Sean O'Byrne

Analysis of Transient Thermal Choking Processes in a Model Scramjet Engine

Shadowgraph flow visualisation and floor static pressure measurements have been used to examine the transient behaviour of a thermally choked combusting flow. Experiments were performed to examine the effect of varying inlet Mach number and fuel-air equivalence ratio on the nature and extent of the interaction. In all cases a sudden increase in static pressure was measured, followed by a highly turbulent region of sonic flow which was seen to propagate upstream along the duct. The nature of the dominant processes causing this pressure discontinuity are still not certain. Some mechanisms which may contribute to this phenomenon are presented. These include separation of the boundary layer in the duct, formation of a detonation and formation of a near-normal shock wave by the region of thermally choked flow.

Coherent Anti-Stokes Raman Spectroscopic Thermometry in a Supersonic Combustor

An experiment has been conducted to acquire data for the validation of computational fluid dynamics codes used in the design of supersonic combustors. The flow in a supersonic combustor, consisting of a diverging duct with a single downstream-angled wall injector, is studied. Combustor entrance Mach number is 2 and enthalpy nominally corresponds to Mach 7 flight. The primary measurement technique is coherent anti-Stokes Raman spectroscopy, but surface pressures and temperatures have also been acquired. Modern design of experiment techniques have been used to maximize the quality of the data set (for the given level of effort) and to minimize systematic errors. Temperature maps are obtained at several planes in the flow for a case in which the combustor is piloted by injecting fuel upstream of the main injector and one case in which it is not piloted. Boundary conditions and uncertainties are characterized.

Dual-Pump Coherent Anti-Stokes Raman Scattering Measurements in a Supersonic Combustor

The dual-pump coherent anti-Stokes Raman scattering (CARS) method was used to measure temperature and the mole fractions of N 2 and O 2 in a supersonic combustor. Experiments were conducted in NASA Langley Research Centers Direct-Connect Supersonic Combustion Test Facility. In this facility, H 2 - and oxygen-enriched air burn to increase the enthalpy of the simulated air test gas. This gas is expanded through a Mach 2 nozzle and into a combustor model consisting of a short constant-area section followed by a small rearward-facing step and another constant-area section. At the end of this straight section, H 2 fuel is injected at Mach 2 and at a 30-deg angle with respect to the freestream. One wall of the duct then expands at a 3-deg angle for over 1 m. The ensuing combustion is probed optically through ports in the side of the combustor. Dual-pump CARS measurements were performed at the facility nozzle exit and at four planes downstream of fuel injection. Maps are presented of the mean temperature, as well as N 2 and O 2 mean mole-fraction fields. Correlations between fluctuations of the different measured parameters are also presented.

OH PLIF Imaging of Supersonic Combustion using Cavity Injection

Planar laser-induced fluorescence of the hydroxyl radical is used to investigate a cavitybased fuel injection system for scramjet operation at a flight Mach number of 11.5. Hydrogen and ethylene fuels are compared over a range of fuel-lean global equivalence ratios. Both fuels show evidence of combustion at each fuel injection pressure. Hydrogen burns diffusively in the shear layer and further downstream, with penetration distance increasing linearly with injection pressure. The combustion in the cavity shear layer sheds regular vortices. Ethylene behaves similarly to hydrogen at the lower equivalence ratios but shows evidence of intense, localized combustion at the highest equivalence ratio. In all cases combustion occurs in the shear layer above the cavity rather than in the recirculating cavity flow.

Dual-Pump CARS Thermometry and Species Concentration Measurements in a Supersonic Combustor

The dual-pump coherent anti-Stokes Raman spectroscopy (CARS) method was used to measure temperature and the absolute mole fractions of N2, O2 and H2 in a supersonic combustor. Experiments were conducted in NASA Langley Research Centers Direct Connect Supersonic Combustion Test Facility. In this facility, hydrogen and air bum to increase the enthalpy of the test gas; O2 is then added to simulate air. This gas is expanded through a Mach 2 nozzle and into a combustor model consisting of a short constant-area section followed by a small rearward facing step and another constant area section. At the end of this straight section H2 fuel is then injected at Mach 2 and at 30 deg. angle with respect to the freestream. One wall of the duct then expands at a 3 deg. angle for over 1 meter. The ensuing combustion is monitored optically through ports in the side of the combustor. CARS measurements were performed at the nozzle exit and at four different planes downstream fuel injection. Maps were obtained of the mean temperature, as well as quantitative N2 and O2 and qualitative H2 mean mole fraction fields. Correlations between fluctuations of the different measured parameters are presented for one of the planes of data.

