Osborne, Robin J. | Engineering Papers

NASA NTRS2017

Development of Augmented Spark Impinging Igniter System for Methane Engines

The Lunar Cargo Transportation and Landing by Soft Touchdown (Lunar CATALYST) program is establishing multiple no-funds-exchanged Space Act Agreement (SAA) partnerships with U.S. private sector entities. The purpose of this program is to encourage the development of robotic lunar landers that can be integrated with U.S. commercial launch capabilities to deliver payloads to the lunar surface. As part of the efforts in Lander Technologies, NASA Marshall Space Flight Center (MSFC) is developing liquid oxygen (LOX) and liquid methane (LCH4) engine technology to share with the Lunar CATALYST partners. Liquid oxygen and liquid methane propellants are attractive owing to their relatively high specific impulse for chemical propulsion systems, modest storage requirements, and adaptability to NASA's Journey to Mars plans. Methane has also been viewed as a possible propellant choice for lunar missions, owing to the performance benefits and as a technology development stepping stone to Martian missions. However, in the development of methane propulsion, methane ignition has historically been viewed as a high risk area in the development of such an engine. A great deal of work has been conducted in the past decade devoted to risk reduction in LOX/CH4 ignition. This paper will review and summarize the history and results of LOX/CH4 ignition programs conducted at NASA. More recently, a NASA-developed Augmented Spark Impinging (ASI) igniter body, which utilizes a conventional spark exciter system, is being tested with LOX/CH4 to help support internal and commercial engine development programs, such as those in Lunar CATALYST. One challenge with spark exciter systems, especially at altitude conditions, is the ignition lead that transmits the high voltage pulse from the exciter to the spark igniter (spark plug). The ignition lead can be prone to corona discharge, reducing the energy delivered by the spark and potentially causing non-ignition events. For the current work, a commercial compact exciter system, which eliminates this high voltage cabling, was tested at altitude conditions. A modified, conventional exciter system with an improved ignition lead was also recently tested at altitude conditions. This test program demonstrated the capability of these exciter systems to operate at altitude. While more extensive testing may be required, these systems or similar ones may be used for future NASA and commercial engine programs.

methane↗

NASA NTRS2017

Development of Augmented Spark Impinging Igniter System for Methane Engines

The Lunar Cargo Transportation and Landing by Soft Touchdown (Lunar CATALYST) program is establishing multiple no-funds-exchanged Space Act Agreement (SAA) partnerships with U.S. private sector entities. The purpose of this program is to encourage the development of robotic lunar landers that can be integrated with U.S. commercial launch capabilities to deliver payloads to the lunar surface. NASA can share technology and expertise under the SAA for the benefit of the CATALYST partners. MSFC seeking to vacuum test Augmented Spark Impinging (ASI) igniter with methane and new exciter units to support CATALYST partners and NASA programs. ASI has previously been used/tested successfully at sea-level, with both O2/CH4 and O2/H2 propellants. Conventional ignition exciter systems historically experienced corona discharge issues in vacuum. Often utilized purging or atmospheric sealing on high voltage lead to remedy. Compact systems developed since PCAD could eliminate the high-voltage lead and directly couple the exciter to the spark igniter. MSFC developed Augmented Spark Impinging (ASI) igniter. Successfully used in several sea-level test programs. Plasma-assisted design. Portion of ox flow is used to generate hot plasma. Impinging flows downstream of plasma. Additional fuel flow down torch tube sleeve for cooling near stoichiometric torch flame. Testing done at NASA GRC Altitude Combustion Stand (ACS) facility 2000-lbf class facility with altitude simulation up to around 100,000 ft. (0.2 psia [10 Torr]) via nitrogen driven ejectors. Propellant conditioning systems can provide temperature control of LOX/CH4 up to test article.

Chemical Propulsion↗

NASA NTRS2013

Component Testing of the J-2X Augmented Spark Igniter (ASI)

In support of the development of the J-2X engine, 201 low pressure, liquid oxygen / liquid hydrogen (LOX/LH2) J-2X Augmented Spark Igniter (ASI) subsystem ignition tests were conducted at Marshall Space Flight Center (MSFC). The main objective of these tests was to start the ASI within the anticipated J-2X engine start box, as well as outside of it, to check for ignition margin. The setup for the J-2X ASI component testing simulated, as much as possible, the tank-head start-up configuration of the ASI within the J-2X Engine. The ignition tests were divided into 124 vacuum start tests to simulate altitude start on a flight engine, and 77 sea-level start tests to simulate the first set of ground tests for the J-2X Engine at Stennis Space Center (SSC). Other ignition parameters that were varied included propellant tank pressures, oxidizer temperature entering the ASI oxidizer feedline, oxidizer valve timing, spark igniter condition (new versus damaged), and oxidizer and fuel feedline orifice sizes. Propellant blowdowns using venturis sized to simulate the ASI resistance allowed calculation of transient propellant mass flow rates as well as global mixture ratio for all ignition tests. Global mixture ratio within the ASI at the time of ignition varied from 0.2 to 1.2. Detailed electronics data obtained from an instrumented ignition lead allowed characterization of the breakdown voltage, sustaining voltage and energy contained in each spark as the ASI propellants ignited. Results indicated that ignition always occurred within the first five sparks when both propellants were present in the ASI chamber.

