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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 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Ignition and combustion characteristics of metallized propellants, phase 2

Experimental and analytical investigations focusing on aluminum/hydrocarbon gel droplet secondary atomization and its effects on gel-fueled rocket engine performance are being conducted. A single laser sheet sizing/velocimetry diagnostic technique, which should eliminate sizing bias in the data collection process, has been designed and constructed to overcome limitations of the two-color forward-scatter technique used in previous work. Calibration of this system is in progress and the data acquisition/validation code is being written. Narrow-band measurements of radiant emission, discussed in previous reports, will be used to determine if aluminum ignition has occurred in a gel droplet. A one-dimensional model of a gel-fueled rocket combustion chamber, described in earlier reports, has been exercised in conjunction with a two-dimensional, two-phase nozzle code to predict the performance of an aluminum/hydrocarbon fueled engine. Estimated secondary atomization effects on propellant burnout distance, condensed particle radiation losses to the chamber walls, and nozzle two phase flow losses are also investigated. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size, and radiation heat losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated, depending on secondary atomization intensity. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two phase flow effects on overall engine performance. Radiation losses yielded a one percent decrease in engine Isp. Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine Isp was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine Isp efficiencies, accounting for radiation and two phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of five.

Mueller, D. C.↗
NASA NTRS1994

New rubber qualification for the igniter adapter

Kirkhill Rubber Company (KRC) has informed Thiokol Corporation that two raw materials used in the asbestos and silica filled acrylonitrile butadience rubber (NBR) formulation per STW 2621 are no longer available from their vendors. Agerite White (Di-beta-naphthyl-paraphenylene diamine), manufactured by B. F. Goodrich, is an antioxidant used in NBR. This raw material makes up roughly 1-2 percent of the finished product. KRC proposed that this raw material be replaced by Agerite Stalite S (mixture of octylated diphenylamines) distributed by R. T. Vanderbilt Co. Protox-166 zinc oxide, manufactured by Zinc Corporation of America, is an activator currently used in NBR. This material also makes up about 1-2 percent of the finished material. Protox-166 is an American process grade zinc oxide. It is proposed by KRC to replace Protox-166 with Kadox-930C, a French process grade zinc oxide. American process grades have an ASTM minimum purity of 99.0 percent; the French process grades have a minimum purity of 99.5 percent. Previous testing per WTP-0270 has demonstrated that the mechanical and thermal properties of the rubber with the new ingredients are comparable to the 'old' rubber. The test results are reported in TWR-61790. One igniter adapter, Part no. 7U77562-02 serial no. 2 was insulated per ETP-1206 using the new rubber formulation and a modified lay up and cure method to demonstrate that there is no impact on this process. The results of this demonstration are reported.

Humpherys, Mark A.↗
NASA NTRS1994

Ignition and combustion characteristics of metallized propellants, phase 2

Secondary atomization and ignition characteristics of aluminum/hydrocarbon gel propellants were investigated. Models of gel droplet shell formation were applied to aluminum/liquid hydrocarbon propellants to examine the effects of solid loading and ultimate particle size on the minimum droplet diameter permitting secondary atomization. A one-dimensional model of a gel-fueled rocket combustion chamber was developed. A model for radiant heat transfer from hot aluminum oxide particles to the chamber walls is included. A two-dimensional, two-phase nozzle code was used to estimate nozzle two-phase losses and overall engine performance.

Mueller, D. C.↗
NASA NTRS1995

Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

Solid State Laser Initiated Ordnance (LIO) offers new technology having potential for enhanced safety, reduced costs, and improved operational efficiency. Concerns over the absence of programmatic applications of the technology, which has prevented acceptance by flight programs, should be abated since LIO has now been operationally implemented by the Laser Initiated Ordnance Sounding Rocket Demonstration (LOSRD) Program. The first launch of solid state laser diode LIO at the NASA Wallops Flight Facility (WFF) occurred on March 15, 1995 with all mission objectives accomplished. This project, Phase 3 of a series of three NASA Headquarters LIO demonstration initiatives, accomplished its objective by the flight of a dedicated, all-LIO sounding rocket mission using a two-stage Nike-Orion launch vehicle. LIO flight hardware, made by The Ensign-Bickford Company under NASA's first Cooperative Agreement with Profit Making Organizations, safely initiated three demanding pyrotechnic sequence events, namely, solid rocket motor ignition from the ground and in flight, and flight termination, i.e., as a Flight Termination System (FTS). A flight LIO system was designed, built, tested, and flown to support the objectives of quickly and inexpensively putting LIO through ground and flight operational paces. The hardware was fully qualified for this mission, including component testing as well as a full-scale system test. The launch accomplished all mission objectives in less than 11 months from proposal receipt. This paper concentrates on accomplishments of the ordnance aspects of the program and on the program's implementation and results. While this program does not generically qualify LIO for all applications, it demonstrated the safety, technical, and operational feasibility of those two most demanding applications, using an all solid state safe and arm system in critical flight applications.

