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Liquid Oxygen-Liquid Methane Ignition Demonstrated for Application to Reaction Control Engines

A workhorse liquid-oxygen/liquid-methane (LOX/LCH4) ignition system was recently tested in NASA Glenn Research Center’s Research Combustion Laboratory. The igniter was an in-house design used to evaluate the ignition processes for LOX/LCH4. The tests examined the flammability of LOX/LCH4 over a range of oxidizer-to-fuel mixture ratios. In addition, ignition pulses were accumulated to examine hardware durability: 1377 individual ignition pulses were successfully demonstrated.

Glenn’s Workhorse igniter was designed around a bluff-body-tipped sparkplug (see the drawing). The exciter unit for the spark plug delivered 200 sparks/sec at 20 kV and 70 to 150 mJ. Design flow rates were in the 10-lbf thrust class. The igniter used three separate propellant feed lines, two for the fuel and one for the oxidizer. LOX was injected with four doublets in the annulus behind the bluff body on the spark plug tip and was excited by the spark. One fuel line injected fuel with four doublets just downstream of the spark plug tip to provide an oxidizer-rich core flow, while the second fuel line injected fuel with tangential swirl to supply film-cooling flow to the chamber wall and an overall fuel-rich torch. The core mixture ratio ranged between 10 and 22, and the overall mixture ratio was between 1.1 and 2.0. Both mixture ratios were selected because their flame temperatures are compatible with conventional materials. Propellant mass flow was controlled by cavitating venturis. Injector element pressure drops were designed to accommodate the desired propellant mass flows as saturated vapor as well as saturated liquid.

Sketch showing spark plug and propellant injection locations
Workhorse igniter design.

The LOX/LCH4 igniter was tested in Glenn’s RCL-21, a rocket test stand that can simulate an altitude of 100,000 ft (10 torr, or 0.2 psia). A liquid nitrogen cooling system was used to condense gaseous oxygen and gaseous methane in small propellant tanks. The cooling lines extended from the tanks to the igniter inlet valves to help ensure that the propellants remained in a liquid state up to the igniter manifold. An additional cooling line was used to cool the igniter body to simulate operating conditions in the cold soak of space.

Performance testing was designed to evaluate the effect of mixture ratio variations on the ignition process. A typical ignition reproduced from the video monitor is shown in the photograph. Single-pulse, 0.5-sec tests were conducted. Results indicated that ignition was possible across the entire range of mixture ratios tested with the igniter body at ambient temperatures.

Photograph showing a luminous torch about 6 inches long
Ignition plume shown in the video monitor.

Another objective was to accumulate ignition pulses to gauge the expected lifetime of the igniter components and valves as well as the repeatability of the ignition pulses. A 0.25-sec pulse was used so that a large number of pulses could be put on the igniter. During these tests, the igniter was operated in multiple pulse strings at a 10-percent duty cycle. The graph is a typical chamber pressure trace showing 10 pulses. The ignition pulse repeatability varied slightly over the course of testing because fluctuations in propellant inlet conditions led to slight variations in the propellant flow rates. This series of tests was halted after the igniter spark plug ceramic insulator failed.

Graph of pressure versus time; pressure trace shows 10 pulses, 0.25-seconds long, with a pressure between 200 and 250 pounds per square inch absolute
Pressure trace from pulse train.

Glenn contacts: Dr. Steven J. Schneider, 216-977-7484, Steven.J.Schneider@nasa.gov, Jeremy W. John, 216-433-6199, Jeremy.W.John@nasa.gov
Authors: Dr. Steven J. Schneider, Jeremy W. John, Joseph G. Zoeckler, Lynn A. Arrington, Jason C. Wendell, and Dale M. Diedrick
Headquarters program office: Exploration Systems Mission Directorate
Programs/projects: Propulsion and Cryogenics Advanced Development project, Reaction Control Engine

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Last updated: December 14, 2007

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