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All sources NASA NTRS DOE OSTI

NASA NTRS1990

Mars in situ propellants: Carbon monoxide and oxygen ignition experiments

Carbon monoxide and oxygen were tested in a standard spark-torch igniter to identify the ignition characteristics of this potential Mars in situ propellant combination. The ignition profiles were determined as functions of mixture ratio, amount of hydrogen added to the carbon monoxide, and oxygen inlet temperature. The experiments indicated that the carbon monoxide and oxygen combination must have small amounts of hydrogen present to initiate reaction. Once the reaction was started, the combustion continued without the presence of hydrogen. A mixture ratio range was identified where ignition occurred, and this range varied with the oxygen inlet temperature.

Linne, Diane L.↗

NASA NTRS1990

Mars in situ propellants - Carbon monoxide and oxygen ignition experiments

Linne, Diane L.↗

NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are summarized. Ignition characterization was accomplished through the analysis, fabrication and testing of a spark initiated torch igniter and prototype 620 1bF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 3.3 to 262 psia and a mixture ratio range of 0.4 to 40. The prototype thruster/igniter assembly was tested over the chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗

NASA NTRS2008

LOX/Methane Main Engine Igniter Tests and Modeling

The LOX/methane propellant combination is being considered for the Lunar Surface Access Module ascent main engine propulsion system. The proposed switch from the hypergolic propellants used in the Apollo lunar ascent engine to LOX/methane propellants requires the development of igniters capable of highly reliable performance in a lunar surface environment. An ignition test program was conducted that used an in-house designed LOX/methane spark torch igniter. The testing occurred in Cell 21 of the Research Combustion Laboratory to utilize its altitude capability to simulate a space vacuum environment. Approximately 750 ignition test were performed to evaluate the effects of methane purity, igniter body temperature, spark energy level and frequency, mixture ratio, flowrate, and igniter geometry on the ability to obtain successful ignitions. Ignitions were obtained down to an igniter body temperature of approximately 260 R with a 10 torr back-pressure. The data obtained is also being used to anchor a CFD based igniter model.

Breisacher, Kevin J.↗

NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are presented. Ignition characterization was accomplished through the analysis, design, fabrication and testing of a spark initiated torch igniter and prototype 620 lbF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗

NASA NTRS2012

Experimental Investigation of Augmented Spark Ignition of a LO2/LCH4 Reaction Control Engine at Altitude Conditions

The use of nontoxic propellants in future exploration vehicles would enable safer, more cost-effective mission scenarios. One promising green alternative to existing hypergols is liquid methane (LCH4) with liquid oxygen (LO2). A 100 lbf LO2/LCH4 engine was developed under the NASA Propulsion and Cryogenic Advanced Development project and tested at the NASA Glenn Research Center Altitude Combustion Stand in a low pressure environment. High ignition energy is a perceived drawback of this propellant combination; so this ignition margin test program examined ignition performance versus delivered spark energy. Sensitivity of ignition to spark timing and repetition rate was also explored. Three different exciter units were used with the engine s augmented (torch) igniter. Captured waveforms indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks. This suggests that rising pressure and flow rate increase spark impedance and may at some point compromise an exciter s ability to complete each spark. The reduced spark energies of such quenched deliveries resulted in more erratic ignitions, decreasing ignition probability. The timing of the sparks relative to the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1 to 6 mJ, though multiple, similarly timed sparks of 55 to 75 mJ were required for reliable ignition. Delayed spark application and reduced spark repetition rate both correlated with late and occasional failed ignitions. An optimum time interval for spark application and ignition therefore coincides with propellant introduction to the igniter.