Schulze, Norman R.↗
NASA NTRS1995

Research on ignition and flame spread of solid materials in Japan

Fire safety is one of the main concerns for crewed missions such as the space station. Materials used in spacecraft may burn even if metalic. There are severe restrictions on the materials used in spacecraft from the view of fire safety. However, such restrictions or safety standards are usually determined based on experimental results under normal gravity, despite large differences between the phenomena under normal and microgravity. To evaluate the appropriateness of materials for use in space, large amount of microgravity fire-safety combustion data is urgently needed. Solid material combustion under microgravity, such as ignition and flame spread, is a relatively new research field in Japan. As the other reports in this workshop describe, most of microgravity combustion research in Japan is droplet combustion as well as some research on gas phase combustion. Since JAMIC, the Japan Microgravity Center, (which offers 10 seconds microgravity time) opened in 1992, microgravity combustion research is robust, and many drop tests relating to solid combustion (paper combustion, cotton string combustion, metal combustion with Aluminium or Magnesium) have been performed. These tests proved that the 10 seconds of microgravity time at JAMIC is useful for solid combustion research. Some experiments were performed before JAMIC opened. For example, latticed paper was burned under microgravity by using a 50 m drop tower to simulate porous material combustion under microgravity. A 50 m tower provides only 2 seconds microgravity time however, and it was not long enough to investigate the solid combustion phenomena.

Ito, Kenichi↗
NASA NTRS1994

Development and qualification testing of a laser-ignited, all-secondary (DDT) detonator

The Indian Head Division, Naval Surface Warfare Center (IHDIV, NSWC) is conducting a qualification program for a laser-ignited, all-secondary (DDT) explosive detonator. This detonator was developed jointly by IHDIV, NSWC and the Department of Energy's EG&G Mound Applied Technologies facility in Miamisburg, Ohio to accept a laser initiation signal and produce a fully developed shock wave output. The detonator performance requirements were established by the on-going IHDIV, NSWC Laser Initiated Transfer Energy Subsystem (LITES) advanced development program. Qualification of the detonator as a component utilizing existing military specifications is the selected approach for this program. The detonator is a deflagration-to-detonator transfer (DDT) device using a secondary explosive, HMX, to generate the required shock wave output. The prototype development and initial system integration tests for the LITES and for the detonator were reported at the 1992 International Pyrotechnics Society Symposium and at the 1992 Survival and Flight Equipment National Symposium. Recent results are presented for the all-fire sensitivity and qualification tests conducted at two different laser initiation pulses.

Blachowski, Thomas J.↗
NASA NTRS1997

Hot Fire Ignition Test with Densified Liquid Hydrogen using a RL10B-2 Cryogenic H2/O2 Rocket Engine

Enhancements to propellants provide an opportunity to either increase performance of an existing vehicle, or reduce the size of a new vehicle. In the late 1980's the National AeroSpace Plane (NASP) reopened the technology chapter on densified propellants, in particular hydrogen. Since that point in time the NASA Lewis Research Center (LERC) in Cleveland, Ohio has been leading the way to provide critical research on the production and transfer of densified propellants. On October 4, 1996 NASA LeRC provided another key demonstration towards the advancement of densified propellants as a viable fuel. Successful ignition of an RL10B-2 engine was achieved with near triple point liquid hydrogen.

McNelis, Nancy B.↗
NASA NTRS1961

Magnetic Ignition of Pulsed Gas Discharges in Air of Low Pressure in a Coaxial Plasma Gun

The effect of an axial magnetic field on the breakdown voltage of a coaxial system of electrodes has been investigated by earlier workers. For low values of gas pressure times electrode spacing, the breakdown voltage is decreased by the application of the magnetic field. The electron cyclotron radius now assumes the role held by the mean free path in nonmagnetic discharges and the breakdown voltage becomes a function of the magnetic flux density. In this paper the dependence of the formative time lag as a function of the magnetic flux density is established and the feasibility of using a magnetic field for igniting high-voltage, high-current discharges is shown through theory and experiment. With a 36 microfarad capacitor bank charged to 48,000 volts, a peak current of 1.3 x 10( exp 6) amperes in a coaxial type of plasma gun was achieved with a current rise time of only 2 microseconds.

