Toshihito Saito

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 8 Testing of a Scramjet Engine with Ramp Compression

To improve combustion efficiency and the inlet started condition, a scramjet model having a ramp combined with a single strut was tested under Mach 8 flight conditions at the Ramjet Engine Test Facility of the National Aerospace Laboratory, Japan. The attached ramp shielded some of the fuel injectors. The fuel flow rate from the open injectors to the flow path was designated as the effective fuel flow rate. In the tests, a combustion efficiency of 90% was attained with vertical injection of hydrogen fuel. The thrust increase was 590 N at the effective equivalence ratio of 1.3. However, because the engine geometry was not optimized, a sufficient increase of the thrust was not attained. High temperature and high pressure necessary for ignition and combustion were achieved by the ramp. When the pressure in the isolator was 160 times as high as that of the freestream air due to combustion, the inlet was in the started condition despite the high pressure. This improved started condition was attained by using the ramp to increase the pressure in the inlet.

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.

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.

R&D on Hydrocarbon-fueled RBCC Engines for a TSTO Launch Vehicle

Combination of a air-breathing engines and rocket engines (termed as 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, T. Kanda, et. al., JPP, 19(2003), 859), and related R&D activities are undergoing at Japan Aerospace Exploration Agency (JAXA), Kakuda Space Center targeting hydrogen as the fuel due to its high Isp and cooling performances. Use of hydrogen fuel, on the other hand, can be costly due to the fuel cost and difficult-to-handle nature of liquefied hydrogen. Thus, hydrocarbon fuel such as ethanol was under consideration for the application to the future reusable launch vehicles termed as the ‘reference system,’ targeting manned operation to LEO. This presentation is to show the status of system analysis, flight demonstration plan, and related research efforts.

Experiments on scramjet engine with ramp-compression inlet at Mach 8

A recent modification to our scramjet engine and its experimental achievements under Mach 8 flight condition is presented. This modified engine has a ramp block, which covers its entire top wall, instead of the strut of the original engine. This block suppressed the inlet-combustor interaction under a large adverse pressure gradient and enlarged the inlet starting condition, compared with conditions of the engine with the strut-compression system. With this wider starting condition, the modified model produced positive net thrust, namely, larger thrust than its own drag. This engine did not go into the unstarted condition nor lose its thrust abruptly, because the airflow on the inlet ramp block separated gradually from its downstream end with an increase of the combustor pressure.

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.

EFFECT OF INGESTED BOUNDARY LAYER ON SCRAMJET ENGINE'S THRUST AND COMBUSTION CHARACTERISTICS

National Aerospace Lab. of Japan has been testing sidewall compression-type scramjet engines at Mach 8 flight condition since 1998. Like typical scramjet engines, our engines also ingest a thick boundary layer developed along facility nozzle surface to simulate an air flow that these engines on the vehicles will take in. Although this layer supports fuel-mixing and - combustion in the engines, it is also an origin of flow separation at the inlet section and engine-unstarting. Therefore, controlling this boundary layer is one of the essential techniques for the scramjet engines. With two engines that have a strut or a rampblock, we experimentally inspected this boundary layer effect on their performances, i.e., thrust, Isp, and their inlets starting/unstarting. Engines are tested in two different types of incoming air flows with a fully-developed boundary layer, with an inviscid region of the layer. Based on these experimental data, we discuss the effect of the ingested boundary layer and evaluate the contribution of the boundary layer to engines total performance. The engine with a strut affects strongly the boundary layer in its limit of starting and thrust performance, but the engine with a rampblock does not. This difference is caused by the upstream influence distance of pressure distribution, which is a function of the ingested boundary layer.

Research works of ethanol propulsion system for the future rocket-plane experimental vehicle

Experimental research works for the ethanol propulsion system has been performed in JAXA/Space Transportation Mission Directorate (STMD) as a part of study for future reusable launch vehicle.. Under this program, 1000 kgf-class (2200 lbf) rocket engine testing, material-matching and critical heat flux (CHF) studies and other related works were conducted in 2009 to 2010. This paper provides an overview of this engine experimental program and the results.