Fumiei Ono

Experimental studies of supersonic film cooling with shock wave interaction

The supersonic film cooling was tested in the Mach 2.35 wind tunnel to investigate the effect of the external shock wave on the film cooling. The coolant was injected with sonic speed. The weak shock wave with the pressure ratio of 1.21 did not reduce the film cooling effectiveness. The stronger shock wave with the pressure ratio of 1.44 decreased the effectiveness of the film cooling in the restricted region. The decrease of the effectiveness was mainly the result of the increase of the adiabatic wall temperature by the decrease of the local Mach number. The increase of the heat transfer coefficient must be considered as well as that of the adiabatic wall temperature. In the region of the interaction, energy and mass were not transferred, but the momentum was transferred from the primary flow to the coolant.

Mach 6 Test of a Scramjet Engine with Multi-Staged Fuel Injection

In this study, a multi-staged fuel injection was applied to a scramjet engine for improving the thrust performance without un-start transition under Mach 6 flight conditions. A boundary-layer bleeding was also applied for suppressing un-start transition. With the multi-staged fuel injection, the engine operated without un-start transition at the fuel equivalence ratio of unity or above. Especially, with fuel injection strut, the thrust increment from the no fuel condition reached to 2880 N at equivalence ratio of 1.45, which was about 1.5 times as large as the maximum thrust obtained with the single-stage injection in the previous tests in Mach 6 flight condition. It was confirmed that the multi-staged injection improved the thrust performance without un-start transition. Nomenclature Cf = friction coefficient Dint = internal drag of engines F = thrust ∆F = thrust increment from the no fuel condition ∆Fint = net thrust by combustion (∆F-Dint) Isp = specific impulse based on net thrust (∆Fint / mf) mf = fuel mass flow rate p = pressure Φ = equivalence ratio Subscripts

Ram and Ejector-Jet Modes Experiments of the Combined Cycle Engine in Mach 4 Flight Conditions

A combustion-capable combined cycle engine model which was constructed based on the rocket and ramjet technology was tested in Mach 4 flight condition. At this speed, engine is designed to shift its operation mode from an ejector-jet to a ramjet. Both modes were simulated by changing the rocket combustion pressure. Even with full rocket exhaust, no effect to the air flow could be observed. The injection point of the secondary fuel affected thrust performance. In the ramjet mode, the pressure rise due to the fuel combustion traveled to the entrance of the combustor, but it stayed near the injection point in the ejector-jet mode.

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.

Oxidation behavior of CVD-SiC in high temperature atmosphere

Heating tests of the silicon carbide thin plates made by chemical vapor deposition process (CVD-SiC) were conducted in the high temperature static air. Heating temperature was varied from 1673K to 1973K, heating period was varied from 1 minute to 40 hours, and ambient pressure was kept at 1 atm. The oxidation rate and the electric impedance of the CVD-SiC specimens were observed under these conditions. Although the oxidation rate was very high at the initial heating period, it rapidly decreased as the silica layer grew on the surface. The oxidation rate of the CVD-SiC increased as the heating temperature rose. When the CVD-SiC specimen contacted with the specimen support made by alumina, the silica layer locally melted at the temperature less than the normal melting point. In the melted area, the rapid oxidation of the CVD-SiC was confirmed. The variation of electrical resistivity of the specimen was significantly affected with the heating temperature and the cumulative heating time.

Effects of propellant characteristics and vehicle structure on rocket space plane

The performance of the space transportation system is usually expressed in terms of the mass ratio of the vehicle and the specific impulse of the propulsion system. Authors introduced a new parameter to show the structural performance of the space transportation system. It is a vehicle bulk dry density defined as a ratio of the vehicle dry mass to the propellant tank volume. A modified vehicle bulk dry density was also introduced. The payload mass is added to evaluate the vehicle dry mass in the latter ‘definition. The effects of the propellant density, the specific impulse and the modified vehicle bulk dry density on the mission capability were discussed in this study.. From those analysis, it was shown that the LOX/LCH, propulsion system allowed the maximum modified vehicle bulk dry density among the practical propellants. By the introduction of the concept of vehicle bulk dry density, it was derived that this propulsion system would accept the most compact and heavy body for a rocket SST0 vehicle. NOMENCLATURE Z SP specific impulse m, initial mass of vehicle (gross lift off mass) mf final mass of vehicle (burn out mass) m,, dq mass of vehicle mpr mass of payload mpmp mass of propellant O/F mixture ratio of oxidizer flow to fuel flow a coefficient appeared in eq.( 12) to (14) /3 constant appeared in eq.(l2) to (14) p. oxidizer density, pF fuel density. *Head, Ramjet Structure Section, Ramjet Propulsion Research Division, Kakuda Research Center, Member AL4A. ’ Senior Researcher, Ramjet Propulsion Division, Kakuda Research Center. 0 ‘Head, Ramjet Systems Section, Ramjet Propulsion Research Division, Kakuda Research Center, Senior member AIAA. ’ Senior Researcher, Ramjet Propulsion Research Division, Kakuda Research Center, Member AIAA. Copyright

Testing of regeneratively cooled light-weight panel

Two types of actively cooled light-weight panels for use in scramjet engines were fabricated. They were made of Hastelloy X, and the photo-etching method was applied for constructing many cooling passages in these panels. One type of panels had a honeycomb structure which was brazed at the back of the panel for reinforcement. A vitiated-airflow generator and a device for supplying coolant were applied to evaluate the cooling ability of the panels. The vitiated airflow had a total temperature of 2000 K 3000 K, a total pressure of 1.0 MPa 1.5 MPa, and a Mach number of 3.1 3.4. The coolant used in this series of tests was water. No damage was observed in these cooled panels after testing. We measured the heated-surface temperature of the panels using a radiation thermometer, and strain at the center of the rear surface of the panel was determined with strain gages. The radiation thermometer indicated a temperature of about 500 K 780 K on the heated-side surface depending on the heat flux. The strain measured during the test was mainly due to thermal stress. The honeycomb structure was able to reduce the apparent strain caused by thermal stress. This indicates an incrcase of rigidity in the panel. *Researcher, Ramjet Structure Section, Ramjet Propulsion ?Senior Researcher, Ramjet Propulsion Research Division. Research Division.