Direct-connect Scramjet Combustor Research Articles

An Experimental Investigation of Low-Frequency Active Excitation in Scramjet Combustor Using a Micro-Pulse Detonation Engine

A micro-pulse detonation engine (μPDE) was designed and installed to a direct-connect scramjet combustor of Pusan National University (PNU-DCSC). The active excitation on the scramjet combustor was experimentally studied using the μPDE operating at frequencies of 10–20 Hz. A vitiation air heater (VAH) was used to supply high-enthalpy vitiated air to the isolator and the scramjet combustor at a Mach number of 2.0, with a total temperature of 1600 K and a total pressure of 1.68 MPa. The exit of μPDE was located at the center of the cavity of the scramjet combustor. Active excitation was performed at equivalence ratios of 0.111 and 0.163, and characteristics were analyzed through Schlieren recording and bottom wall pressure measurement. As a result, when the detonation emitted from the μPDE entered the scramjet combustor, it instantly formed a shock train and moved forward within the scramjet combustor. The flame instantaneously changed from the cavity shear layer flame to the cavity flame. Through bottom wall pressure measurement, it was also observed that active excitation resulted in a significant pressure increase near the cavity compared to when active excitation was not performed. This revealed combustion characteristics, indicating improved combustion efficiency from the pressure increase in the scramjet combustor.

A Study on a Vitiated Air Heater for a Direct-Connect Scramjet Combustor and Preliminary Test on the Scramjet Combustor Ignition

Vitiation air heater (VAH) combustion characteristics for a direct-connect scramjet combustor (DCSC) were experimentally studied. The VAH consists of a head, modular chamber, and circular-to-rectangular shape transition (CRST) nozzle. The CRST nozzle transforms the circular cross-sectioned rocket-type VAH into a rectangular cross-sectioned scramjet combustor. The CRST nozzle exit Mach numbers at the top, middle, and bottom were measured using a tungsten wedge. The oblique shock formed by the wedge was captured using Schlieren visualization and recorded with a high-speed camera. The θ-β-M relation showed that the exit Mach number was 2.04 ± 0.04 with a chamber pressure of 1.685 ± 0.07 MPa. With the VAH design point verified, preliminary scramjet combustor ignition tests were conducted. As the fuel was not auto-ignited by the vitiated air, the forced ignition method, in which VAH ignition flame ignites the scramjet fuel, was used. The Schlieren images showed that a cavity shear layer combustion mode was formed and also showed that the forced ignition method could be used as a reference model for the ignitor-ignition method.

Circular Scramjet Combustor Transition Lengths at Mach 8 and 12.8 Flight Conditions

This paper presents an experimental study of laminar-to-turbulent boundary-layer transition in a circular duct at high supersonic Mach number. The experiments were conducted as an essential precursor for scramjet drag reduction studies using a direct-connect scramjet combustor. The experiments were carried out at the University of Queensland T4 shock tunnel using a 33.2-mm-diam and 500-mm-long circular duct at flow stagnation enthalpies of 3.2 and . The experimental work was supported with two-dimensional and three-dimensional Reynolds-averaged Navier–Stokes simulations. Comparisons between numerical and experimental results show initial rise in heat transfer at the first shock-wave boundary-layer interaction, followed by an unsteady laminar region interspersed with possible turbulent spots. Heat transfer measurements along the wall that show sustained transition to turbulence did not occur within the full length of the duct. Transition to turbulence within the circular duct is significantly delayed when compared with estimations based on flat-plate results. These results indicate that there is value in revisiting the assumption of a fully turbulent boundary layer for previous semi-direct-connect scramjet combustion experiments that have used transition length correlations derived from flat-plate experiments within the same facility. The findings from this investigation have implications for conclusions drawn from experimental investigations of boundary-layer-combustion-induced drag reduction in circular ducts.

Experimental and numerical investigations of combustion mode transition in a direct–connect scramjet combustor

Direct–connect experiment and numerical simulation investigations on combustion mode transition were completed for inflow conditions of Mach number 3.0, total temperature 1500 K and total pressure 2.1 MPa. Pressure at centerline of combustor upper wall was measured. Numerical simulation results, containing Mach number distribution of combustion chamber, shock wave structure of isolator, and heat release rate of combustor, were compared at various equivalence ratios. The quantificational distinguish criterion of combustion mode (po/pi) was constructed using wall pressure po of outlet and pi of inlet for isolator. The combustor works in scram-mode when the pressure ratio is below 1.5, while it is in ram-mode when the value of po/pi is over 1.5. Two types of combustion mode could also be identified base on heat release distribution of combustion chamber and flow configuration of isolator. In scram-mode, the heat release is distributed equally, the flow in combustor is supersonic, and there is no large separation region in isolator. In ramjet mode, the heat release is distributed concentrated, the flow at the entrance of combustor and in the combustor are both subsonic, and there is large separation region and shock train in isolator.

