Coil-On-Plug Ignition for LOX/Methane Liquid Rocket Engines in Thermal Vacuum Environments

NASA Technical Reports Server (NTRS)

Melcher, John C.; Atwell, Matthew J.; Morehead, Robert L.; Hurlbert, Eric A.; Bugarin, Luz; Chaidez, Mariana

2017-01-01

A coil-on-plug ignition system has been developed and tested for Liquid Oxygen (LOX) / liquid methane rocket engines operating in thermal vacuum conditions. The igniters were developed and tested as part of the Integrated Cryogenic Propulsion Test Article (ICPTA), previously tested as part of the Project Morpheus test vehicle. The ICPTA uses an integrated, pressure-fed, cryogenic LOX/methane propulsion system including a reaction control system (RCS) and a main engine. The ICPTA was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. In order to successfully demonstrate ignition reliability in the vacuum conditions and eliminate corona discharge issues, a coil-on-plug ignition system has been developed. The ICPTA uses spark-plug ignition for both the main engine igniter and the RCS. The coil-on-plug configuration eliminates the conventional high-voltage spark plug cable by combining the coil and the spark-plug into a single component. Prior to ICPTA testing at Plum Brook, component-level reaction control engine (RCE) and main engine igniter testing was conducted at NASA Johnson Space Center (JSC), which demonstrated successful hot-fire ignition using the coil-on-plug from sea-level ambient conditions down to 10(exp.-2) torr. Integrated vehicle hot-fire testing at JSC demonstrated electrical and command/data system performance. Lastly, Plum Brook testing demonstrated successful ignitions at simulated altitude conditions at 30 torr and cold thermal-vacuum conditions at 6 torr. The test campaign successfully proved that coil-on-plug technology will enable integrated LOX/methane propulsion systems in future spacecraft.

Very Low Thrust Gaseous Oxygen-hydrogen Rocket Engine Ignition Technology

NASA Technical Reports Server (NTRS)

Bjorklund, Roy A.

1983-01-01

An experimental program was performed to determine the minimum energy per spark for reliable and repeatable ignition of gaseous oxygen (GO2) and gaseous hydrogen (GH2) in very low thrust 0.44 to 2.22-N (0.10 to 0.50-lb sub f) rocket engines or spacecraft and satellite attitude control systems (ACS) application. Initially, the testing was conducted at ambient conditions, with the results subsequently verified under vacuum conditions. An experimental breadboard electrical exciter that delivered 0.2 to 0.3 mj per spark was developed and demonstrated by repeated ignitions of a 2.22-N (0.50-lb sub f) thruster in a vacuum chamber with test durations up to 30 min.

Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

NASA Technical Reports Server (NTRS)

Gregory, John W.; Straight, David M.

1961-01-01

Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

Orbit transfer rocket engine technology program: Oxygen materials compatibility testing

NASA Technical Reports Server (NTRS)

Schoenman, Leonard

1989-01-01

Particle impact and frictional heating tests of metals in high pressure oxygen, are conducted in support of the design of an advanced rocket engine oxygen turbopump. Materials having a wide range of thermodynamic properties including heat of combustion and thermal diffusivity were compared in their resistance to ignition and sustained burning. Copper, nickel and their alloys were found superior to iron based and stainless steel alloys. Some materials became more difficult to ignite as oxygen pressure was increased from 7 to 21 MPa (1000 to 3000 psia).

Breakdown voltage determination of gaseous and near cryogenic fluids with application to rocket engine ignition

NASA Astrophysics Data System (ADS)

Nugent, Nicholas Jeremy

Liquid rocket engines extensively use spark-initiated torch igniters for ignition. As the focus shifts to longer missions that require multiple starts of the main engines, there exists a need to solve the significant problems associated with using spark-initiated devices. Improving the fundamental understanding of predicting the required breakdown voltage in rocket environments along with reducing electrical noise is necessary to ensure that missions can be completed successfully. To better understand spark ignition systems and add to the fundamental research on spark development in rocket applications, several parameter categories of interest were hypothesized to affect breakdown voltage: (i) fluid, (ii) electrode, and (iii) electrical. The fluid properties varied were pressure, temperature, density and mass flow rate. Electrode materials, insert electrode angle and spark gap distance were the electrode properties varied. Polarity was the electrical property investigated. Testing how breakdown voltage is affected by each parameter was conducted using three different isolated insert electrodes fabricated from copper and nickel. A spark plug commonly used in torch igniters was the other electrode. A continuous output power source connected to a large impedance source and capacitance provided the pulsing potential. Temperature, pressure and high voltage measurements were recorded for the 418 tests that were successfully completed. Nitrogen, being inert and similar to oxygen, a propellant widely used in torch igniters, was used as the fluid for the majority of testing. There were 68 tests completed with oxygen and 45 with helium. A regression of the nitrogen data produced a correction coefficient to Paschen's Law that predicts the breakdown voltage to within 3000 volts, better than 20%, compared to an over prediction on the order of 100,000 volts using Paschen's Law. The correction coefficient is based on the parameters most influencing breakdown voltage: fluid density, spark gap distance, electrode angles, electrode materials and polarity. The research added to the fundamental knowledge of spark development in rocket ignition applications by determining the parameters that most influence breakdown voltage. Some improvements to the research should include better temperature measurements near the spark gap, additional testing with oxygen and testing with fuels of interest such as hydrogen and methane.

Supercomputer modeling of hydrogen combustion in rocket engines

NASA Astrophysics Data System (ADS)

Betelin, V. B.; Nikitin, V. F.; Altukhov, D. I.; Dushin, V. R.; Koo, Jaye

2013-08-01

Hydrogen being an ecological fuel is very attractive now for rocket engines designers. However, peculiarities of hydrogen combustion kinetics, the presence of zones of inverse dependence of reaction rate on pressure, etc. prevents from using hydrogen engines in all stages not being supported by other types of engines, which often brings the ecological gains back to zero from using hydrogen. Computer aided design of new effective and clean hydrogen engines needs mathematical tools for supercomputer modeling of hydrogen-oxygen components mixing and combustion in rocket engines. The paper presents the results of developing verification and validation of mathematical model making it possible to simulate unsteady processes of ignition and combustion in rocket engines.

Saturn Apollo Program

NASA Image and Video Library

1965-07-10

Marshall Space Flight Center's rocket development has always included component testing. Pictured here is a Cell 114-B burn stack. The C114-B is part of the gas generators used to test heat exchanges for the F-1 engine. On the initial firing of the C114-B the spark ignition would not light. The rocket propellant mixed with the liquid oxygen gelled creating a bomb. After several attempts at ignition, the spark ignited and blew up the stand. Subsequent testings were completed on newly constructed stands and no further mishaps were reported.

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

NASA Technical Reports Server (NTRS)

Deans, Matthew

2012-01-01

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.

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

NASA Technical Reports Server (NTRS)

Deans, Matthew C.; Schneider, Steven J.

2012-01-01

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.

Delta II Mars Pathfinder

NASA Technical Reports Server (NTRS)

1998-01-01

Final preparations for lift off of the DELTA II Mars Pathfinder Rocket are shown. Activities include loading the liquid oxygen, completing the construction of the Rover, and placing the Rover into the Lander. After the countdown, important visual events include the launch of the Delta Rocket, burnout and separation of the three Solid Rocket Boosters, and the main engine cutoff. The cutoff of the main engine marks the beginning of the second stage engine. After the completion of the second stage, the third stage engine ignites and then cuts off. Once the third stage engine cuts off spacecraft separation occurs.

Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Thomas, Matthew E.; Bossard, John A.; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

2001-01-01

The fiber optically coupled laser ignition approach summarized is under consideration for use in igniting bi-propellant rocket thrust chambers. This laser ignition approach is based on a novel dual pulse format capable of effectively increasing laser generated plasma life times up to 1000 % over conventional laser ignition methods. In the dual-pulse format tinder consideration here an initial laser pulse is used to generate a small plasma kernel. A second laser pulse that effectively irradiates the plasma kernel follows this pulse. Energy transfer into the kernel is much more efficient because of its absorption characteristics thereby allowing the kernel to develop into a much more effective ignition source for subsequent combustion processes. In this research effort both single and dual-pulse formats were evaluated in a small testbed rocket thrust chamber. The rocket chamber was designed to evaluate several bipropellant combinations. Optical access to the chamber was provided through small sapphire windows. Test results from gaseous oxygen (GOx) and RP-1 propellants are presented here. Several variables were evaluated during the test program, including spark location, pulse timing, and relative pulse energy. These variables were evaluated in an effort to identify the conditions in which laser ignition of bi-propellants is feasible. Preliminary results and analysis indicate that this laser ignition approach may provide superior ignition performance relative to squib and torch igniters, while simultaneously eliminating some of the logistical issues associated with these systems. Further research focused on enhancing the system robustness, multiplexing, and window durability/cleaning and fiber optic enhancements is in progress.

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/ Hydrogen Rocket Engine

DTIC Science & Technology

1993-10-12

achieved on the same test. Figure 5-1 is a photograph of the RS-44 ICE engine as it achieved mainstage 3 operation on test 87-017-006. Nominal and...engine transition to mainstage main chamber pressure, 3 Igniter operation satisfactory 87-017-004 1.09 Igniter operation / engine Redline cutoff due...ok, transition stage satisfactory, 3 maInstage operation achieved, all other objectives met. 3 I I I -56- 1 CR 194443 Figure 7-22 RS-44 ENGtNE

Ignition Delay Experiments with Small-scale Rocket Engine at Simulated Altitude Conditions Using Various Fuels with Nitric Acid Oxidants / Dezso J. Ladanyi

NASA Technical Reports Server (NTRS)

Ladanyi, Dezso J

1952-01-01

Ignition delay determinations of several fuels with nitric oxidants were made at simulated altitude conditions utilizing a small-scale rocket engine of approximately 50 pounds thrust. Included in the fuels were aniline, hydrazine hydrate, furfuryl alcohol, furfuryl mercaptan, turpentine, and mixtures of triethylamine with mixed xylidines and diallyaniline. Red fuming, white fuming, and anhydrous nitric acids were used with and without additives. A diallylaniline - triethylamine mixture and a red fuming nitric acid analyzing 3.5 percent water and 16 percent NO2 by weight was found to have a wide temperature-pressure ignition range, yielding average delays from 13 milliseconds at 110 degrees F to 55 milliseconds at -95 degrees F regardless of the initial ambient pressure that ranged from sea-level pressure altitude of 94,000 feet.

Combustion dynamics in cryogenic rocket engines: Research programme at DLR Lampoldshausen

NASA Astrophysics Data System (ADS)

Hardi, Justin S.; Traudt, Tobias; Bombardieri, Cristiano; Börner, Michael; Beinke, Scott K.; Armbruster, Wolfgang; Nicolas Blanco, P.; Tonti, Federica; Suslov, Dmitry; Dally, Bassam; Oschwald, Michael

2018-06-01

The Combustion Dynamics group in the Rocket Propulsion Department at the German Aerospace Center (DLR), Lampoldshausen, strives to advance the understanding of dynamic processes in cryogenic rocket engines. Leveraging the test facilities and experimental expertise at DLR Lampoldshausen, the group has taken a primarily experimental approach to investigating transient flows, ignition, and combustion instabilities for over one and a half decades. This article provides a summary of recent achievements, and an overview of current and planned research activities.

Orbital transfer rocket engine technology program: Soft wear ring seal technology

NASA Technical Reports Server (NTRS)

Lariviere, Brian W.

1992-01-01

Liquid oxygen (LOX) compatibility tests, including autogenous ignition, promoted ignition, LOX impact tests, and friction and wear tests on different PV products were conducted for several polymer materials as verification for the implementation of soft wear ring seals in advanced rocket engine turbopumps. Thermoplastics, polyimide based materials, and polyimide-imide base materials were compared for oxygen compatibility, specific wear coefficient, wear debris production, and heat dissipation mechanisms. A thermal model was generated that simulated the frictional heating input and calculated the surface temperature and temperature distribution within the seal. The predictions were compared against measured values. Heat loads in the model were varied to better match the test data and determine the difference between the measured and the calculated coefficients of friction.

Extended temperature range ACPS thruster investigation

NASA Technical Reports Server (NTRS)

Blubaugh, A. L.; Schoenman, L.

1974-01-01

The successful hot fire demonstration of a pulsing liquid hydrogen/liquid oxygen and gaseous hydrogen/liquid oxygen attitude control propulsion system thruster is described. The test was the result of research to develop a simple, lightweight, and high performance reaction control system without the traditional requirements for extensive periods of engine thermal conditioning, or the use of complex equipment to convert both liquid propellants to gas prior to delivery to the engine. Significant departures from conventional injector design practice were employed to achieve an operable design. The work discussed includes thermal and injector manifold priming analyses, subscale injector chilldown tests, and 168 full scale and 550 N (1250 lbF) rocket engine tests. Ignition experiments, at propellant temperatures ranging from cryogenic to ambient, led to the generation of a universal spark ignition system which can reliably ignite an engine when supplied with liquid, two phase, or gaseous propellants. Electrical power requirements for spark igniter are very low.

Two-phase flow in the cooling circuit of a cryogenic rocket engine

NASA Astrophysics Data System (ADS)

Preclik, D.

1992-07-01

Transient two-phase flow was investigated for the hydrogen cooling circuit of the HM7 rocket engine. The nuclear reactor code ATHLET/THESEUS was adapted to cryogenics and applied to both principal and prototype experiments for validation and simulation purposes. The cooling circuit two-phase flow simulation focused on the hydrogen prechilling and pump transient phase prior to ignition. Both a single- and a multichannel model were designed and employed for a valve leakage flow, a nominal prechilling flow, and a prechilling with a subsequent pump-transient flow. The latter case was performed in order to evaluate the difference between a nominal and a delayed turbo-pump start-up. It was found that an extension of the nominal prechilling sequence in the order of 1 second is sufficient to finally provide for liquid injection conditions of hydrogen which, as commonly known, is undesirable for smooth ignition and engine starting transients.

