Masao Takegoshi

Performance of a Rocket-Ramjet Combined-Cycle Engine Model in Ejector Mode Operation

A rocket-ramjet combined cycle engine model, embedding twin rocket chamber on top wall side of a scramjet flow-pass, was tested in its ejector-mode operation under sea-level static conditions. The rocket chamber was driven with gaseous hydrogen and oxygen at nominal operation condition of 3 MPa in chamber pressure and 6.5~7.5 in mixture ratio. Gaseous hydrogen was also injected through secondary injector orifices to pressurize the ramjet combustor. Mixing between the hot rocket plume and cold airflow as well as combustion of residual fuel within the plume with the airflow caused entropy and static pressure increases in the constant-area mixing duct in our original flow-pass design, resulting in a high back-pressure to the incoming airflow and a limited airflow rate. Thus, the mixing duct was re-designed to have divergence from its onset to compensate the pressure-rise. With this modified flow-pass configuration, the airflow rate was increased by 40%. However, this flow-pass geometry resulted in generation of low speed area, through which pressure-rise due to secondary combustion (and flow-pass exit contraction to simulate secondary combustion) penetrate the mixing duct and reduced the incoming airflow rate. A contraction on the incoming airflow enabled choking condition of the incoming airflow to sustaining the airflow rate, while the rate itself was reduced. Contraction at the exit of the engine enhanced mixing, however, choking condition was not attained due to the higher pressure level associated with the high exit contraction. Balancing these factors and additional mixing enhancement are required. The engine performances were summarized.

Performances of a Rocket Chamber for the Combined-cycle Engine at Various Conditions

A gaseous hydrogen/gaseous oxygen rocket chamber was designed to fit in a rocketramjet combined-cycle engine model, and its performance was evaluated experimentally. Such a rocket chamber is required to operate in very wide ranges of chamber pressure (Pc) and mixture ratio (O/F). For stable operation, the injector has a choking point and a diffuser in the downstream portion. The design point of the injector is Pc = 5.0 MPa and O/F = 7 when the injection pressure of both the fuel and the oxidizer is 7 MPa. Stable operation and a C-star efficiency of 0.91 were attained in the rocket mode operation at O/F = 6.5 - 7.5 and Pc = 3 - 5 MPa. Stable operation and a C-star efficiency of 0.93 were attained in the ramjet mode operation at O/F = 4.5 - 7 and Pc= 0.6. This stable operation was attained by supplying oxygen from two, three or four of eight injectors. The C-star efficiency was 0.94 with four oxygen injector elements at O/F = 0.89, and 0.92 with three oxygen injector elements at O/F = 0.49. No thermal damage was observed on the oxygen post and faceplate with flush face oxygen post in all operating conditions. The fundamental design of the rocket chamber and injector for the combined-cycle engine was completed in this study.

Firing-Tests of a Rocket Combustor for Combined Cycle Engine at various conditions

A gaseous hydrogen / gaseous oxygen rocket chamber was designed to fit to a RocketBased-Combined-Cycle engine model, and its performances were evaluated experimentally. The rocket chamber was required to operate at a very wide operation range in terms of chamber pressure (Pc) and mixture ratio (O/F); 0.6 MPa & 6 for ‘ramjet-mode’ operation, 0.6 MPa & 0.5 for ‘scramjet-mode’ operation, and 5 MPa & 7 for ‘ejector-rocket-mode’ operation. For stable operation, both gaseous hydrogen injectors and gaseous oxygen injectors, which were aligned co-axially, had choking point and diffuser at downstream portion. The number of the oxygen injector in use could be selected. The outer hydrogen injector showed lower discharge coefficient and lower durability against back-pressure than the inner oxygen injector. The hot-firing tests with a heat-sink type combustion chamber showed stable operation with the C-star efficiency of 87% for the ramjet-mode operation and 83% for the scramjet-mode operation. The hot-firing tests with a water-cooled combustion chamber also showed stable operation with the C-star efficiency of 95% for the ejector-rocket-mode operation. The water-cooled chamber showed enough durability in the ejector-rocket-mode operation, however, serious thermal damages upon the tip of injector elements and the chamber faceplate were observed, requiring more modification in the design around the injector and the faceplate.

Injectors and Combustion Performance of Rocket Thruster for Rocket-Ramjet Combined-Cycle Engine Model

The required operating conditions for the present rocket thrust chamber for rocketramjet combined-cycle engine are 1) chamber pressure Pc = 5.0 MPa, mixture ratio O/F = 7 for the ejector-jet mode, the scramjet mode, and the rocket mode, 2) Pc = 0.6 MPa, O/F = 3 for the ramjet mode. Stable operation of a rocket engine in a wide range of both Pc and O/F is required in order to adjust the operation mode in accordance with flight speed. Gaseous hydrogen and gaseous oxygen were used as the propellant. In the previous study, the hydrogen flow rate during firing tests decreased due to the thermal deformation of the faceplate. In this study, a new designed injector which has eight hydrogen injection holes arranged around an oxygen post was proposed for the rocket thrust chamber of the rocketramjet combined-cycle engine. The hydrogen flow rate during firing tests was almost the same as that during the cold flow tests. Performances of 0.84 to 0.88 in C* efficiency was achieved under the condition of O/F = 6.5 to 7.5 using the thrust chamber of L* = 0.33 m. However performances of about 1.0 in C* efficiency were achieved under the conditions of O/F = 4 to 6.5 using the thrust chamber of L* = 0.99 m.

Evaluation of Metallic-tube-cooled C/C Composite Structure by Rocket Combustor

In this study, the cooling performance of a C/C composite material structure with metallic tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The temperature of the C/C composite materials cooled by stainless steel tubes in the combustor attained a stable state at about 40 seconds after ignition. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2900 K and heat flux to the combustion chamber wall was about 6.5 MW/m 2 . No thermal damage was observed on the stainless steel tubes which were in contact with the C/C composite materials. The heat flux to the C/C composite wall was about 59% less than that of a water-cooled copper alloy combustor wall. The result shows that the amount of the engine coolant can be reduced. Results of the heating test showed that such a metallic-tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure, as well as indicating the possibility of reducing the amount of the coolant even if the thermal load to the engine is high. Thus, application of the metallic-tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined cycle engine is expected.