Ageneralreviewoftheemergenceandmaturingofsupersoniccombustionramjet (scramjet)enginetechnologies overthelast 40 years is presented. The review isgiven in terms of theeffortscompleted or proceeding in theUnited States, Russia, France, Germany, Japan, Australia, and other countries. A brief account of the challenges of scramjet combustor development is given, and attention is directed toward other source references for more detailed accounts of technology evolution. Two emerging scramjet applications are identie ed, namely, hydrogenfueledhigh-speed engines foraccess to spaceand hydrocarbon-fueled enginesforhypersonicair-launched missiles.

Scramjet Engines: The First Forty Years

A flux splitting scheme is proposed for the general nonequilibrium flow equations with an aim at removing numerical dissipation of Van-Leer-type flux-vector splittings on a contact discontinuity. The scheme obtained is also recognized as an improved Advection Upwind Splitting Method (AUSM) where a slight numerical overshoot immediately behind the shock is eliminated. The proposed scheme has favorable properties: high-resolution for contact discontinuities; conservation of enthalpy for steady flows; numerical efficiency; applicability to chemically reacting flows. In fact, for a single contact discontinuity, even if it is moving, this scheme gives the numerical flux of the exact solution of the Riemann problem. Various numerical experiments including that of a thermo-chemical nonequilibrium flow were performed, which indicate no oscillation and robustness of the scheme for shock/expansion waves. A cure for carbuncle phenomenon is discussed as well.

A Flux Splitting Scheme with High-Resolution and Robustness for Discontinuities(Proceedings of the 12th NAL Symposium on Aircraft Computational Aerodynamics)

Recent advances in the understanding of the structure, propagation, and extinction of laminar flames under the influence of stretch, as manifested by the existence of flame curvature, flow nonuniformity, and flame motion, are reviewed. The emphasis is on premixed flames because of the richness and subtlety of the phenomena involved. The review distinguishes the influences of the tangential and normal velocity gradients on the flame response, both at the hydrodynamic scale and within the flame structure, and emphasizes the importance of the preferential diffusion nature of heat and mass transport, as well as the extent to which the flame can freely adjust its location in response to stretch in order to achieve complete reaction. It is then demonstrated that stretch has only minimal effect on an adiabatic, unrestrained, diffusionally-neutral flame with complete reaction in that the temperature, propagation rate, and thickness of the flame are invariant to stretch, and that stretch alone cannot extinguish such a flame. In the presence of preferential diffusion and/or when the flame movement is restrained, the response of the flame to stretch becomes more sensitive and extinction is also possible. The concept of flame stretch is applied to interpret such practical flame phenomena as flame stabilization and flame-front instability, determination of laminar flame speeds and flammability limits, concentration and temperature modifications in flame chemistry, and modeling of turbulent flames. The properties of stretched diffusion flames are then briefly discussed. The review closes with suggestions for further research.

Dynamics of Stretched Flames

We present a practical implementation of calibration-free wavelength-modulation spectroscopy with second harmonic detection (WMS-2f) for measurements of gas temperature and concentration in harsh environments. The method is applicable to measurements using lasers with synchronous wavelength and intensity modulation (such as injection current-tuned diode lasers). The key factors that enable measurements without the on-site calibration normally associated with WMS are (1) normalization of the WMS-2f signal by the first harmonic (1f) signal to account for laser intensity, and (2) the inclusion of laser-specific tuning characteristics in the spectral-absorption model that is used to compare with measured 1f-normalized, WMS-2f signals to infer gas properties. The uncertainties associated with the calibration-free WMS method are discussed, with particular emphasis on the influence of pressure and optical depth on the WMS signals. Many of these uncertainties are also applicable to calibrated WMS measurements. An example experimental setup that combines six tunable diode laser sources between 1.3 and 2.0 mum into one probe beam for measurements of temperature, H(2)O, and CO(2) is shown. A hybrid combination of wavelength and frequency demultiplexing is used to distinguish among the laser signals, and the optimal set of laser-modulation waveforms is presented. The system is demonstrated in the harsh environment of a ground-test scramjet combustor. A comparison of direct absorption and 1f-normalized, WMS-2f shows a factor of 4 increase in signal-to-noise ratio with the WMS technique for measurements of CO(2) in the supersonic flow. Multidimensional computational fluid-dynamics (CFD) calculations are compared with measurements of temperature and H(2)O using a simple method that accounts for the influence of line-of-sight (LOS) nonuniformity on the absorption measurements. The comparisons show the ability of the LOS calibration-free technique to gain useful information about multidimensional CFD models.

Wavelength-modulation spectroscopy for measurements of gas temperature and concentration in harsh environments

Dynamics of stretched flames

To investigate the sensitivity of combustion to the test gas, an H2-fueled scramjet engine was tested at a Mach 6  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 dif 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  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  ame was anchored near the backward-facing step on the sidewalls. However, the  ame partially detached from the step on the top wall in the combustor. The detached  ame may make the combustion kinetically controlled to produce the sensitivity to the test air.

Comparison of Scramjet Engine Performance in Mach 6 Vitiated and Storage-Heated Air

To avoid combustor ‐inlet interaction,a staged supersoniccombustorwitha strutforthee rst-stageinjection and second-stage wall injectors at its divergent section was introduced to increase the pressure in the divergent section under the limitation of the peak pressure at its minimum cross-sectional area section. The combustor was tested experimentally in a direct-connect wind-tunnel facility with a combustion heater that supplied Mach 2.5 aire ow witha total temperatureof1500 K.With thestaged injections,thepressurerisedueto thesecond-stagecombustion was isolated from that due to the e rst-stage combustion. Fuel with an equivalence ratio more than unity, twice that of the allowable e ow rate with the e rst-stage injection alone, could be injected without the interaction, and the maximum thrust increment was augmented by 100%. The effects of thesecond-stageinjector placements were also examined, and it was found that the second-stage injection farther upstream resulted in the interaction, but that the injection further downstream resulted in low thrust due to low combustion efe ciency. Nomenclature A = cross-sectional area of e ow pass dF = thrust increment dIsp = specie c impulse based on thrust increment F = thrust P = pressure P.X/ = measured wall pressure at X, mm T = temperature X = streamwise location from step Y = spanwise location from duct center Z = lateral location from duct center ° = specie c heat ratio " = Crocco factor (see Ref. 10) A = equivalence ratio Subscripts i = combustor inlet o = combustor outlet 0 = stagnant condition 1 = e rst stage 2 = second stage

Combustion Tests of a Staged Supersonic Combustor with a Strut

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.

Mach 6 Testing of a Scramjet Engine Model

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.

Tohru Mitani

Papers

Comparison of Scramjet Engine Performance in Mach 6 Vitiated and Storage-Heated Air

Combustion Tests of a Staged Supersonic Combustor with a Strut

Mach 6 Testing of a Scramjet Engine Model

Reaction and Mixing-Controlled Combustion in Scramjet Engines

A flame inhibition theory by inert dust and spray