Test Stand at the Rocket Engine Test Facility

NASA Image and Video Library

1973-02-21

The thrust stand in the Rocket Engine Test Facility at the National Aeronautics and Space Administration (NASA) Lewis Research Center in Cleveland, Ohio. The Rocket Engine Test Facility was constructed in the mid-1950s to expand upon the smaller test cells built a decade before at the Rocket Laboratory. The $2.5-million Rocket Engine Test Facility could test larger hydrogen-fluorine and hydrogen-oxygen rocket thrust chambers with thrust levels up to 20,000 pounds. Test Stand A, seen in this photograph, was designed to fire vertically mounted rocket engines downward. The exhaust passed through an exhaust gas scrubber and muffler before being vented into the atmosphere. Lewis researchers in the early 1970s used the Rocket Engine Test Facility to perform basic research that could be utilized by designers of the Space Shuttle Main Engines. A new electronic ignition system and timer were installed at the facility for these tests. Lewis researchers demonstrated the benefits of ceramic thermal coatings for the engine’s thrust chamber and determined the optimal composite material for the coatings. They compared the thermal-coated thrust chamber to traditional unlined high-temperature thrust chambers. There were more than 17,000 different configurations tested on this stand between 1973 and 1976. The Rocket Engine Test Facility was later designated a National Historic Landmark for its role in the development of liquid hydrogen as a propellant.

Control Room at the NACA’s Rocket Engine Test Facility

NASA Image and Video Library

1957-05-21

Test engineers monitor an engine firing from the control room of the Rocket Engine Test Facility at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Engine Test Facility, built in the early 1950s, had a rocket stand designed to evaluate high-energy propellants and rocket engine designs. The facility was used to study numerous different types of rocket engines including the Pratt and Whitney RL-10 engine for the Centaur rocket and Rocketdyne’s F-1 and J-2 engines for the Saturn rockets. The Rocket Engine Test Facility was built in a ravine at the far end of the laboratory because of its use of the dangerous propellants such as liquid hydrogen and liquid fluorine. The control room was located in a building 1,600 feet north of the test stand to protect the engineers running the tests. The main control and instrument consoles were centrally located in the control room and surrounded by boards controlling and monitoring the major valves, pumps, motors, and actuators. A camera system at the test stand allowed the operators to view the tests, but the researchers were reliant on data recording equipment, sensors, and other devices to provide test data. The facility’s control room was upgraded several times over the years. Programmable logic controllers replaced the electro-mechanical control devices. The new controllers were programed to operate the valves and actuators controlling the fuel, oxidant, and ignition sequence according to a predetermined time schedule.

Aircraft Research and Technology for Antimisting Kerosene Conference, February 18-19, 1981.

DTIC Science & Technology

1981-06-01

carrier turbine aircraft fatal accidents from 1964 through 1976. Since antimisting fuel is intended to inhibit ignition and flame propagation when fuel is...been shown to be possible and rapid, although costly and complex. One item that should be added at this point is in the event turbine engine power...port side with 0.28 percent FM-9 fuel was ignited by the rocket motors. When the turbine engine separated from the wing, localized fire remained with

NASA Engineer Examines the Design of a Regeneratively-Cooled Rocket Engine

NASA Image and Video Library

1958-12-21

An engineer at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a drawing showing the assembly and details of a 20,000-pound thrust regeneratively cooled rocket engine. The engine was being designed for testing in Lewis’ new Rocket Engine Test Facility, which began operating in the fall of 1957. The facility was the largest high-energy test facility in the country that was capable of handling liquid hydrogen and other liquid chemical fuels. The facility’s use of subscale engines up to 20,000 pounds of thrust permitted a cost-effective method of testing engines under various conditions. The Rocket Engine Test Facility was critical to the development of the technology that led to the use of hydrogen as a rocket fuel and the development of lightweight, regeneratively-cooled, hydrogen-fueled rocket engines. Regeneratively-cooled engines use the cryogenic liquid hydrogen as both the propellant and the coolant to prevent the engine from burning up. The fuel was fed through rows of narrow tubes that surrounded the combustion chamber and nozzle before being ignited inside the combustion chamber. The tubes are visible in the liner sitting on the desk. At the time, Pratt and Whitney was designing a 20,000-pound thrust liquid-hydrogen rocket engine, the RL-10. Two RL-10s would be used to power the Centaur second-stage rocket in the 1960s. The successful development of the Centaur rocket and the upper stages of the Saturn V were largely credited to the work carried out Lewis.

Analysis and Evaluation of German Attainments and Research in the Liquid Rocket Engine Field. Volume 8. Rocket Engine Control and Safety Circuits

DTIC Science & Technology

1951-02-01

the pressure switch (16) is activated. This causes the-electrical circuit to open valve (11) and start the igniter (17). The nitrogen pressure...activates the pressure switch (11) at approximately 7 psi before it flows through the Injector (9) into the chamber. ATI-85«’󈧕 - -A 11...precluded. Accordingly, pressure switch (11) is inserted in the system in parallel (electrically) with the flow indicator (17), and the circuit may

Ignition Delays of Alkyl Thiophosphites with White and Red Fuming Nitric Acids Within Temperature Range 80 to -105 F

NASA Technical Reports Server (NTRS)

Miller, Riley O; Ladanyi, Dezso J

1953-01-01

Ignition delays of alkyl thiophosphites were obtained in a modified open-cup apparatus and a small-scale rocket engine apparatus. At -40 F, mixed alkyl thiophosphites gave short delays with white fuming nitric acid containing 2 percent water and red fuming nitric acids of widely varying compositions. At -40 F and higher, triethyl trithiophosphite blended with as much as 40 percent n-heptane gave satisfactory self-igniting properties at temperatures as low as -76 F.

Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Anderson, William E.

2005-01-01

The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

Developing Avionics Hardware and Software for Rocket Engine Testing

NASA Technical Reports Server (NTRS)

Aberg, Bryce Robert

2014-01-01

My summer was spent working as an intern at Kennedy Space Center in the Propulsion Avionics Branch of the NASA Engineering Directorate Avionics Division. The work that I was involved with was part of Rocket University's Project Neo, a small scale liquid rocket engine test bed. I began by learning about the layout of Neo in order to more fully understand what was required of me. I then developed software in LabView to gather and scale data from two flowmeters and integrated that code into the main control software. Next, I developed more LabView code to control an igniter circuit and integrated that into the main software, as well. Throughout the internship, I performed work that mechanics and technicians would do in order to maintain and assemble the engine.

Combustion-wave ignition for rocket engines

NASA Technical Reports Server (NTRS)

Liou, Larry C.

1992-01-01

The combustion wave ignition concept was experimentally studied in order to verify its suitability for application in baffled sections of a large booster engine combustion chamber. Gaseous oxygen/gaseous methane (GOX/GH4) and gaseous oxygen/gaseous hydrogen (GOX/GH2) propellant combinations were evaluated in a subscale combustion wave ignition system. The system included four element tubes capable of carrying ignition energy simultaneously to four locations, simulating four baffled sections. Also, direct ignition of a simulated Main Combustion Chamber (MCC) was performed. Tests were conducted over a range of mixture ratios and tube geometries. Ignition was consistently attained over a wide range of mixture ratios. And at every ignition, the flame propagated through all four element tubes. For GOX/GH4, the ignition system ignited the MCC flow at mixture ratios from 2 to 10 and for GOX/GH2 the ratios is from 2 to 13. The ignition timing was found to be rapid and uniform. The total ignition delay when using the MCC was under 11 ms, with the tube-to-tube, as well as the run-to-run, variation under 1 ms. Tube geometries were found to have negligible effect on the ignition outcome and timing.

