Development of a High-Pressure Gaseous Burner for Calibrating Optical Diagnostic Techniques

Source: NTRS

ABSTRACT In this work-in-progress report, we show the development of a unique high-pressure burner facility (up to 60 atm) that provides steady, reproducible premixed flames with high precision, while having the capability to use multiple fuel/oxidizer combinations. The highpressure facility has four optical access ports for applying different laser diagnostic techniques and will provide a standard reference flame for the development of a spectroscopic database in high-pressure/temperature conditions. Spontaneous Raman scattering (SRS) was the first diagnostic applied, and was used to successfully probe premixed hydrogen-air flames generated in the facility using a novel multi-jet micro-premixed array burner element. The SRS spectral data include contributions from H2, N2, O2, and H2O and were collected over a wide range of equivalence ratios ranging from 0.16 to 4.9 at an initial pressure of 10-atm via a spatially resolved point SRS measurement with a high-performance optical system. Temperatures in fuel-lean to stoichiometric conditions were determined from the ratio of the Stokes to anti-Stokes scattering of the Q-branch of N2, and those in fuel-rich conditions via the rotational temperature of H2. The SRS derived temperatures using both techniques were consistent and indicated that the flame temperature was approximately 500 K below that predicted by adiabatic equilibrium, indicating a large amount of heat-loss at the measurement zone. The integrated vibrational SRS signals show that SRS provides quantitative number density data in high-pressure H2-air flames.

  • 45th AIAA Aerospace Sciences Meeting and Exhibit; 01/2007
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    Journal of Propulsion and Power 01/2010; 26(1):186-189. DOI:10.2514/1.44640 · 0.61 Impact Factor
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    ABSTRACT: A new, high-pressure facility was used to investigate GO(2)/GH(2) single-element, shear, coaxial injectors with operating conditions typical of rocket engines. Oxygen-to-hydrogen mass flow ratios of 4 and 6 at operational chamber pressures of 6.2, 4.9, 4.5, and 2.8 MPa were investigated by 1) keeping the propellant mass flow rates constant while changing the exhaust nozzle diameter, and 2) by keeping the exhaust nozzle diameter constant and changing the propellant mass flow rates to change the chamber pressure. Axial heat fluxes, injector face temperature, and exit nozzle temperature were measured. The injector's outer diameter was 2.7 mm and the chamber cross section was square with a side L = 2.5 cm. Maximum heat release occurred at 2.4 L from the injector face. The injector face temperatures showed little to no dependence on chamber pressure except for high mass flow ratios. A clear dependence exists, however, on the chamber length. The profiles of heat flux and chamber wall temperatures indicated no pressure dependence and only a slight dependence on propellant injection velocities when the mass flows were kept constant. A scaling of heat flux values based on fuel mass flow rate, instead of chamber pressure, is, therefore, suggested. The lack of pressure dependence and only slight dependence on the propellant injection velocities suggested that the basic dynamic structures of the combusting flow were mainly dominated by the chamber geometry.
    Journal of Spacecraft and Rockets 05/2007; 44(3):633-639. DOI:10.2514/1.26678 · 0.47 Impact Factor


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