Thom, Karlheinz↗
NASA NTRS2000

Predictability of Gaseous Impact Ignition Sensitivity from Autoignition Temperature Data

An analysis of cumulative binomial probabilities for a quantal-type test like ASTM G74 indicates that for a probability of reaction in a single trial of 0.05, the probability of obtaining zero reactions in 20 trials is only approximately 36 percent, Because the lack of precision with the ASTM G74 test logic could be potentially misleading, an investigation was conducted to determine if pneumatic impact sensitivity could be predicted more reliably from other, more precise test methods. Pneumatic impact sensitivity tests were conducted following ASTM G74, but employing more rigorous test logics; autoignition temperature tests were conducted per ASTM G72. Polytetrafluoroethylene (PTFE), Viton A, polyethylene, polyoxymethylene polyamide, and silicone were the polymers used. After a fixed number of pneumatic impact tests was performed at set pressures, a plot of reaction frequency over a 3.4 to 24.1 MPa impact pressure range indicated that PTFE, Viton A, and silicone would outperform other polymers. The 50 percent pneumatic impact reactivity level was determined with the Bruceton test method. The data obtained resulted in a strong correlation between the 50 percent reactivity level and the autoignition temperature of polymers, indicating similar ignition mechanisms.

Hirsh, David↗
NASA NTRS2001

Rocket Launch-Induced Vibration and Ignition Overpressure Response

Rocket-induced vibration and ignition overpressure response environments are predicted in the low-frequency (5 to 200 hertz) range. The predictions are necessary to evaluate their impact on critical components, structures, and facilities in the immediate vicinity of the rocket launch pad.

Caimi, Raoul E.↗
NASA NTRS2001

Three-Dimensional Ignition and Flame Propagation Above Liquid Fuel Pools: Computational Analysis

A three-dimensional unsteady reactive Navier-Stokes code is developed to study the ignition and flame spread above liquid fuels initially below the flashpoint temperature. Opposed air flow to the flame spread due to forced and/or natural convection is considered. Pools of finite width and length are studied in air channels of prescribed height and width. Three-dimensional effects of the flame front near the edge of the pool are captured in the computation. The formation of a recirculation zone in the gas phase similar to that found in two-dimensional calculations is also present in the three-dimensional calculations. Both uniform spread and pulsating spread modes are found in the calculated results.

Cai, Jinsheng↗
NASA NTRS2001

Rocket Launch-Induced Vibration and Ignition Overpressure Response

Rocket-induced vibration and ignition overpressure response environments are predicted in the low-frequency (5 to 200 hertz) range. The predictions are necessary to evaluate their impact on critical components, structures, and facilities in the immediate vicinity of the rocket launch pad.

Caimi, Raoul↗
NASA NTRS2002

Review of Air Vitiation Effects on Scramjet Ignition and Flameholding Combustion Processes

This paper offers a detailed review and analysis of more than 100 papers on the physics and chemistry of scramjet ignition and flameholding combustion processes, and the known effects of air vitiation on these processes. The paper attempts to explain vitiation effects in terms of known chemical kinetics and flame propagation phenomena. Scaling methodology is also examined, and a highly simplified Damkoehler scaling technique based on OH radical production/destruction is developed to extrapolate ground test results, affected by vitiation, to flight testing conditions. The long term goal of this effort is to help provide effective means for extrapolating ground test data to flight, and thus to reduce the time and expense of both ground and flight testing.

Pellett, G. L.↗
NASA NTRS2003

Measuring the Internal Environment of Solid Rocket Motors During Ignition

A new instrumentation system has been developed to measure the internal environment of solid rocket test motors during motor ignition. The system leverages conventional, analog gages with custom designed, electronics modules to provide safe, accurate, high speed data acquisition capability. To date, the instrumentation system has been demonstrated in a laboratory environment and on subscale static fire test motors ranging in size from 5-inches to 24-inches in diameter. Ultimately, this system is intended to be installed on a full-scale Reusable Solid Rocket Motor. This paper explains the need for the data, the components and capabilities of the system, and the test results.

Weisenberg, Brent↗
NASA NTRS1947

Practical Possibilities of High-Altitude Flight with Exhaust-Gas Turbines in Connection with Spark Ignition Engines Comparative Thermodynamic and Flight Mechanical Investigations

As a means of preparing for high-altitude flight with spark-ignition engines in conjunction with exhaust-gas turbosuperchargers, various methods of modifying the exhaust-gas temperatures, which are initially higher than a turbine can withstand are mathematically compared. The thermodynamic results first obtained are then examined with respect to the effect on flight speed, climbing speed, ceiling, economy, and cruising range. The results are so presented in a generalized form that they may be applied to every appropriate type of aircraft design and a comparison with the supercharged engine without exhaust-gas turbine can be made.

Weise, A.↗