Kleinhenz, Julie↗

NASA NTRS2012

Spark Ignition Characteristics of a L02/LCH4 Engine at Altitude Conditions

The use of non-toxic propellants in future exploration vehicles would enable safer, more cost effective mission scenarios. One promising "green" alternative to existing hypergols is liquid methane/liquid oxygen. To demonstrate performance and prove feasibility of this propellant combination, a 100lbf LO2/LCH4 engine was developed and tested under the NASA Propulsion and Cryogenic Advanced Development (PCAD) project. Since high ignition energy is a perceived drawback of this propellant combination, a test program was performed to explore ignition performance and reliability versus delivered spark energy. The sensitivity of ignition to spark timing and repetition rate was also examined. Three different exciter units were used with the engine s augmented (torch) igniter. Propellant temperature was also varied within the liquid range. Captured waveforms indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks (in quiescent, room air). The escalating pressure and flow environment increases spark impedance and may at some point compromise an exciter s ability to deliver a spark. Reduced spark energies of these sparks result in more erratic ignitions and adversely affect ignition probability. The timing of the sparks relative to the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1-6mJ, though multiple, similarly timed sparks of 55-75mJ were required for reliable ignition. An optimum time interval for spark application and ignition coincided with propellant introduction to the igniter and engine. Shifts of ignition timing were manifested by changes in the characteristics of the resulting ignition.

Kleinhenz, Julie↗

NASA NTRS2012

Spark Ignition Characteristics of a LO2/LCH4 Engine at Altitude Conditions

The use of non-toxic propellants in future exploration vehicles would enable safer, more cost effective mission scenarios. One promising "green" alternative to existing hypergols is liquid methane/liquid oxygen. To demonstrate performance and prove feasibility of this propellant combination, a 100lbf LO2/LCH4 engine was developed and tested under the NASA Propulsion and Cryogenic Advanced Development (PCAD) project. Since high ignition energy is a perceived drawback of this propellant combination, a test program was performed to explore ignition performance and reliability versus delivered spark energy. The sensitivity of ignition to spark timing and repetition rate was also examined. Three different exciter units were used with the engine's augmented (torch) igniter. Propellant temperature was also varied within the liquid range. Captured waveforms indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks (in quiescent, room air). The escalating pressure and flow environment increases spark impedance and may at some point compromise an exciter.s ability to deliver a spark. Reduced spark energies of these sparks result in more erratic ignitions and adversely affect ignition probability. The timing of the sparks relative to the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1-6mJ, though multiple, similarly timed sparks of 55-75mJ were required for reliable ignition. An optimum time interval for spark application and ignition coincided with propellant introduction to the igniter and engine. Shifts of ignition timing were manifested by changes in the characteristics of the resulting ignition.

Kleinhenz, Julie↗

NASA NTRS1997

Demonstration of Oxygen and Carbon Monoxide Propellants for Mars Missions

Currently, proposed planetary exploration missions must be small, with low costs and a short development time. Relatively high-risk technologies are being accepted for such missions if they meet these guidelines. For a Mars sample-return mission, one of the higher risk technologies is the use of return propellants produced from indigenous materials such as the Martian atmosphere. This consists of 96 percent carbon dioxide, which can be processed into oxygen and carbon monoxide. This year, the NASA Lewis Research Center completed the experimental evaluation and subscale technology development of an oxygen/carbon monoxide propellant combination. Previous research included ignition characterization, combustion performance, and heat transfer characterization with gaseous propellants at room temperature. In this year s tests, we studied the ignition characteristics and combustion of oxygen and carbon monoxide at near liquid temperatures. The mixture ratio boundaries for oxygen and carbon monoxide were determined as a function of propellant temperature in a spark torch igniter. With both propellants at room temperature, the ignition range was between 0.50 and 1.44; and with both propellants chilled to near-liquid temperatures, it was between 2.4 and 3.1. Statistical analysis of the mean value of the ignition boundaries provided models that describe the combination of oxygen temperature, carbon monoxide temperature, and mixture ratio that resulted in ignition. This range is the larger boxed area shown in the figure. The smaller boxed area indicates the range at which there is a 90-percent confidence that ignition will occur. The relatively small range at only 90-percent confidence indicates that using the oxygen/carbon monoxide combination as its own ignition source may not be the best design for a remote engine operating on Mars. Tests also were performed in a simulated small rocket engine that used oxygen/hydrogen combustion gases as the ignition source for oxygen/carbon monoxide. In these experiments, the oxygen/carbon monoxide was successfully ignited in eight of eight tests at a mixture ratio of 0.52. In addition, the oxygen/carbon monoxide maintained steady combustion after the oxygen/hydrogen ignition source was removed, verifying that the oxygen/carbon monoxide rocket engine should continue to be included in mission plans as return propulsion from Mars.

