Atsuo Murakami

Ignition and Combustion Performance of Scramjet Combustors with Fuel Injection Struts

Ignition and combustion performance of a scramjet combustor with a fuel injection strut was experimentally investigated with Mach 2.5 vitiated air. Five strut models with different leading-edge geometry were tested without fuel injection to select the less flow-disturbing configuration. The nonreacting flowfields were also investigated by computation with a two-dimensional Navier—Stokes code. Using the selected strut, combustion and ignition tests were conducted. A pitot pressure and gas composition survey was carried out to deduce mixing and combustion efficiencies. It was found that mixing and combustion with a less flow-disturbing strut was considerably worse than those with a more flow-disturbing strut. Autoignition and forced ignition with plasma torches were tested for hydrogen. Ignition characteristics of parallel and perpendicular injection were quite different. The plasma igniters could successfully ignite both parallel and perpendicular fuel jets without a noticeable time delay between both sides of the strut.

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.

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.

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.

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.

Mixing and Combustion Control Strategies For Efficient Scramjet Operation in Wide Range of Flight Mach Numbers

In this paper, we present our main results of the firing tests of a newly proposed hydrogen fueled dual-mode scramjet combustor. The present scramjet combustor aims at obtaining a better engine performance, working characteristics and operability in wide range of flight Mach numbers. To realize such a scramjet, we especially focus on the following technical issues 1) use of parallel/low angle fuel injection, 2) control of combustor boundary layer separation, 3) good ignition/flameholding ability as well as efficient fuel/air mixing and combustion in the supersonic core flow by the streamwise vortices, and 4) selective operability of supersonic/subsonic combustion modes and efficient combustor operation in these modes. Involving these technical issues, our basic idea for the combustor of such efficient scramjet performance and operability is a multiple staged combustor characterized by the combination use of the “Alternating-Wedge strut” injector and wall-mounted ramp injectors both of which generate streamwise vortices. Firing test results showed a superior ability of this type of combustor to perform a supersonic combustion in the combustor core flow and operability in supersonic/subsonic combustion modes with high thrust performance in wide range of flight Mach numbers.

Mach 4 testing of scramjet inlet models

Six scramjet inlet models were tested in a Mach 4 wind tunnel. Wall and pitot pressure were measured and schlieren photographs were taken. Parameters of these models are side-plate sweep angle, contraction ratio, and cowl geometry. The shock pattern inside one of the models, as shown by schlieren photographs, coincides with the calculations. Both mass-capture ratio and total pressure recovery are 50-70%. There seems to be a better sweep angle and a better cowl length for maximum total pressure recovery.

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

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.

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