Simultaneous sensing of temperature, CO, and CO2 in a scramjet combustor using quantum cascade laser absorption spectroscopy

A mid-infrared laser absorption sensor was developed for gas temperature and carbon oxide (CO, CO2) concentrations in high-enthalpy, hydrocarbon combustion flows. This diagnostic enables non-intrusive, in situ measurements in harsh environments produced by hypersonic propulsion ground test facilities. The sensing system utilizes tunable quantum cascade lasers capable of probing the fundamental mid-infrared absorption bands of CO and CO2 in the 4–5 µm wavelength domain. A scanned-wavelength direct absorption technique was employed with two lasers, one dedicated to each species, free-space fiber-coupled using a bifurcated hollow-core fiber for remote light delivery on a single line of sight. Scanned-wavelength modulation spectroscopy with second-harmonic detection was utilized to extend the dynamic range of the CO measurement. The diagnostic was field-tested on a direct-connect scramjet combustor for ethylene–air combustion. Simultaneous, laser-based measurements of carbon monoxide and carbon dioxide provide a basis for evaluating combustion completion or efficiency with temporal and spatial resolution in practical hydrocarbon-fueled engines.

Spatially Resolved Water Measurements in a Scramjet Combustor Using Diode Laser Absorption

A two-color tunable diode laser sensor for detection of absorption near was used to simultaneously determine temperature, column density, and velocity in a direct-connect scramjet combustor. Axial and transverse spatial resolutions were achieved by translating the measurement line of sight across and along the flowpath with measurements spanning six axial planes and 24 measurement lines of sight in each plane. Results are presented for a uniform noncombusting -seeded case as well as for air combustion at equivalence ratios of 0.17 and 0.46. A combination of wavelength-modulation spectroscopy and scanned-wavelength direct absorption were used to mitigate nonuniform gas conditions along the measurement line of sight present in the combustion cases. Comparisons of tunable diode laser absorption spectroscopy measurements with computational fluid dynamics simulations are included and illustrate the use of line-of-sight tunable diode laser absorption spectroscopy to validate computational efforts. Additional comparisons with coherent anti-Stokes Raman scattering measurements in the same model scramjet corroborate many key findings from tunable diode laser absorption spectroscopy.

Collaborative Experimental and Computational Study of a Dual-Mode Scramjet Combustor

Advanced computational models of hypersonic air-breathing combustion processes are being developed to better understand and predict the complex flows within a dual-mode scramjet combustor. However, the accuracy of these models can only be quantified through comparison to experimental databases. Moreover, the quality of computational results is dependent on accurate and detailed knowledge of the combustor inflow and boundary conditions. Toward these ends, this paper describes results from a collaboration of experimental and computational investigators. Detailed computational fluid dynamics and finite element analyses were performed throughout the design and implementation of experiments involving a direct-connect scramjet combustor operating at steady state during long duration testing. The test section hardware was designed to provide substantial access for optical laser diagnostics. Measurement locations included the inflow plane and several locations downstream of fuel injection. A suite of advanced in-stream diagnostics were applied, many of which are described in companion papers. Significant results in this paper include measured static wall pressures and temperatures, stereoscopic particle image velocimetry, and focused schlieren imaging. Validated thermal finite element calculations in the scramjet hardware and temperature maps of the flow path boundaries are also presented. Comparison of experimental results with computational fluid dynamics predictions are discussed in a separate paper.

Dual-Pump Coherent Anti-Stokes Raman Spectroscopy Measurements in a Dual-Mode Scramjet

In this paper, the authors describe dual-pump coherent anti-Stokes Raman spectroscopy (CARS) measurements of mixing and combustion in a direct-connect scramjet combustor operating at equivalent flight Mach typical of the ramjet–scramjet transition, in the scram mode. Measurements were performed in the University of Virginia’s scramjet test facility in which the air is heated by electrical resistance heaters. The CARS technique is used to acquire temporally and spatially resolved measurements of temperature and species mole fraction. Measurements were at four planes: one upstream of an fuel injector and three downstream. Contour plots of mean flow and standard deviation statistics are presented for cases with and without reaction of the fuel. The vibrational temperature at the exit of the facility nozzle, and in the freestream of the scramjet combustor, is elevated compared with the rotational temperature; the air– vibrational temperature is the same as the facility stagnation temperature. There are spatial nonuniformities of temperature exiting the heater. The mixing of the fuel jet from the single ramp wall injector and the growth of the plume downstream is shown. The flame is attached at the injector and propagates from the wall adjacent to the fuel plume, around the periphery of the plume, before engulfing the whole plume. Computational fluid dynamics modeling shows that the flow can be predicted well using hybrid large-eddy simulation/Reynolds-averaged Navier–Stokes methods with relatively simple subgrid models, provided facility effects are properly accounted for.

Issues associated with long-duration high-enthalpy scramjet combustor testing

Long-duration direct-connect combustor tests are an essential element in the development of an effective supersonic combustion ramjet (scramjet). While test techniques and analysis methodology have been established for simulated flight Mach numbers at Mach 8 and below, phenomena associated with higher flight simulations require additional attention. In this article, technical issues associated with long-duration high-enthalpy directconnect scramjet combustor tests are discussed. Since ground tests form the basis by which flight hardware is designed, it is important to quantify the differences between flight and ground test conditions. Analyses are presented herein which characterize the differences between ground test and flight combustor inlet properties. In particular, the effects of dynamic pressure and chemical nonequilibrium kinetics are investigated. Analytical results will also be presented which characterize aspects of the performance of a generic combustor operating at Mach 8,11, and 12 flight simulations. One-, two-, and three-stream mixing models are used to assess combustor performance, and the predictions derived from the various models are contrasted.