Kouichiro Tani

COMPARISON OF SCRAMJET ENGINE PERFORMANCE IN MACH 6 VITIATED AND STORAGE-HEATED AIR

To investigate the sensitivity of combustion to the test gas, an H2-fueled scramjet engine was tested at a Mach 6 x8f ight condition with the air supplied by a combustion heater (V mode) and a storage heater (S mode). The fuel self-ignited without the assistance of igniters in the V mode. However, self-ignition was difx8e cult in the S mode. The easier ignition with vitiated air was caused by radicals supplied from the combustion heater. The combustion behavior was also affected by the test air, which suggests that the combustion was not fully mixing-controlled. As the fuel x8f ow rate increased, the combustion changed from a weak mode, delivering a lower thrust, to an intensive mode, with a higher thrust. Gas sampling showed that the weak combustion was caused by autoignition in the boundary layer on the engine walls. In the intensive combustion mode, the x8f ame was anchored near the backward-facing step on the sidewalls. However, the x8f ame partially detached from the step on the top wall in the combustor. The detached x8f ame may make the combustion kinetically controlled to produce the sensitivity to the test air.

Effectiveness of plasma torches for ignition and flameholding in scramjet

A newly developed plasma torch with a feed stock of air or oxygen was investigated experimentally in order to determine its effectiveness on ignition and flameholding in a scramjet combustor. This design comes from the viewpoint of total system design of scramjet engine and vehicle because it is preferable to utilize incoming air or onboard propellants as a feed stock. Three patterns of fuel injection were tested 1) from one orifice: 2) from four orifices on one wall; and 3) from all nine orifices on both walls. Ignition and flameholding phenomena were examined through direct photographs of internal and exit flames of the combustor and by wall temperature measurements. The specially devised plasma torch with air or oxygen was able to operate stably without any support gas, for example, argon. Ignition limit curves, with and without the plasma torch, were obtained on a plane relating the air total temperature to the fuel equivalence ratio for the three patterns of fuel injection, and then compared to each other. For a wide range of experimental conditions, it was shown that the effectiveness of an air or oxygen plasma torch was comparable to that of a nitrogen or argon-hydrogen plasma torch. For single-wall injection, it was observed that the plasma torch ignited the fuel jet located directly downstream, and the flame thus formed ignited adjacent fuel jets. In double-wall injection, however, ignition of the fuel injected from the wall opposite the plasma torch was unsuccessful. It was also found that, under some conditions, flameholding can be continued even after the plasma torch is turned off, most notably in the case of single-wall injection. The occurrence or nonoccurrence of this phenomenon is also shown in the ignition limit curves diagram.

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.

Effects of injector geometry on scramjet combustor performance

An experiment was conducted to investigate the effect of injector/combustor geometry on combustion-induced peak wall pressure and associated upstream influence, as well as on mixing/combustion characteristics at an entrance Mach number of 2.5. The length of the constant area section downstream of injection orifices had a strong influence on the above-mentioned characteristics. However, the sweep of the rearward-facing steps on both side walls had little effect on these characteristics, nor did reversing them have any effect. The peak wall pressure and the length of the upstream influence agreed qualitatively with predictions of an analytical model and an empirical formula developed at Johns Hopkins University. Fuel jets injected from the model with the longest constant area section and the fuel equivalence ratio of unity, coalesced at a very early stage downstream of the fuel injection orifices. This coalescence led to a decrease in mixing rate downstream, despite the higher degree of mixing near the injection orifices. The combustion efficiencies were higher than those obtained at NASA Langley in the upstream region due to the higher mixing rate near the injection orifices.

