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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 19 records
All sourcesNASA NTRSDOE OSTI
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 NTRS1999

X-33 Combustion-Wave Ignition System Tested

The NASA Lewis Research Center, in cooperation with Rocketdyne, the Boeing Company, tested a novel rocket engine ignition system, called the combustion-wave ignition system, in its Research Combustion Laboratory. This ignition system greatly simplifies ignition in rocket engines that have a large number of combustors. The particular system tested was designed and fabricated by Rocketdyne for the national experimental spacecraft, X-33, which uses Rocketdyne s aerospike rocket engines. The goal of the tests was to verify the system design and define its operational characteristics. Results will contribute to the eventual successful flight of X-33. Furthermore, the combustion-wave ignition system, after it is better understood and refined on the basis of the test results and, later, flight-proven onboard X-33, could become an important candidate engine ignition system for our Nation s next-generation reusable launch vehicle.

Liou, Larry C.↗
NASA NTRS1971

Advanced ignition systems

Low cost, lightweight ignition system for hydrogen oxygen engine system incorporating multicombustor thrust chambers

Source record↗
NASA NTRS2002

Technology Development of a Fiber Optic-Coupled Laser Ignition System for Multi-Combustor Rocket Engines

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). The first two years of the project focus on comprehensive assessments and evaluations of a novel dual-pulse laser concept, flight- qualified laser system, and the technology required to integrate the laser ignition system to a rocket chamber. With collaborations of the Department of Energy/Los Alamos National Laboratory (LANL) and CFD Research Corporation (CFDRC), MSFC has conducted 26 hot fire ignition tests with lab-scale laser systems. These tests demonstrate the concept feasibility of dual-pulse laser ignition to initiate gaseous oxygen (GOX)/liquid kerosene (RP-1) combustion in a rocket chamber. Presently, a fiber optic- coupled miniaturized laser ignition prototype is being implemented at the rocket chamber test rig for future testing. Future work is guided by a technology road map that outlines the work required for maturing a laser ignition system. This road map defines activities for the next six years, with the goal of developing a flight-ready laser ignition system.

Trinh, Huu P.↗
NASA NTRS1977

A sustained-arc ignition system for internal combustion engines

A sustained-arc ignition system was developed for internal combustion engines. It produces a very-long-duration ignition pulse with an energy in the order of 100 millijoules. The ignition pulse waveform can be controlled to predetermined actual ignition requirements. The design of the sustained-arc ignition system is presented in the report.

Birchenough, A. G.↗
NASA NTRS2013

Development and Testing of a Green Monopropellant Ignition System

This paper will detail the development and testing of a "green" monopropellant booster ignition system. The proposed booster ignition technology eliminates the need for a pre-heated catalyst bed, a high wattage power source, toxic pyrophoric ignition fluids, or a bi-propellant spark ignitor. The design offers the simplicity of a monopropellant feed system features non-hazardous gaseous oxygen (GOX) as the working fluid. The approach is fundamentally different from all other "green propellant" solutions in the aerospace in the industry. Although the proposed system is more correctly a "hybrid" rocket technology, since only a single propellant feed path is required, it retains all the simple features of a monopropellant system. The technology is based on the principle of seeding an oxidizing flow with a small amount of hydrocarbon.1 The ignition is initiated electrostatically with a low-wattage inductive spark. Combustion gas byproducts from the hydrocarbon-seeding ignition process can exceed 2400 C and the high exhaust temperature ensures reliable main propellant ignition. The system design is described in detail in the Hydrocarbon-Seeded Ignition System Design subsection.

Whitmore, Stephen A.↗
NASA NTRS1978

Space Shuttle SRM Ignition System

This paper presents the Space Shuttle SRM Ignition System, which consists of a large solid propellant main igniter, a small solid propellant initiating igniter and an electromechanical safety and arming device containing two NASA Standard Initiators and a B-KNO3 pyrotechnic booster charge. In development motors, the igniter also has a valve through which CO2 is injected for post-firing quench of the SRM. The igniter has redundant, testable seals at all pressurized joints and three major reusable components; the case, the adapter, and the S&A device. Two development problem areas are discussed. One problem area was transverse mode combustion instability in the main igniter with maximum amplitude of 340 psi peak-to-peak at a frequency of 1500 Hz, which was reduced by a propellant grain configuration change and a change from a 2% aluminum content propellant to a formulation containing 10% aluminum. The other problem area was an excessively rapid rise of thrust in the SRM, which was reduced by reducing the igniter mass flow rate. This mass flow rate reduction was accomplished by removing portions of the grain starpoints in the head end.

Bolieau, C. W.↗
NASA NTRS1972

A fluidic sounding rocket motor ignition system.

Fluidic sounding rocket motor ignition has been found to be feasible using a system without stored energy and with the complete absence of electrical energy and wiring. The fluidic ignitor is based on a two component aerodynamic resonance heating device called the pneumatic match. Temperatures in excess of 600 C were generated in closed resonance tubes which were excited by a free air jet from a simple convergent nozzle. Using nitrocellulose interface material, ignition of boron potassium nitrate was accomplished with air supply pressures as low as 45 psi. This paper describes an analytical and experimental program which established a fluidic rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source.

Marchese, V. P.↗
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

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 Ibf engine.

Deans, Matthew C.↗
NASA NTRS2002

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). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the DPLIS concept having two main advantages over existing laser ignition concepts. First, the DPLIS can be potentially optimized its laser pulse format to maximize the initial plasma volume, the plasma lifetime, as well as the flame kernel growth rate. Characterization studies of the laser pulse format are now underway with experiments of igniting gaseous hydrogen/air in a Hencken burner. Once ignition is achieved, the flame is open to the atmosphere. This open environment allows easy access for diagnostics of the ignition phenomenon. The quick turn-around time of conducting experiments on this burner make it more amenable for conducting a large number of experiments for statistical analysis of the sensitivity of the test parameters. The results from these experiments will help optimize the laser format for future testing in an H2/O2 subscale rocket chamber.

Trinh, Huu P.↗
NASA NTRS1981

Propulsion system ignition overpressure for the Space Shuttle

Liquid and solid rocket motor propulsion systems create an overpressure wave during ignition, caused by the accelerating gas particles pushing against or displacing the air contained in the launch pad or launch facility and by the afterburning of the fuel-rich gases. This wave behaves as a blast or shock wave characterized by a positive triangular-shaped first pulse and a negative half-sine wave second pulse. The pulse travels up the space vehicle and has the potential of either overloading individual elements or exciting overall vehicle dynamics. The latter effect results from the phasing difference of the wave from one side of the vehicle to the other. This overpressure phasing, or delta P environment, because of its frequency content as well as amplitude, becomes a design driver for certain panels (e.g., thermal shields) and payloads for the Space Shuttle. The history of overpressure effects on the Space Shuttle, the basic overpressure phenomenon, Space Shuttle overpressure environment, scale model overpressure testing, and techniques for suppressing the overpressure environments are considered.

Ryan, R. S.↗