LOx / LCH4: A Unifying Technology for Future Exploration

NASA Technical Reports Server (NTRS)

Falker, John; Terrier, Douglas; Clayton, Ronald G.; Banker, Brian; Ryan, Abigail

2015-01-01

Reduced mass due to increasing commonality between spacecraft subsystems such as power and propulsion have been identified as critical to enabling human missions to Mars. This project represents the first ever integrated propulsion and power system testing and lays the foundations for future sounding rocket flight testing, which will yield the first in-space ignition of a LOx / LCH4 rocket engine.

Liquid and gelled sprays for mixing hypergolic propellants using an impinging jet injection system

NASA Astrophysics Data System (ADS)

James, Mark D.

The characteristics of sprays produced by liquid rocket injectors are important in understanding rocket engine ignition and performance. The includes, but is not limited to, drop size distribution, spray density, drop velocity, oscillations in the spray, uniformity of mixing between propellants, and the spatial distribution of drops. Hypergolic ignition and the associated ignition delay times are also important features in rocket engines, providing high reliability and simplicity of the ignition event. The ignition delay time is closely related to the level and speed of mixing between a hypergolic fuel and oxidizer, which makes the injection method and conditions crucial in determining the ignition performance. Although mixing and ignition of liquid hypergolic propellants has been studied for many years, the processes for injection, mixing, and ignition of gelled hypergolic propellants are less understood. Gelled propellants are currently under investigation for use in rocket injectors to combine the advantages of solid and liquid propellants, although not without their own difficulties. A review of hypergolic ignition has been conducted for selected propellants, and methods for achieving ignition have been established. This research is focused on ignition using the liquid drop-on-drop method, as well as the doublet impinging jet injector. The events leading up to ignition, known as pre-ignition stage are discussed. An understanding of desirable ignition and combustion performance requires a study of the effects of injection, temperature, and ambient pressure conditions. A review of unlike-doublet impinging jet injection mixing has also been conducted. This includes mixing factors in reactive and non-reactive sprays. Important mixing factors include jet momentum, jet diameter and length, impingement angle, mass distribution, and injector configuration. An impinging jet injection system is presented using an electro-mechanically driven piston for injecting liquid and gelled hypergolic propellants. A calibration of the system is done with water in preparation for hypergolic injection, and characteristics of individual water and gelled JP-8 jets are studied at velocities in the range of 3 ft/s to 61 ft/s. The piston response is also analyzed to characterize the startup and steady state liquid jet velocities using orifices of 0.02" in diameter. Using this injection system, water and gelled JP-8 sprays are formed and compared across injection velocities of 30 ft/s to 121 ft/s. The comparison includes sheet shape and disintegration, total number of drops, drop size distributions, drop eccentricity, most populated drop bin size, and mean drop sizes. A test matrix for investigating the effects of mixing on ignition of MMH and IRFNA through different injection conditions are presented. First, water and IRFNA are injected to create a spray in the combustion chamber in order to verify effectiveness of test procedures and the test hardware. Next, injection of the hypergolic propellants MMH and IRFNA are done in accordance to the test matrix, although ignition was not observed as expected. These injections are followed by simple drop-on-drop tests to investigate propellant quality and ignition delay. Drop tests are performed with propellants IRFNA/MMH, and again with H2O2/Block 0 as possible propellant replacements for the proposed test plan.

Liquid Oxygen Rotating Friction Ignition Testing of Aluminum and Titanium with Monel and Inconel for Rocket Engine Propulsion System Contamination Investigation

NASA Technical Reports Server (NTRS)

Peralta, S.; Rosales, Keisa R.; Stoltzfus, Joel M.

2009-01-01

Metallic contaminant was found in the liquid oxygen (LOX) pre-valve screen of the shuttle main engine propulsion system on two orbiter vehicles. To investigate the potential for an ignition, NASA Johnson Space Center White Sands Test Facility performed (modified) rotating friction ignition testing in LOX. This testing simulated a contaminant particle in the low-pressure oxygen turbo pump (LPOTP) and the high-pressure oxygen turbo pump (HPOTP) of the shuttle main propulsion system. Monel(R) K-500 and Inconel(R) 718 samples represented the LPOTP and HPOTP materials. Aluminum foil tape and titanium foil represented the contaminant particles. In both the Monel(R) and Inconel(R) material configurations, the aluminum foil tape samples did not ignite after 30 s of rubbing. In contrast, all of the titanium foil samples ignited regardless of the rubbing duration or material configuration. However, the titanium foil ignitions did not propagate to the Monel and Inconel materials.

Optical engine initiation: multiple compartment applications

NASA Astrophysics Data System (ADS)

Hunt, Jeffrey H.

2009-05-01

Modern day propulsion systems are used in aerospace applications for different purposes. The aerospace industry typically requires propulsion systems to operate in a rocket mode in order to drive large boost vehicles. The defense industry generally requires propulsion systems to operate in an air-breathing mode in order to drive missiles. A mixed system could use an air-breathing first stage and a rocket-mode upper stage for space access. Thus, propulsion systems can be used for high mass payloads and where the payload is dominated by the fuel/oxidizer mass being used by the propulsion system. The pulse detonation wave engine (PDWE) uses an alternative type of detonation cycle to achieve the same propulsion results. The primary component of the PDWE is the combustion chamber (or detonation tube). The PDWE represents an attractive propulsion source since its engine cycle is thermodynamically closest to that of a constant volume reaction. This characteristic leads to the inference that a maximum of the potential energy of the PDWE is put into thrust and not into flow work. Consequently, the volume must be increased. The technical community has increasingly adopted the alternative choice of increasing total volume by designing the engine to include a set of banks of smaller combustion chambers. This technique increases the complexity of the ignition subsystem because the inter-chamber timing must be considered. Current approaches to igniting the PDWE have involved separate shock or blast wave initiators and chemical additives designed to enhance detonatibility. An optical ignition subsystem generates a series of optical pulses, where the optical pulses ignite the fuel/oxidizer mixture such that the chambers detonate in a desired order. The detonation system also has an optical transport subsystem for transporting the optical pulses from the optical ignition subsystem to the chambers. The use of optical ignition and transport provides a non-toxic, small, lightweight, precisely controlled detonation system.

Modeling of Transient Flow Mixing of Streams Injected into a Mixing Chamber

NASA Technical Reports Server (NTRS)

Voytovych, Dmytro M.; Merkle, Charles L.; Lucht, Robert P.; Hulka, James R.; Jones, Gregg W.