Osborne, Robin J.↗

NASA NTRS2010

Igniter Testing and Development for Liquid Oxygen/Liquid Methane (LOX/LCH4) and Liquid Oxygen/Gaseous Hydrogen (LOX/GH2) Injectors

In support of NASA's efforts to advance methane technology for future exploration missions, two different igniter concepts were developed and tested at NASA's Marshall Space Flight Center for ignition of oxygen/methane propellants. The first type of igniter, an impinging injection, spark-initiated torch igniter (impinging igniter), used traditional spark plug technology to produce a small, but very intense, pulsed plasma as the ignition source. The second type of igniter, a microwave-generated plasma igniter (microwave igniter), used a microwave magnetron to produce a continuous, larger-volume plasma as the ignition source. Two new designs for both types of igniters were leveraged off past oxygen/methane igniter designs that were tested at MSFC and documented in a 2008 JANNAF paper. In this most recent work, the two new igniters were to be tested at the component level as well as inside a liquid oxygen/liquid methane (LOX/LCH4) 40-element swirl coaxial injector. Although the microwave igniter achieved ignition in every test, it encountered technical difficulties that prevented it from being tested with the injector. The impinging igniter required several design iterations as well as testing at different mixture ratios and mass flow rates before reliable operation with minimal ignition delay was achieved. The final design of the impinging igniter was successfully tested in the LOX/LCH4 injector. The same igniter was later re-configured and component-tested to demonstrate reliable ignition of oxygen/hydrogen propellants. Future tests of the impinging igniter will include ignition of a LOX/hydrogen two-stage swirl injector, which is one of several injector concepts being tested in the Lunar Lander Descent Engine Test Bed at MSFC. All igniter and injector tests have been funded by the Propulsion & Cryogenics Advanced Development project managed by Glenn Research Center and sponsored by NASA s Exploration Technology Development Program.

Osborne, Robin J.↗

NASA NTRS2008

Igniters for Liquid Oxygen/Liquid Methane Technology Development

As part of NASA's technology development of liquid methane / liquid oxygen engines for future exploration missions, two different igniters were recently studied at NASA Marshall Space Flight Center. The first igniter tested was an impinging injection, spark-initiated torch igniter, and the second was a microwave-generated plasma igniter. The purpose of the ignition tests was to define the ignition limits under vacuum conditions and characterize the transient start-up performance as a function of propellant mixture ratio (MR), mass flow rates, inlet temperatures, and pre-ignition chamber pressure. In addition, for the impinging igniter two different spark plugs were tested, and for the microwave igniter the magnetron filament warm-up time and the magnetron input power were both varied. The results gathered from these tests indicated that the impinging igniter is capable of operating over an MR range of 2 - 27, with methane and oxygen inlet temperatures as low as -161 F and -233 F, respectively. The microwave igniter was tested over an MR range of 2 - 9, with methane and oxygen inlet temperatures as low as -90 F and -200 F, respectively. The microwave igniter achieved ignition over this range, although an upper ignition limit was determined for the oxidizer mass flow rate. In general, the torch exhaust temperatures for the microwave igniter were not as high as those attained with the impinging igniter. The microwave igniter, however, was hot-fired 17 times and was still operational, whereas the impinging igniter spark plugs experienced thermal shock and erosion over nine hot-fire tests. It was concluded that for the microwave igniter better mixing of the propellants might be required in order to both raise the torch exhaust temperature and decrease the required magnetron input power, and for the impinging igniter the spark plug position within the igniter chamber should be varied in future tests to identify a more optimal location. All of the igniter tests were supported by the Propulsion & Cryogenics Advanced Development project, which is part of NASA's Exploration Technology Development Program.

Osborne, Robin J.↗

NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner, shown schematically in Figure 1 and similar in design to apparatuses used by other researchers to study solid and liquid monopropellants, will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗

NASA NTRS2001

Measuring Rocket Engine Temperatures with Hydrogen Raman Spectroscopy

Optically accessible, high pressure, hot fire test articles are available at NASA Marshall for use in development of advanced rocket engine propellant injectors. Single laser-pulse ultraviolet (UV) Raman spectroscopy has been used in the past in these devices for analysis of high pressure H2- and CH4-fueled combustion, but relies on an independent pressure measurement in order to provide temperature information. A variation of UV Raman (High Resolution Hydrogen Raman Spectroscopy) is under development and will allow temperature measurement without the need for an independent pressure measurement, useful for flows where local pressure may not be accurately known. The technique involves the use of a spectrometer with good spectral resolution, requiring a small entrance slit for the spectrometer. The H2 Raman spectrum, when created by a narrow linewidth laser source and obtained from a good spectral resolution spectrograph, has a spectral shape related to temperature. By best-fit matching an experimental spectrum to theoretical spectra at various temperatures, a temperature measurement is obtained. The spectral model accounts for collisional narrowing, collisional broadening, Doppler broadening, and collisional line shifting of each Raman line making up the H2 Stokes vibrational Q-branch spectrum. At pressures from atmospheric up to those associated with advanced preburner components (5500 psia), collisional broadening though present does not cause significant overlap of the Raman lines, allowing high resolution H2 Raman to be used for temperature measurements in plumes and in high pressure test articles. Experimental demonstrations of the technique are performed for rich H2-air flames at atmospheric pressure and for high pressure, 300 K H2-He mixtures. Spectrometer imaging quality is identified as being critical for successful implementation of technique.

Wehrmeyer, Joseph A.↗