termed as Rocket Based Combined Cycle engine, RBCC) gives an opportunity to reduce onboard oxygen consumption and to increase system weight margins for various application from cruiser to launch vehicles. For launch vehicle applications, combination of a ramjet/scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines was proposed (termed as Rocket-Ramjet Combined Cycle engine), and related R&D activities are undergoing at Japan Aerospace Exploration Agency, Kakuda Space Center. This presentation is to address the current status of the activities and R&D plans in near future.

Kiyoshi Nojima
(Tohoku University, Japan)

We experimentally investigated the mixing characteristics of the secondary fuel in a rocket-ramjet combined cycle engine operating under a sea-level static condition, to improve the combustion efficiency of the secondary fuel. We used a 1/5-scale engine model, which had two rocket chambers in a compression ramp on the topwall. The rocket exhaust was simulated by cold N2 gas at the chamber pressure of 3 MPa. The model had diverging section and constant section in the secondary combustor. Helium gas, simulating the secondary fuel, was injected in the diverging section. Gas sampling measurement through the wall pressure taps and planar Mie scattering images of the particles seeded only in the secondary fuel (Fig. 1) indicates that not only the fuel jets at the topwall side but those at the cowl side could ignite from the viewpoint of mixing, because the fuel jets at the cowl side sufficiently contacted with the rocket plumes. At the cowl side, the fuel jets penetrated deeply, covered the considerable flow-path area and mixed sufficiently with breathed air to achieve high combustion efficiency. Therefore, we believe that low flame holding performance of the engine induces low combustion efficiency of the secondary fuel.