2006-01-01

Ignition is recognized as one the critical drivers in the reliability of multiple-start rocket engines. Residual combustion products from previous engine operation can condense on valves and related structures thereby creating difficulties for subsequent starting procedures. Alternative ignition methods that require fewer valves can mitigate the valve reliability problem, but require improved understanding of the spatial and temporal propellant distribution in the pre-ignition chamber. Current design tools based mainly on one-dimensional analysis and empirical models cannot predict local details of the injection and ignition processes. The goal of this work is to evaluate the capability of the modern computational fluid dynamics (CFD) tools in predicting the transient flow mixing in pre-ignition environment by comparing the results with the experimental data. This study is a part of a program to improve analytical methods and methodologies to analyze reliability and durability of combustion devices. In the present paper we describe a series of detailed computational simulations of the unsteady mixing events as the cold propellants are first introduced into the chamber as a first step in providing this necessary environmental description. The present computational modeling represents a complement to parallel experimental simulations' and includes comparisons with experimental results from that effort. A large number of rocket engine ignition studies has been previously reported. Here we limit our discussion to the work discussed in Refs. 2, 3 and 4 which is both similar to and different from the present approach. The similarities arise from the fact that both efforts involve detailed experimental/computational simulations of the ignition problem. The differences arise from the underlying philosophy of the two endeavors. The approach in Refs. 2 to 4 is a classical ignition study in which the focus is on the response of a propellant mixture to an ignition source, with emphasis on the level of energy needed for ignition and the ensuing flame propagation issues. Our focus in the present paper is on identifying the unsteady mixing processes that provide the propellant mixture in which the ignition source is to be placed. In particular, we wish to characterize the spatial and temporal mixture distribution with a view toward identifying preferred spatial and temporal locations for the ignition source. As such, the present work is limited to cold flow (pre-ignition) conditions

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

NASA Technical Reports Server (NTRS)

Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.

1993-01-01

A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

Nickel-coated Aluminum Particles: A Promising Fuel for Mars Missions

NASA Technical Reports Server (NTRS)

Shafirovich, Evgeny; Varma, Arvind

2004-01-01

Combustion of metals in carbon dioxide is a promising source of energy for propulsion on Mars. This approach is based on the ability of some metals (e.g. Mg, Al) to burn in CO2 atmosphere and suggests use of the Martian carbon dioxide as an oxidizer in jet or rocket engines. Analysis shows that CO2/metal propulsion will reduce significantly the mass of propellant transported from Earth for long-range mobility on Mars and sample return missions. Recent calculations for the near-term missions indicate that a 200-kg ballistic hopper with CO2/metal rocket engines and a CO2 acquisition unit can perform 10-15 flights on Mars with the total range of 10-15 km, i.e. fulfill the exploration program typically assigned for a rover. Magnesium is currently recognized as a candidate fuel for such engines owing to easy ignition and fast burning in CO2. Aluminum may be more advantageous if a method for reducing its ignition temperature is found. Coating it by nickel is one such method. It is known that a thin nickel layer of nickel on the surface of aluminum particles can prevent their agglomeration and simultaneously facilitate their ignition, thus increasing the efficiency of aluminized propellants.

Design study of RL10 derivatives. Volume 2: Engine design characteristics, appendices. [development of rocket engine for application to space tug propulsion system

NASA Technical Reports Server (NTRS)

1973-01-01

Calculations, curves, and substantiating data which support the engine design characteristics of the RL-10 engines are presented. A description of the RL-10 ignition system is provided. The performance calculations of the RL-10 derivative engines and the performance results obtained are reported. The computer simulations used to establish the control system requirements and to define the engine transient characteristics are included.

Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

NASA Technical Reports Server (NTRS)

Shi, John J.

2005-01-01

At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.

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

NASA Technical Reports Server (NTRS)

Collins, Jacob; Hurlbert, Eric; Romig, Kris; Melcher, John; Hobson, Aaron; Eaton, Phil

2009-01-01

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.

An Ignition Torch Based on Photoignition of Carbon Nanotubes at Elevated Pressure (Briefing Charts)

DTIC Science & Technology

2016-01-04

Ignition Capsule A 10 mg low pressure ignition torch as it ignites a fuel spray We use PITCH to ignite subscale test rockets at 130 K and ~35 atm (~500...distribution is unlimited High Pressure PITCH Applied to a H2/O2 Subscale Rocket Injector Top: a high-pressure chamber for test of subscale rocket injector...to a high-pressure test combustion chamber via a 20 cm extension tube (OD=6 mm) Click >>> 9 DISTRIBUTION STATEMENT A. Approved for public release

Design of a Film Cooling Experiment for Rocket Engines

DTIC Science & Technology

2010-03-01

concentrations inside the UCC (22)............................................................ 25 Figure 7: PIV data in the UCC (23...64 Figure 38: UCC /FCR igniter ............................................................................................. 65 Figure 39: Ethylene...TDLAS Tunable Diode Laser Absorption Spectroscopy UCC Ultra Compact Combustor μm micrometers VI Virtual Instrument Xe Xenon ZnSe

Molded composite pyrogen igniter for rocket motors. [solid propellant ignition

NASA Technical Reports Server (NTRS)

Heier, W. C.; Lucy, M. H. (Inventor)

1978-01-01

A lightweight pyrogen igniter assembly including an elongated molded plastic tube adapted to contain a pyrogen charge was designed for insertion into a rocket motor casing for ignition of the rocket motor charge. A molded plastic closure cap provided for the elongated tube includes an ignition charge for igniting the pyrogen charge and an electrically actuated ignition squib for igniting the ignition charge. The ignition charge is contained within a portion of the closure cap, and it is retained therein by a noncorrosive ignition pellet retainer or screen which is adapted to rest on a shoulder of the elongated tube when the closure cap and tube are assembled together. A circumferentially disposed metal ring is provided along the external circumference of the closure cap and is molded or captured within the plastic cap in the molding process to provide, along with O-ring seals, a leakproof rotary joint.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, soars high after takeoff from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes. Release of the Pegasus XL rocket is scheduled for 8:40 a.m. EST.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer aircraft descends toward the Skid Strip at Cape Canaveral Air Force Station in Florida. The aircraft carried a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, for launch. With the aircraft flying off shore, the Pegasus rocket was released. Five seconds later, the solid propellant engine ignited and boosted the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes. Release of the Pegasus XL rocket occurred at 8:37 a.m. EST.

Enhanced development of a catalyst chamber for the decomposition of up to 1.0 kg/s hydrogen peroxide

NASA Astrophysics Data System (ADS)

Božić, Ognjan; Porrmann, Dennis; Lancelle, Daniel; May, Stefan

2016-06-01

A new innovative hybrid rocket engine concept is developed within the AHRES program of the German Aerospace Center (DLR). This rocket engine based on hydroxyl-terminated polybutadiene (HTPB) with metallic additives as solid fuel and high test peroxide (HTP) as liquid oxidizer. Instead of a conventional ignition system, a catalyst chamber with a silver mesh catalyst is designed to decompose the HTP. The newly modified catalyst chamber is able to decompose up to 1.0 kg/s of 87.5 wt% HTP. Used as a monopropellant thruster, this equals an average thrust of 1600 N. The catalyst chamber is designed using the self-developed software tool SHAKIRA. The applied kinetic law, which determines catalytic decomposition of HTP within the catalyst chamber, is given and commented. Several calculations are carried out to determine the appropriate geometry for complete decomposition with a minimum of catalyst material. A number of tests under steady state conditions are carried out, using 87.5 wt% HTP with different flow rates and a constant amount of catalyst material. To verify the decomposition, the temperature is measured and compared with the theoretical prediction. The experimental results show good agreement with the results generated by the design tool. The developed catalyst chamber provides a simple, reliable ignition system for hybrid rocket propulsion systems based on hydrogen peroxide as oxidizer. This system is capable for multiple reignition. The developed hardware and software can be used to design full scale monopropellant thrusters based on HTP and catalyst chambers for hybrid rocket engines.

Rocket engine injectorhead with flashback barrier

NASA Technical Reports Server (NTRS)

Mungas, Gregory S. (Inventor); Fisher, David J. (Inventor); Mungas, Christopher (Inventor)

2012-01-01

Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation.