Linne, Diane L.↗

NASA NTRS2003

Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) for Rocket Engine Ignition System Application

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Laser ignition has been used at MSFC in recent test series to successfully ignite RP1/GOX propellants in a subscale rocket chamber, and other past studies by NASA GRC have demonstrated the use of laser ignition for rocket engines. Despite the progress made in the study of this ignition method, the logistics of depositing laser sparks inside a rocket chamber have prohibited its use. However, recent advances in laser designs, the use of fiber optics, and studies of multi-pulse laser formats3 have renewed the interest of rocket designers in this state-of the-art technology which offers the potential elimination of torch igniter systems and their associated mechanical parts, as well as toxic hypergolic ignition systems. In support of this interest to develop an alternative ignition system that meets the risk-reduction demands of Next Generation Launch Technology (NGLT), characterization studies of a dual pulse laser format for laser-induced spark ignition are underway at MSFC. Results obtained at MSFC indicate that a dual pulse format can produce plasmas that absorb the laser energy as efficiently as a single pulse format, yet provide a longer plasma lifetime. In an experiments with lean H2/air propellants, the dual pulse laser format, containing the same total energy of a single laser pulse, produced a spark that was superior in its ability to provide sustained ignition of fuel-lean H2/air propellants. The results from these experiments are being used to optimize a dual pulse laser format for future subscale rocket chamber tests. Besides the ignition enhancement, the dual pulse technique provides a practical way to distribute and deliver laser light to the combustion chamber, an important consideration given the limitation of peak power that can be delivered through optical fibers. With this knowledge, scientists and engineers at Los Alamos National Laboratory and CFD Research Corporation have designed and fabricated a miniaturized, first-generation optical prototype of a laser ignition system that could be the basis for a laser ignition system for rocket applications. This prototype will be tested at MSFC in future subscale rocket ignition tests.

Osborne, Robin↗

NASA NTRS1993

Design and fabrication of a hydrogen/oxygen thrust chamber assembly

The combustion system for the Advance Expander Test Bed engine consists of a thrust mount, an injector with a H2-O2 torch igniter, a combustion chamber, and a conical nozzle extension. These components constitute the thrust chamber assembly. The design, fabrication, and testing of these components are described in detail.

Hernacki, Thomas R.↗

NASA NTRS1994

Hydrogen-oxygen Torch Ignitor

The hydrogen-oxygen torch igniter described herein has been successfully used for many years at various NASA Lewis Research Center rocket test facilities to provide ignition for rocket engine research hardware. This igniter is inexpensive, simple to operate, and has demonstrated very good reliability. It has been used as an ignition source for rocket engines that utilized a variety of propellant combinations; some of these engines developed up to 40,000 lb of thrust.