Conceptual Study of a Rocket-Ramjet Combined-Cycle Engine for an Aerospace Plane

Operating conditions of a rocket-ramjet combined-cycle engine for a single-stage-to-orbit aerospace plane were studied. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. Characteristics of the engine operating conditions were studied analytically. The thrust augmentation effect of the ejector-jet mode was found to be small at low subsonic speed and to increase with an increase of the flight Mach number. Study of the effective impulse function clarified that higher specific impulse was preferable in supersonic flight, whereas greater thrust coefficient was preferable in hypersonic flight. The mentioned characteristics were examined by simulation of engine operating in an aerospace plane flight. Transportation of a mass into orbit was compared among several engines with different combinations of thrust and specific impulse. The mass which could be carried into orbit was larger with a ramjet mode of higher specific impulse and with a scramjet mode of greater thrust.

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.

Experimental study on autoignition in a scramjet combustor

A UTOIGNITION characteristics of hydrogen fuel in a scramjet combustor were examined using a direct-connect test apparatus with particular reference to the effect of fuel injection patterns. Autoignition behavior fell into four distinct categories, separated by the three bounding curves. One of the boundaries was independent of fuel injection patterns, while the others significantly depended on them. In the case of fuel injection from a single wall or from a single orifice, local flame quenching caused by expansion wave emanating from the step on the opposite wall was observed. Compared with injection from a single orifice, injection from multiple orifices appreciably enhanced autoignition. The reason for this is that fuel jets from adjacent orifices and from the opposite wall tended to attenuate the local flame quenching caused by the expansion waves from the opposite wall. Ignition limit curves derived from the present experiment were compared with an autoignition criteria proposed by Huber et al. They agree well in the case of fuel injection from a single orifice. For the case of injection from multiple orifices, however, the agreement is poor.

Aerodynamic Characteristics of the Combined Cycle Engine in an Ejector Jet Mode

Aerodynamic performances of the combined cycle engine model was tested in a subsonic to transonic flow region. In this low speed, the engine works in an ejector jet mode. To simulate ejector eect, gas nitrogen was exhausted from the nozzle which located downstream of the inlet. Wall pressure distributions and Pitot pressure at the exit of the model were measured to grasp basic features of the engine flow field. The fluctuating Pitot pressure was also measured to estimate the mass flow oscillation. Two dimensional CFD was carried out to compare the experimental results and the flow structure inside the inlet was examined. With the current configuration, high pressure region on the ramp which caused the increment of drag, was formed and well predicted by CFD.

Mach 6 Test of a Scramjet Engine with Multi-Staged Fuel Injection

In this study, a multi-staged fuel injection was applied to a scramjet engine for improving the thrust performance without un-start transition under Mach 6 flight conditions. A boundary-layer bleeding was also applied for suppressing un-start transition. With the multi-staged fuel injection, the engine operated without un-start transition at the fuel equivalence ratio of unity or above. Especially, with fuel injection strut, the thrust increment from the no fuel condition reached to 2880 N at equivalence ratio of 1.45, which was about 1.5 times as large as the maximum thrust obtained with the single-stage injection in the previous tests in Mach 6 flight condition. It was confirmed that the multi-staged injection improved the thrust performance without un-start transition. Nomenclature Cf = friction coefficient Dint = internal drag of engines F = thrust ∆F = thrust increment from the no fuel condition ∆Fint = net thrust by combustion (∆F-Dint) Isp = specific impulse based on net thrust (∆Fint / mf) mf = fuel mass flow rate p = pressure Φ = equivalence ratio Subscripts

Aerodynamic Performances of a Combined Cycle Inlet

Ramp-compression-type inlets for the combined cycle engine were investigated in a Mach 4 wind tunnel. Geometries of the inlet models were changed to investigate their effects on the aerodynamic performances and the starting characteristics. A bent cowl improved the starting characteristic by weakening the shock from the cowl leading edge to suppress the shock-induced separation and by reducing the internal contraction ratio of the cowl duct, thus, avoiding choking at the inlet throat. With the current geometries, the inlet started with a slightly larger internal contraction ratio than that restricted by the Kantrowitz-Donaldson limit. The ratio of the height of the cowl duct of the inlet to the incoming boundary layer was found to affect the starting characteristic. The model with a higher capture ratio showed better total pressure recovery performance.