State and prospects of solid propellant rocket development

NASA Astrophysics Data System (ADS)

Kukushkin, V. Kh.

1992-07-01

An overview is presented of aspects of solid-propellant rocket engine (SPRE) development with individual treatment given to sustainer and spacecraft SPRE technologies. The paper focuses on low-modulus fuels of composite solid propellant, requirements for adhesion stability, and enhancement of the power characteristics of solid propellants. R&D activities are described that relate to the use of SPREs with extending nozzles and to the design of ultradimensional nozzles for upper-stage engines. Other developments for the SPREs include engines with separate loading and pasty fuel applications, and progress is reported in the direction of detonation SPREs. The SPREs using pasty propellants provide good control over thrust characteristics and fuel qualities. A device is incorporated that assures fuel burning in the combustion region and reliable ignition during restarting of these engines.

The alleged contributions of Pedro E. Paulet to liquid-propellant rocketry

NASA Technical Reports Server (NTRS)

Ordway, F. I., III

1977-01-01

The first practical working liquid propellant rocket motor was claimed by Pedro E. Paulet, a South American engineer from Peru (1895). He operated a conical motor, 10 centimeters in diameter, using nitrogen peroxide and gasoline as propellants and measuring thrust up to 90 kilograms, and apparently used spark ignition and intermittent propellant injection. The test device which he used contained elements of later test stands, such as a spring thrust-measuring device. However, he did not publish his work until twenty-five years later. Evidence is examined concerning this only known claim to liquid propellant rocket engine experiments in the nineteenth century.

Ignition and combustion of aluminum/magnesium alloy particles in O2 at high pressures

NASA Technical Reports Server (NTRS)

Roberts, Ted A.; Burton, Rodney L.; Krier, Herman

1993-01-01

The ignition and combustion of Al, Mg, and Al/Mg alloy particles in 99 percent O2/1 percent N2 mixtures is investigated at high temperatures and pressures for rocket engine applications. The 20-micron particles contain 0, 5, 10, 20, 40, 60, 80, and 100 wt pct Mg alloyed with Al, and are ignited in oxygen using the reflected shock in a single-pulse shock tube near the endwall. Using this technique, the ignition delay and combustion times of the particles are measured at temperatures up to 3250 K as a function of Mg content for oxygen pressures of 8.5, 17, and 34 atm. An ignition model is developed that employs a simple lumped capacitance energy equation and temperature and pressure dependent particle and gas properties. Good agreement is achieved between the measured and predicted trends in the ignition delay times.

Low-temperature Ignition-delay Characteristics of Several Rocket Fuels with Mixed Acid in Modified Open-cup-type Apparatus

NASA Technical Reports Server (NTRS)

Miller, Riley O

1950-01-01

Summaries of low-temperature self-ignition data of various rocket fuels with mixed acid (nitric plus sulfuric) are presented. Several fuels are shown to have shorter ignition-delay intervals and less variation in delay intervals at moderate and sub-zero temperatures than crude N-ethylaniline (monoethylaniline),a rocket fuel in current use.

Development of a new generation solid rocket motor ignition computer code

NASA Technical Reports Server (NTRS)

Foster, Winfred A., Jr.; Jenkins, Rhonald M.; Ciucci, Alessandro; Johnson, Shelby D.

1994-01-01

This report presents the results of experimental and numerical investigations of the flow field in the head-end star grain slots of the Space Shuttle Solid Rocket Motor. This work provided the basis for the development of an improved solid rocket motor ignition transient code which is also described in this report. The correlation between the experimental and numerical results is excellent and provides a firm basis for the development of a fully three-dimensional solid rocket motor ignition transient computer code.

Fuel Chemistry and Combustion Distribution Effects on Rocket Engine Combustion Stability

DTIC Science & Technology

2012-01-25

by Crowe et al. (1963). The small solid rocket motors are fired into the collection tank with the nozzle [Crowe et al. (1963)] and without nozzle...explosions at the end of the droplet lifetime. Upon ignition , a neat droplet of JP-8 will burn orange, and the droplet will regress until all of the...pixel location were estimated by applying a time shift and amplitude scaling factor to the pressure measurements made at the aft end of chamber

The Development of the Wall Momentum Erosive Burning Scaling Law and Macro Scale Erosive Burning Model

DTIC Science & Technology

2010-05-01

burn rate in excess of 2 in/sec at 1000 psi, and Mach numbers that reach 1.0 at the aft end at ignition . Typically, motors with high burning rate...37 VI I. INTRODUCTION Interior ballistics of solid propellant rocket motors continues to be an engineering discipline that is...and one open source paper published between 2005 and 2009 [2, 3, 13]. II. BACKGROUND Erosive burning is a term used in the solid rocket motor

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft takes off from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Nickel-Coated Aluminum Particles: A Promising Fuel for Mars Missions

NASA Technical Reports Server (NTRS)

Shafirovich, Evgeny; Varma, Arvind

2004-01-01

Combustion of metals in carbon dioxide is a promising source of energy for propulsion on Mars. This approach is based on the ability of some metals (e.g. Mg, Al) to burn in CO2 atmosphere and suggests use of the Martian carbon dioxide as an oxidizer in jet or rocket engines. Analysis shows that CO2/metal propulsion will reduce significantly the mass of propellant transported from Earth for long-range mobility on Mars and sample return missions. Recent calculations for the near-term missions indicate that a 200-kg ballistic hopper with CO2/metal rocket engines and a CO2 acquisition unit can perform 10-15 flights on Mars with the total range of 10-15 km, i.e. fulfill the exploration program typically assigned for a rover. Magnesium is currently recognized as a candidate fuel for such engines owing to easy ignition and fast burning in CO2. Aluminum may be more advantageous if a method for reducing its ignition temperature is found. Coating it by nickel is one such method. It is known that a thin nickel layer of nickel on the surface of aluminum particles can prevent their agglomeration and simultaneously facilitate their ignition, thus increasing the efficiency of aluminized propellants. Combustion of single Ni-coated Al particles in different gas environments (O2, CO2, air) was studied using electrodynamic levitation and laser ignition. It was shown that the combustion mechanisms depend on the ambient atmosphere. Combustion in CO2 is characterized by the smaller size and lower brightness of flame than in O2, and by phenomena such as micro-flashes and fragment ejection. The size and brightness of flame gradually decrease as the particle burns.

Review on factors affecting the performance of pulse detonation engine

NASA Astrophysics Data System (ADS)

Tripathi, Saurabh; Pandey, Krishna Murari

2018-04-01

Now a day's rocket engines (air-breathing type) are being used for aerospace purposes but the studies have shown that these are less efficient, so alternatives are being searched for these. Pulse Detonation Engine (PDE) is one such efficient engine which can replace the rocket engines. In this review paper, different researches have been cited. As can be observed from various researches, insertion of obstacles is better. Deflagration to Detonation(DDT) transition process is found to be most important factor. So a lot of researches are being done considering this DDT chamber. Also, the ignition chamber and ejector were found to improve the effectiveness of PDE. The PDE works with a range of Mach 0-4. Flame acceleration is also found to increase the DDT process. Use of valve and valveless engine has also been compared. Various other factors have been focused in this review paper which is found to boost PDE performance.

