Sadatake Tomioka

Mach 6 Testing of a Scramjet Engine Model

Testing of a subscale scramjet research engine model was carried out in the Mach 6 Ramjet Engine Test Facility of the National Aerospace Laboratory, Kakuda Research Center. The engine had a sidewall compression-type inlet. The fuel was hydrogen. With the attachment of a short strut on the top wall, intensive combustion with high-combustion efe ciency was attained, and the engine-produced thrust canceled the drag. The e ame was held in the low-velocity region around the step, even after the ignitors had been turned off. When the fuel e ow rate was small, there was a different combustion mode with weak combustion and little thrust. The unstart condition seemed to begin around the cowl. Tangential injection of fuel inhibited intensive combustion.

Sonic injection from diamond-shaped orifices into a supersonic crossflow

The plume from a diamond-shaped, sonic injector orifice was studied experimentally in a Mach 3 crossflow. The structure of the plume, as well as the near injector flowfield, were examined by flow visualization techniques,and penetration height growth and maximum concentration decay were evaluated from aerothermodynamic probing measurements at two downstream stations. For the transverse injection, the jet-to-freestream dynamic pressure ratio was varied from 0.3 to 2.0. At lower dynamic pressure ratios, the plume from the diamond-shaped injector penetrated farther across the main flow compared to that from an equivalent circular injector, whereas an increase in the dynamic pressure ratio resulted in a deterioration of the plumes sharpness, and penetration of the plume became comparable to that from the circular injector. Angled injections were applied to the diamond-shaped orifice to enhance the penetration at a high dynamic pressure ratio of 2.0. Giving sweepback angle to the orifice was as effective as the case with circular injectors in enhancing the penetration. Adding a moderate yaw angle to the sweptback, diamond-shaped orifice resulted in greatly enhanced penetration, unlike the case with the circular injector. In all cases, the decay rate of the maximum concentration was almost insensitive to the orifice shape.

Analyses and Application of Gas Sampling to Scramjet Engine Testing

Gas sampling has been used in combustor studies and in scramjet engine testing. Because the gas sampling is based on the assumption that the gas composition is frozen in the sampling process, the critical Damkohler numbers necessary to quench reactions in the gas-sampling probes were evaluated using a reduced kinetic model. The phase plane analysis showed that reactions in probes can be extinguished if the probe Damkohler number is less than about 10. The analytical results were cone rmed by numerical calculations using full kinetics. The shock swallowing into sampling probes was examined using numerical simulations for the low-Reynolds-number e ow. These theoretical results were verie ed by experiments using four kinds of probes with various cone gurations in a Mach 2.5 supersonic combustor. Based on the results, e ne sampling probes with a tip diameter less than 0.3 mm are recommended for scramjet testing. Based on these calibration studies, gas sampling was successfully applied to scramjet engine testing under a e ight Mach number up to 8, to reveal interesting features in the internal e ow in swept-back engines.

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.

Mechanism and Control of Combustion-Mode Transition in a Scramjet Engine

A sidewall compression scramjet engine operated in two combustion modes underMach 6 flight condition, weakand intensive-combustionmodes.Theweakmode occurredbelow the overall fuel equivalence ratio ( ) of around0.4. Transition from the weak mode to the intensive mode occurred at 0:4, accompanied by a sudden increase in thrust. Mechanisms of the transition were numerically investigated in this study. Simulations captured the sudden increase in thrust at the mode transition. In the weak mode, combustion occurred in only a region near the topwall where an igniter was installed. The combustion region expanded toward the cowl with boundary-layer separation at the mode transition. Simulations demonstrated that low ignition capability resulted in the weak mode. This study demonstrated that the presence of additional igniters on the sidewalls improved the ignition capability and achieved the intensive mode in the entire range.

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.

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.

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.