Kan Kobayashi

Supersonic Flow Ignition by Plasma Torch and H2/O2 Torch

The effectiveness of an H 2 /O 2 torch igniter, which injects H 2 /O 2 combustion gas, and a plasma torch igniter in promoting ignition of an H 2 /air mixture within a supersonic combustor was investigated analytically and experimentally. First, bulk temperature and bulk radical concentration in an O 2 plasma torch igniter were estimated by asymptotic analysis. Those in the H 2 /O 2 torch igniter were also obtained by equilibrium calculation. Second, gas sampling was conducted to determine the concentration of igniter injectant and the equivalence ratio in the region of ignition, namely, the recirculation region at the base of a rearward-facing step. Then, ignition time of the mixture in the region was analytically estimated by using a two-step H 2 /O 2 reaction mechanism and the results just-mentioned. The estimation showed that the ignition promotion effect of the plasma torch igniter was higher than that of the H 2 /O 2 torch igniter at a given input energy. However, the H 2 /O 2 torch igniter was also expected to sufficiently promote ignition because it was easy to increase the input energy by increasing the mass flow rate. Thus, an ignition experiment with the H 2 /O 2 torch igniter was conducted in Mach 2.5 airflow. Comparison of the results with those of the plasma torch igniters showed that the H 2 /O 2 torch igniter also had a sufficient ignition promotion effect as expected from analytical results.

Effects of injection configuration on performance of a staged supersonic combustor

Performance of a staged supersonic combustor with two-staged wall injections was examined experimentally in a direct-connected wind-tunnel facility with a combustion heater, which supplied Mach 2.5 airflow with a total temperature of 1500 K. The fuel flow rate through the first-stage wall injectors in the combustors minimum cross-sectional area section was limited to avoid combustor-inlet interaction, and the second-stage injection in the following divergent section was used to increase the pressure in this section. The performance was compared with that of the combustor with first-stage strut/second-stage wall injection configuration, while the strut was kept installed in both cases. The first-stage wall injection at a fuel flow rate caused a pressure rise almost equivalent to that of the first-stage strut injection, whereas it resulted in the combustor-inlet interaction at a lower fuel flow rate. Staged injection with the first-stage wall injection resulted in lower combustion efficiencies of the second-stage fuel mainly because of poor jet penetration, poor fine-scale mixing, and slower reaction and resulted in poorer combustor performances than those with the first-stage strut injection. Installation of the strut was found to have little effect on the performance of the combustor with the two-staged wall injections.

Boundary-Layer Control in Mach 4 and Mach 6 Scramjet Engines

Boundary-layer ingestion in airframe-integrated scramjet engines causes engine stall (engine unstart hereafter) and restricts engine performance. To improve the unstart characteristics in engines, boundary-layer bleed anda two-staged injection of fuel were examined in Mach 4 and Mach 6 engines. The bleed system consisting of a porous plate, an air cooler, a metering orifice, and an on/off valve was designed for each of the engines. First, a method to determine bleed rates was developed. Porous plates were designed to extract air from the Mach 4 engine at a rate of 200-g/s and from the Mach 6 engine at a rate of 30 g/s. Air coolers were then optimized based on the bleed rates. The exhaust air could be cooled below 600 K in the porous plates and the compact air coolers. The Mach 4 engine tests showed that the 200-g/s bleed (3% of the captured air) doubled an engine operating range (limit fuel rate) and thrust. With two-staged injection of H 2 , the engine operating range was extended from Φ0.3 to Φ0.95, and the maximum thrust was increased from 1200 to 2560 N. The Mach 6 tests showed that the 30-g/s bleed (0.6% of the captured air) extended the start limit from Φ0.48 to Φ1 and increased the maximum thrust from 1620 to 2460 N.

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.

Distributed Fuel Injection for Performance Improvement of Staged Supersonic Combustor

Introduction S UBSCALE model scramjet engines have been tested at Japan Aerospace Exploration Agency—Kakuda Space Propulsion Center (formerly National Aerospace Laboratory—Kakuda Space Propulsion Laboratory).1 The models consisted of a sidecompression-type inlet, a constant (minimum) cross-sectional area isolator with steps at its exit, a constant cross-sectional area combustor with fuel injectors, a diverging combustor, and an internal nozzle. In the tests, occurrence of combustor-inlet interaction2 resulted in a limit to the fuel flow rate and the engine thrust. To mitigate this interaction, a staged combustor, which had firststage fuel injectors on a strut in the minimum cross-sectional area

Vitiation Effects on Scramjet Engine Performance in Mach 6 Flight Conditions

Quasi-one-dimensional analyses with either chemical equilibrium or finite-rate reaction were conducted to evaluate the effects of contamination due to flow vitiation on engine performances, namely, thrust production. The incoming flow state was calculated based on measurements, and a higher total enthalpy for the freestream through a vitiation heater compared to that through a storage heater was found, the difference in the flow condition being responsible for two-thirds of the reduction in the thrust production with vitiation observed in the engine tests. With the finite-rate reaction calculation, the possibility of a further reduction due to the contamination effect on the combustion mechanism was found, which was responsible for one-third of the measured reduction. The one-dimensional analyses were further pursued to find matched test conditions with the storage heater to that with the vitiation heater in view of the thrust production and pressure distribution within the engine, and both freestream enthalpy and fuel equivalence ratio should be adjusted to attain the matched conditions.

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.

On acoustic damping of a cylindrical chamber in resonant modes

Acoustic damping of a cylindrical chamber with open and closed ends in resonant modes is analytically and numerically investigated to understand the low damping characteristic of the chamber without chocked nozzle. First, on the basis of the analytic solution of resonant acoustic modes inside a cylinder, the damping by radiation from the open end is calculated analytically using simple acoustic source modeling for velocity fluctuation. The effect of viscosity is also considered as an attenuation mechanism. The values of acoustic damping calculated for the first longitudinal and tangential modes are in good agreement with the corresponding values obtained using numerical simulation. The damping is also investigated for a configuration of the chamber with an injector installed off-center. Finally, we numerically and semi-analytically investigate the acoustic damping for a configuration that includes a hot-gas injection. The obtained mode is found to be a spinning tangential mode and the radiated wave also has a spinning feature. The damping for the spinning tangential mode is found to be larger than that for the symmetric dipole-like radiation under a uniform standing condition, but much smaller than the chamber with a chocked nozzle. Therefore, the chamber with an open end has the low damping characteristic suitable for intentionally generating oscillatory combustion.