Premature ignition of a rocket motor.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Moore, Darlene Ruth

During preparation for a rocket sled track (RST) event, there was an unexpected ignition of the zuni rocket motor (10/9/08). Three Sandia staff and a contractor were involved in the accident; the contractor was seriously injured and made full recovery. The data recorder battery energized the low energy initiator in the rocket.

Fluid-solid coupled simulation of the ignition transient of solid rocket motor

NASA Astrophysics Data System (ADS)

Li, Qiang; Liu, Peijin; He, Guoqiang

2015-05-01

The first period of the solid rocket motor operation is the ignition transient, which involves complex processes and, according to chronological sequence, can be divided into several stages, namely, igniter jet injection, propellant heating and ignition, flame spreading, chamber pressurization and solid propellant deformation. The ignition transient should be comprehensively analyzed because it significantly influences the overall performance of the solid rocket motor. A numerical approach is presented in this paper for simulating the fluid-solid interaction problems in the ignition transient of the solid rocket motor. In the proposed procedure, the time-dependent numerical solutions of the governing equations of internal compressible fluid flow are loosely coupled with those of the geometrical nonlinearity problems to determine the propellant mechanical response and deformation. The well-known Zeldovich-Novozhilov model was employed to model propellant ignition and combustion. The fluid-solid coupling interface data interpolation scheme and coupling instance for different computational agents were also reported. Finally, numerical validation was performed, and the proposed approach was applied to the ignition transient of one laboratory-scale solid rocket motor. For the application, the internal ballistics were obtained from the ground hot firing test, and comparisons were made. Results show that the integrated framework allows us to perform coupled simulations of the propellant ignition, strong unsteady internal fluid flow, and propellant mechanical response in SRMs with satisfactory stability and efficiency and presents a reliable and accurate solution to complex multi-physics problems.

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/Hydrogen Rocket Engine

NASA Technical Reports Server (NTRS)

Sutton, R. F.; Lariviere, B. W.

1993-01-01

An advanced LOX/LH2 expander cycle rocket engine, producing 15,000 lbf thrust for Orbital Transfer Vehicle missions, was tested to determine ignition, transition, and main stage characteristics. Detail design and fabrication of the pump fed RS44 integrated component evaluator (ICE) was accomplished using company discretionary resources and was tested under this contracted effort. Successful demonstrations were completed to about the 50 percent fuel turbopump power level (87,000 RPM), but during this last test, a high pressure fuel turbopump (HPFTP) bearing failed curtailing the test program. No other hardware were affected by the HPFTP premature shutdown. The ICE operations matched well with the predicted start transient simulations. The tests demonstrated the feasibility of a high performance advanced expander cycle engine. All engine components operated nominally, except for the HPFTP, during the engine hot-fire tests. A failure investigation was completed using company discretionary resources.

Rocket Launch-Induced Vibration and Ignition Overpressure Response

NASA Technical Reports Server (NTRS)

Caimi, Raoul; Margasahayam, Ravi; Nayfeh, Jamal; Thompson, Karen (Technical Monitor)

2001-01-01

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.

Starting of rocket engine at conditions of simulated altitude using crude monoethylaniline and other fuels with mixed acid

NASA Technical Reports Server (NTRS)

Ladanyi, Dezso J; Sloop, John L; Humphrey, Jack C; Morrell, Gerald

1950-01-01

Experiments were conducted at sea level and pressure altitude of about 55,000 feet at various temperatures to determine starting characteristics of a commercial rocket engine using crude monoethylaniline and other fuels with mixed acid. With crude monoethylaniline, ignition difficulties were encountered at temperatures below about 20 degrees F. With mixed butyl mercaptans, water-white turpentine, and x-pinene, no starting difficulties were experienced at temperatures as low as minus 74 degrees F. Turpentine and x-pinene, however, sometimes left deposits on the injector face. With blends containing furfuryl alcohol and with other blends, difficulties were experienced either from appreciable deposits or from starting.

Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2003-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2002-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

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

NASA Technical Reports Server (NTRS)

McNelis, Nancy B.; Haberbusch, Mark S.

1997-01-01

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.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer touches down at the Skid Strip at Cape Canaveral Air Force Station in Florida. The aircraft carried a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, for launch. With the aircraft flying off shore, the Pegasus rocket was released at 8:37 a.m. EST. Five seconds later, the solid propellant engine ignited and boosted the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft is being readied for takeoff from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Launch Attempt - Prepared for Takeoff - Scrubb

NASA Image and Video Library

2016-12-12

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft is ready for takeoff from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft is ready for takeoff from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Launch Attempt; Scrubbed - Take Off

NASA Image and Video Library

2016-12-12

An Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft takes off from the Skid Strip at Cape Canaveral Air Force Station, Florida. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Ignition transient analysis of solid rocket motor

NASA Technical Reports Server (NTRS)

Han, Samuel S.

1990-01-01

To predict pressure-time and thrust-time behavior of solid rocket motors, a one-dimensional numerical model is developed. The ignition phase of solid rocket motors (time less than 0.4 sec) depends critically on complex interactions among many elements, such as rocket geometry, heat and mass transfer, flow development, and chemical reactions. The present model solves the mass, momentum, and energy equations governing the transfer processes in the rocket chamber as well as the attached converging-diverging nozzle. A qualitative agreement with the SRM test data in terms of head-end pressure gradient and the total thrust build-up is obtained. Numerical results show that the burning rate in the star-segmented head-end section and the erosive burning are two important parameters in the ignition transient of the solid rocket motor (SRM).

Rocket Engine Innovations Advance Clean Energy

NASA Technical Reports Server (NTRS)

2012-01-01

During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

Shuttle Propulsion Overview - The Design Challenges

NASA Technical Reports Server (NTRS)

Owen, James W.

2011-01-01

The major elements of the Space Shuttle Main Propulsion System include two reusable solid rocket motors integrated into recoverable solid rocket boosters, an expendable external fuel and oxidizer tank, and three reusable Space Shuttle Main Engines. Both the solid rocket motors and space shuttle main engines ignite prior to liftoff, with the solid rocket boosters separating about two minutes into flight. The external tank separates, about eight and a half minutes into the flight, after main engine shutdown and is safely expended in the ocean. The SSME's, integrated into the Space Shuttle Orbiter aft structure, are reused after post landing inspections. The configuration is called a stage and a half as all the propulsion elements are active during the boost phase, with only the SSME s continuing operation to achieve orbital velocity. Design and performance challenges were numerous, beginning with development work in the 1970's. The solid rocket motors were large, and this technology had never been used for human space flight. The SSME s were both reusable and very high performance staged combustion cycle engines, also unique to the Space Shuttle. The multi body side mount configuration was unique and posed numerous integration and interface challenges across the elements. Operation of the system was complex and time consuming. This paper describes the design challenges and key areas where the design evolved during the program.

Demonstration of the B4C/NaIO4/PTFE Delay in the U.S. Army Hand-Held Signal

DTIC Science & Technology

2015-05-20

Figure 1. Partial cross section diagram of a hand-held signal showing the rocket motor , delay element, expelling charge, and pyrotechnic payload as...The black powder-based rocket motor , consisting of propellant pellets (G) encased in a cardboard tube, contains an axial core hole to accommodate the...that ignites the rocket motor . Simultaneously, the delay element is ignited and burns for an interval (preferably 5−6 s) before it ignites the black

Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

NASA Technical Reports Server (NTRS)

Melcher, John C., IV; Allred, Jennifer K.