Fluorescence Velocimetry of the Hypersonic, Separated Flow over a Cone

Planar laser-induced fluorescence of nitric oxide is used to measure a component of the velocity field for the Mach 7 flow around a 30-deg half-angle, 50-mm-diam cone mounted to a long, 38-mm-diam shaft, or sting. Transverse velocities are measured in the freestream, the shock layer, and the separated region at the junction between the cone and the sting. For most of the flowfield, the uncertainty of the measurements is between ±50 and ±100 m/s for velocities ranging from -300 to 1300 m/s, corresponding to a minimum uncertainty of ±5%. The measurements are compared with the commercial computational fluid dynamics (CFD) code CFD-FASTRAN . The agreement between the theoretical model and the experiment is reasonably good. CFD accurately predicts the size and shape of the shock layer and separated region behind the cone as well as the magnitude of the gas velocity near the reattachment shock. However, the magnitude of the velocity in the shock layer and gas expansion differ somewhat from that predicted by CFD. The discrepancies are attributed to a small systematic error associated with laser-beam attenuation and also to inexact modeling of the flowfield by CFD

CARS Temperature and Species Concentration Measurements in a Supersonic Combustor with Normal Injection

The dual-pump coherent anti-Stokes Raman spectroscopy (CARS) method was used to measure temperature and the absolute mole fractions of N2, O2 and H2 in a supersonic combustor. Experiments were conducted in the NASA Langley Direct-Connect Supersonic Combustion Test Facility. CARS measurements were performed at the facility nozzle exit and at three planes downstream of fuel injection. Processing the CARS measurements produced maps of the mean temperature, as well as quantitative N2 and O2 and qualitative H2 mean mole fraction fields at each plane. The CARS measurements were also used to compute correlations between fluctuations of the different simultaneously measured parameters. Comparisons were made between this 90 degree angle fuel injection case and a 30 degree fuel injection case previously presented at the 2004 Reno AIAA Meeting.

Diode-laser-based near-resonantly enhanced flow visualization in shock tunnels

Two new near-resonantly enhanced flow visualization techniques suitable for hypersonic low-density flows in shock or arc tunnels have been developed using seeded lithium (Li) metal as the refractivity-enhancing species. Two semiconductor lasers, single-longitudinal-mode and multimode, are compared with respect to their suitability as light sources for the technique. Transient wake-flow structures around a cylinder and a model of a planetary entry vehicle are visualized to demonstrate the capabilities of this comparatively inexpensive and simple visualization system. The images show flow features which are undetectable with conventional schlieren, shadowgraph, or interferometry techniques. Furthermore, the effect of density inhomogeneities along the line-of-sight outside the region of interest can be reduced by enhancing the refractivity only in selected parts of the flowfield.

Nonintrusive Temperature and Velocity Measurements in a Hypersonic Nozzle Flow

Distributions of nitric oxide vibrational temperature, rotational temperature andvelocity have been measured in the hypersonic freestream at the exit of a conical noz-zle, using planar laser-induced fluorescence. Particular attention has been devoted toreducing the major sources of systematic error that can affect fluorescence tempera-ture measurements, including beam attenuation, transition saturation effects, laser modefluctuations and transition choice. Visualization experiments have been performed toimprove the uniformity of the nozzle flow. Comparisons of measured quantities with asimple one-dimensional computation are made, showing good agreement between mea-surements and theory given the uncertainty of the nozzle reservoir conditions and thevibrational relaxation rate.

SUPERSONIC COMBUSTION EXPERIMENTS FOR CFD MODEL DEVELOPMENT AND VALIDATION (INVITED)

This paper briefly reviews a number of experiments suitable for the screening, development and validation of turbulence models used in computational fluid dynamics (CFD). Two sets of experiments conducted for this purpose by the authors are presented. The first is a compressible axisymmetric jet in which a central jet of helium mixes with a surrounding coflow of air. The second is a supersonic combustor. Various instrumentation techniques have been employed with the axisymmetric jet, whereas the dual pump coherent antiStokes Raman spectroscopy technique has been applied to the combustor. Experiences in trying to calculate these flows, and set model constants, are reviewed.