Repas, George A.↗

NASA NTRS1999

Oxygen/Alcohol Dual Thrust RCS Engines

A non-toxic dual thrust RCS engine offers significant operational, safety, and performance advantages to the space shuttle and the next generation RLVs. In this concept, a single engine produces two thrust levels of 25 and 870 lbf. The low thrust level is provided by the spark torch igniter, which, with the addition of 2 extra valves, can also be made to function as a vernier. A dual thrust RCS engine allows 38 verniers to be packaged more efficiently on a vehicle. These 38 vemiers improve translation and reduce cross coupling, thereby providing more pure roll, pitch, and yaw maneuvers of the vehicle. Compared to the 6 vemiers currently on the shuttle, the 38 dual thrust engines would be 25 to 40% more efficient for the same maneuvers and attitude control. The vernier thrust level also reduces plume impingement and contamination concerns. Redundancy is also improved, thereby improving mission success reliability. Oxygen and ethanol are benign propellants which do not create explosive reaction products or contamination, as compared to hypergolic propellants. These characteristics make dual-thrust engines simpler to implement on a non-toxic reaction control system. Tests at WSTF in August 1999 demonstrated a dual-thrust concept that is successful with oxygen and ethanol. Over a variety of inlet pressures and mixture ratios at 22:1 area ratio, the engine produced between 230 and 297 sec Isp, and thrust levels from 8 lbf. to 50 lbf. This paper describes the benefits of dual-thrust engines and the recent results from tests at WSTF.

Angstadt, Tara↗

NASA NTRS2012

Development and Testing of a Methane/Oxygen Catalytic Microtube Ignition System for Rocket Propulsion

This study sought to develop a catalytic ignition advanced torch system with a unique catalyst microtube design that could serve as a low energy alternative or redundant system for the ignition of methane and oxygen rockets. Development and testing of iterations of hardware was carried out to create a system that could operate at altitude and produce a torch. A unique design was created that initiated ignition via the catalyst and then propagated into external staged ignition. This system was able to meet the goals of operating across a range of atmospheric and altitude conditions with power inputs on the order of 20 to 30 watts with chamber pressures and mass flow rates typical of comparable ignition systems for a 100 lbf engine.

Deans, Matthew↗

NASA NTRS2012

Development and Testing of a Methane/Oxygen Catalytic Microtube Ignition System for Rocket Propulsion

Deans, Matthew C.↗

NASA NTRS2009

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Collins, Jacob↗

DOE OSTI2024

Case study of UAS ignition of prescribed fire in a mixedwood on the William B. Bankhead National Forest, Alabama

Abstract Background For at least four decades, practitioners have recognized advantages of aerial versus ground ignition for maximizing the effectiveness of prescribed fires. For example, larger areas can be ignited in less time, or ignition energy may be variously targeted over an area in accordance with the uneven distribution of fuels. The maturation of wireless communication, geopositioning systems, and unmanned aerial systems (UAS) has enhanced those advantages, and UAS approaches also provide further advantages relative to helicopter ignitions, such as reduced risk to human safety, lower operating costs, and higher operational flexibility. In a long running study at the Bankhead National Forest in northcentral Alabama, prescribed fire has been used for nearly 20 years. Most of the burns have been hand-ignited via drip torches, while some have been aerially ignited via helicopter. In March 2022, for the first time, a UAS was used to ignite prescribed fires across a landscape that included a long-term research stand. This field note relates comparisons of both fire behavior and fuel consumption metrics for the UAS-ignited burn versus previous burns on the same stand, and versus burns of other research stands in the same year. Results The UAS-ignited prescribed fire experienced burn effects similar to those from ground-ignited prescribed fires on the same stand in previous years, as well as those from ground-ignited prescribed fires on other stands in the same year. Conclusion This post hoc analysis suggests that UAS ignition approaches may be sufficient for achieving prescribed burn goals, thereby enabling practitioners to realize the advantages offered by that ignition mode.

Environmental Sciences & Ecology↗

NASA NTRS1973

Integrated thruster assembly program

The program is reported which has provided technology for a long life, high performing, integrated ACPS thruster assembly suitable for use in 100 typical flights of a space shuttle vehicle over a ten year period. The four integrated thruster assemblies (ITA) fabricated consisted of: propellant injector; a capacitive discharge, air gap torch type igniter assembly; fast response igniter and main propellant valves; and a combined regen-dump film cooled chamber. These flightweight 6672 N (1500 lb) thruster assemblies employed GH2/GO2 as propellants at a chamber pressure of 207 N/sq cm (300 psia). Test data were obtained on thrusted performance, thermal and hydraulic characteristics, dynamic response in pulsing, and cycle life. One thruster was fired in excess of 42,000 times.

Source record↗