Tohru Mitani

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.

Reaction and Mixing-Controlled Combustion in Scramjet Engines

H2-fueled scramjet engines were tested under Mach 4 (M4) to Mach 8 (M8) e ight conditions, and the local equivalenceratio andcombustion efe ciency weremeasured by gassampling attheengineexit.Correlationbetween the local values of equivalence ratio and combustion efe ciency showed that the M4 combustion was principally reaction controlled and the reaction and the mixing-controlled combustion coexisted in the M6 condition. The M8 combustion in the engines was rate controlled by the mixing of H 2. Comparison of engine performance in the air, supplied by a combustion heater and a storage heater, indicated that the performance strongly depended on the air-heating methods. The dependence of engine performance on air-heating methods could be explained by the e nding that the M6 engine combustion was partially reaction controlled. The wall-heating rate and pressure distribution in the M6 tests also supported the shift from the partially reaction-controlled to the mixing-controlled combustion as the fuel rate was increased. The mixing-controlled combustion suggested weak facility dependence intheM8 condition.Becausereaction in the M8 condition is notsufe ciently fast, the main combustion region might be blown downstream in the engine.

Load Oscillations Caused by Unstart of Hypersonic Wind Tunnels and Engines

Large-amplitude load oscillations were observed during the tests of a hypersonic engine model in a freejet-type wind tunnel. To clarify the cause of the oscillations and their characteristics, oscillating wall pressures and loads on a drag model and engine models were investigated. Flowe eld was observed by shadowgraph to determine the cause of the large starting loads. Power spectral density functions and probability functions of wall pressures and loads were derived by the fast Fourier transform. The amplitude of the unsteady frontal pressure was correlated with the dynamic pressure. The magnitude of the starting load was related to the drag coefe cient of the models, and the expected maximum peak loads of a large-scale ramjet engine test facility were evaluated. Engine unstart loads were also simulated by means of secondary e ow injection into a small-scale model of a ramjet engine. With these methods, characteristics of engine unstart loads and the possibility of sensing engine unstart in its early phase were studied. Engine unstart could be sensed with pressure measurement around the engine throat before it became severe. Furthermore, engine unstart loads associated with scramjet engine combustion were related to the drag coefe cient of the engine.

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.

Mixing Augmentation of Transverse Injection in Scramjet Combustor

A method for augmenting the mixing of transverse injection in a scramjet combustor is suggested, and its effectiveness is checked by numerical methods. The intent was to promote streamwise vorticity by modifying the injector geometry using a cavity to increase the mixing rate and penetration. A three-dimensional Navier ‐ Stokes code adopting the upwind method of Edwards’ s low diffusion e ux splitting scheme was used. The k‐! SST turbulence model was used to calculate the turbulent viscosity. The injector models proposed showed remarkable increases of streamwise vorticity and, as a result, signie cant improvements in mixing characteristics such as mixing rate and penetration. However, the proposed models also showed additional losses of stagnation pressure. Themixing characteristicsarestrongly related to the jet-to-cross e ow momentum e ux ratio J. In the case of higher values of J, slower mixing rates, higher penetration, and more losses of stagnation pressure are shown.

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.

Drags in Scramjet Engine Testing: Experimental and Computational Fluid Dynamics Studies

No-fuel internal drag of a side-compression scramjet engine was evaluated by using two one-e fth-subscaled models, one forwall pressure measurement and theother forforce measurement under conditionsof a e ight Mach number of 4. The pressure and frictional drags in various parts of the models were estimated from these windtunnel tests. Comparison between the pressure measurement and the force measurement revealed that the drag derived by these wind-tunnel tests agreed within 5%. After examining the consistency between the pressure and the force experiments, these results were used to calibrate a newly developed computational e uid dynamics code. The frictional drag and the heating rate on the engine internal walls were evaluated with the unstructured-grid code to be compared with those obtained from the one-e fth-subscale model and the full-scale engine. The total drag coefe cient of the scramjet engine, including the installation drag, was found to be 0.281 and the internal drag coefe cient was found to be 0.093. Consequently, two-thirds of the total drag measured in engine testing in the Ramjet Engine Test Facility was produced by the external e ow over the engine module. Subtracting the external drag, the internal performance delivered by the H 2-fueled scramjet engine is discussed.