Kanenori Kato

Experimental Study on Combustion Modes in a Supersonic Combustor

obliquefuelinjectionwasemployedtomitigatefueljet/airflowinteraction,theinteractionwasfoundtohaveasizable effect on the ignition limit. A simple model to predict the minimum combustor length needed to attain supersonic or dual-mode combustion was suggested, and it agreed relatively well with the experimental results. The model was applied to experimental results of a subscale scramjet engine tested under Mach 8 flight conditions at the Japan Aerospace Exploration Agency, Kakuda Space Center, in order to predict the occurrence of combustion (large pressure rise) within a constant area combustor in the engine, and it agreed well with the experimental results.

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.

Downstream Ramjet-Mode Combustion in a Dual-Mode Scramjet Engine

Downstream combustion ramjet-mode operation in a dual-mode engine combustor was studied experimentally with a Mach 2.5 wind tunnel. In this downstream combustion ramjet mode, subsonic combustion was attained in the downstream straight-duct section without a geometrical throat. The experimental results are presented. The total temperature and the total pressure in the vitiation air heater were 800 K and 1.0 MPa, respectively. Pitot pressure and gas sampling were measured on the exit plane of the combustion model. The combustion conditions of the downstream combustion ramjet mode were compared with those of the usual upstream combustion ramjet mode. Better thrust performance and combustion were observed in the upstream combustion ramjet mode. However, in the case of the downstream combustion ramjet mode, injection of a larger amount of fuel was possible and a large impulse function was attained. The downstream straight duct was required for sufficient reaction. The fuel did not go far upstream in the separation region in pseudo-shock. Nomenclature ER = equivalence ratio M = Mach number Pi = wall pressure at the entrance of the combustor Pp = pitot pressure Pw = wall pressure x = distance from backward-facing step y = vertical distance from the bottom z = transverse distance from side wall surface ηc = combustion efficiency φ = equivalence ratio

Performance of a RBCC Combustor Operating in Ramjet Mode

*† ‡ § ¶ # The direct-connect combustion tests and the numerical simulations of the rocket-ramjet combined cycle combustor, requiring large bases to install rocket chambers, were conducted to determine whether the combustor could be operated in ramjet mode. The rocket engines in the combustor were operated at fuel rich condition as fuel injectors for the freestream. For the rocket operation alone, the combustor was operated in scramjet mode. By the secondary fuel injection from the cowl, the supersonic freestream was decelerated to subsonic speed by the combustion-generated shock train in front of the secondary injection. The subsonic combustion of the secondary injected fuel resulted in higher combustion and thrust performances. The further upstream secondary injection was suitable for the ramjet operation. By the increase in fuel flow rate of the secondary injection, the thermal load on both the rocket chamber and throat was reduced without decrease in thrust. Gas sampling measurements at the exit of the combustor revealed that the merger of the secondary fuel with the rocket plumes resulted in the decrease of the combustion and thrust performances.

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.

Skin-Friction Measurements in Supersonic Combustion Flows of a Scramjet Combustor

7m -1 , 3.3 MW, respectively. The measurements were conducted using a skin friction gage. This type device used a small movable head mounted flush to a wall such that the head was assumed to be exposed to the same shear stress from the flow as the surrounding wall. At first, wall skinfrictions were measured in the case with straight ducts without combustion, evaluating measurement accuracy which was found to be ±5%. Then the testing was conducted in a supersonic combustor to examine the effects of combustion on wall skin friction. The output from the gage during test time showed that the wall skin friction with supersonic combustion was as large as that for a corresponding non-combusting flow. It is difficult to explain with accuracy the changes due to combustion by such a method that combined analyzing the state of main flow and one dimensional skin friction prediction for flat plate. It is needed that establishing such more accurate prediction of skin friction that would account for effect of combustion in the hypersonic airbreathing engine.

Experimental Study of Combined Cycle Engine Combustor in Scramjet-Mode

A rocket based combined cycle engine was tested in scramjet-mode. The combustor model had two rockets in the combustor section. These rockets were used not only for thrust production unit, but also for fuel supply unit. The test facility consisted of an air supply section, a combustor section, a straight duct section, and a divergent duct section. Total pressure of the supplied air was 1 MPa, and the total temperature of the air was 2400 K. In the tests, in order examine the effect of the length of the straight duct section, four kinds of length of straight duct were connected downstream to the rockets. Two combustion patterns were observed for different lengths of straight duct. Pitot pressure and gas sampling were measured on the exit plane of the combustor model. Better combustion conditions were attained in the case of a longer duct. Nomenclature A = cross section Fc = impulse function of combustion gas L = length of straight duct section (including the straight part of combustor downstream of rocket) Pw = wall pressure Pwi = wall pressure at the entrance of the combustor W = width of the straight duct section x = distance from the position of the rocket nozzle exit along the airflow direction y = vertical distance from the bottom z = transverse distance from sidewall surface (rocket side) (O/F)r = mass flow ratio of oxygen to fuel in the rocket φ = equivalence ratio to airflow ηc = combustion efficiency

Mach-8 Tests of a Combined-Cycle Engine Combustor

Under the Mach-8 flight condition, combined-cycle-engine combustor tests were conducted. The combustor model had a two-rocket fuel injector, which supplied fuel-rich, precombustion gas as fuel. Gaseous hydrogen and oxygen were used as propellants. Two combustion patterns were observed for various lengths of the straight duct connected to the rocket injector section. In the longer straight-duct configurations, combustion gas was decelerated to subsonic speed and choked thermally at the exit of the straight duct. In the shorter straight-duct configurations, combustion progressed in the downstream divergent duct. Pitot pressure and gas-sampling parameters were measured at the exit plane of the combustor model. Results showed that a certain length of a constant cross-sectional area combustor section was required to sustain combustion of the rocket plume injected parallel to the airflow.

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.

Performance of a Rocket-Ramjet Combined-Cycle Engine Model under Ramjet-Mode Operations

A rocket-ramjet combined cycle engine model, embedding twin rocket chamber on top wall side of a scramjet flow-pass, was tested under ramjet- (dual-) mode operation at M6 flight conditions. The rocket chamber was driven with gaseous hydrogen and oxygen at nominal operation condition of 0.6 MPa in chamber pressure and 6.0 in mixture ratio. Gaseous hydrogen was also injected through injector orifices within the ‘ramjet’ combustor to pressurize the combustor. A simple, one-dimensional calculation was conducted based on wall pressure measurement to evaluate combustion mode and engine performance. Engine exhaust samplings as well as Pitot pressure measurements were conducted to evaluate the calculation results. Different combustion modes were observed, which were found to have sizable effects on the engine performance.