2009-01-01

Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.

Launch of the STS 51-F Challenger

NASA Image and Video Library

1985-07-29

51F-S-157 (29 July 1985) --- Just moments following ignition, the Space Shuttle Challenger, mated to its two solid rocket boosters and an external fuel tank, soars toward a week-long mission in Earth orbit. Note the diamond shock effect in the vicinity of the three main engines. Launch occurred at 5:00 p.m. (EDT), July 29, 1985.

Development of Ionic Liquid Monopropellants for In-Space Propulsion

NASA Technical Reports Server (NTRS)

Blevins, John A.; Osborne, Robin; Drake, Gregory W.

2005-01-01

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 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.

The hard start phenomena in hypergolic engines. Volume 5: RCS engine deformation and destruct tests

NASA Technical Reports Server (NTRS)

Miron, Y.; Perlee, H. E.

1974-01-01

Tests were conducted to determine the causes of Apollo Reaction Control (RCS) engine failures. Stainless steel engines constructed for use in the destructive tests are described. The tests conducted during the three phase investigation are discussed. It was determined that the explosive reaction that destroys the RCS engines occurs at the time of engine ignition and is apparently due to either the detonation of the heterogeneous constituents of the rocket engine, consisting primarily of unreacted propellant droplets and vapors, and/or the detonation of explosive materials accumulated on the engine walls from previous pulses. Photographs of the effects of explosions on the simulated RCS engines are provided.

Deimos Methane-Oxygen Rocket Engine Test Results

NASA Astrophysics Data System (ADS)

Engelen, S.; Souverein, L. J.; Twigt, D. J.

This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

Fundamental rocket injector/spray programs at the Phillips Laboratory

NASA Astrophysics Data System (ADS)

Talley, D. G.

1993-11-01

The performance and stability of liquid rocket engines is determined to a large degree by atomization, mixing, and combustion processes. Control over these processes is exerted through the design of the injector. Injectors in liquid rocket engines are called upon to perform many functions. They must first of all mix the propellants to provide suitable performance in the shortest possible length. For main injectors, this is driven by the tradeoff between the combustion chamber performance, stability, efficiency, and its weight and cost. In gas generators and preburners, however, it is also driven by the possibility of damage to downstream components, for example piping and turbine blades. This can occur if unburned fuel and oxidant later react to create hot spots. Weight and cost considerations require that the injector design be simple and lightweight. For reusable engines, the injectors must also be durable and easily maintained. Suitable atomization and mixing must be produced with as small a pressure drop as possible, so that the size and weight of pressure vessels and turbomachinery can be minimized. However, the pressure drop must not be so small as to promote feed system coupled instabilities. Another important function of the injectors is to ensure that the injector face plate and the chamber and nozzle walls are not damaged. Typically this requires reducing the heat transfer to an acceptable level and also keeping unburned oxygen from chemically attacking the walls, particularly in reusable engines. Therefore the mixing distribution is often tailored to be fuel-rich near the walls. Wall heat transfer can become catastrophically damaging in the presence of acoustic instabilities, so the injector must prevent these from occurring at all costs. In addition to acoustic stability (but coupled with it), injectors must also be kinetically stable. That is, the flame itself must maintain ignition in the combustion chamber. This is not typically a problem with main injectors, but can be a consideration in preburners, where the desire to keep turbine inlet temperatures as cool as possible can make it advantageous for the preburners to operate as far from stoichiometry as can be tolerated.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft gains altitude after takeoff from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft begins its takeoff from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft touches down at the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane provided photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket was released at 8:37 a.m. EST. Five seconds later, the solid propellant engine ignited and boosted the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft takes off from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Launch Attempt - Prepared for Takeoff - Scrubb

NASA Image and Video Library

2016-12-12

A pathfinder aircraft prepares for takeoff from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft soars high after takeoff from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft descends for touchdown at the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane provided photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket was released at 8:37 a.m. EST. Five seconds later, the solid propellant engine ignited and boosted the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Design, Activation, and Operation of the J2-X Subscale Simulator (JSS)

NASA Technical Reports Server (NTRS)

Saunders, Grady P.; Raines, Nickey G.; Varner, Darrel G.

2009-01-01

The purpose of this paper is to give a detailed description of the design, activation, and operation of the J2-X Subscale Simulator (JSS) installed in Cell 1 of the E3 test facility at Stennis Space Center, MS (SSC). The primary purpose of the JSS is to simulate the installation of the J2-X engine in the A3 Subscale Rocket Altitude Test Facility at SSC. The JSS is designed to give aerodynamically and thermodynamically similar plume properties as the J2-X engine currently under development for use as the upper stage engine on the ARES I and ARES V spacecraft. The JSS is a scale pressure fed, LOX/GH fueled rocket that is geometrically similar to the J2-X from the throat to the nozzle exit plane (NEP) and is operated at the same oxidizer to fuel ratios and chamber pressures. This paper describes the heritage hardware used as the basis of the JSS design, the newly designed rocket hardware, igniter systems used, and the activation and operation of the JSS.

Gas-Generator Augmented Expander Cycle Rocket Engine

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2011-01-01

An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

Development of Ionic Liquid Monopropellants for In-Space Propulsion

NASA Technical Reports Server (NTRS)

Blevins, John A.; Drake, Gregory W.; Osborne, Robin J.

2005-01-01

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.

Investigation of Cleanliness Verification Techniques for Rocket Engine Hardware

NASA Technical Reports Server (NTRS)

Fritzemeier, Marilyn L.; Skowronski, Raymund P.

1994-01-01

Oxidizer propellant systems for liquid-fueled rocket engines must meet stringent cleanliness requirements for particulate and nonvolatile residue. These requirements were established to limit residual contaminants which could block small orifices or ignite in the oxidizer system during engine operation. Limiting organic residues in high pressure oxygen systems, such as in the Space Shuttle Main Engine (SSME), is particularly important. The current method of cleanliness verification for the SSME uses an organic solvent flush of the critical hardware surfaces. The solvent is filtered and analyzed for particulate matter followed by gravimetric determination of the nonvolatile residue (NVR) content of the filtered solvent. The organic solvents currently specified for use (1, 1, 1-trichloroethane and CFC-113) are ozone-depleting chemicals slated for elimination by December 1995. A test program is in progress to evaluate alternative methods for cleanliness verification that do not require the use of ozone-depleting chemicals and that minimize or eliminate the use of solvents regulated as hazardous air pollutants or smog precursors. Initial results from the laboratory test program to evaluate aqueous-based methods and organic solvent flush methods for NVR verification are provided and compared with results obtained using the current method. Evaluation of the alternative methods was conducted using a range of contaminants encountered in the manufacture of rocket engine hardware.

Unique thermocouple to measure the temperatures of squibs, igniters, propellants, and rocket nozzles

NASA Astrophysics Data System (ADS)

Nanigian, Jacob; Nanigian, Dan

2006-05-01

The temperatures produced by the various components in the propulsion system of rockets and missiles determine the performance of the rocket. Since these temperatures occur very rapidly and under extreme conditions, standard thermocouples fail before any meaningful temperatures are measured. This paper describes the features of a special family of high performance thermocouples, which can measure these transient temperatures with millisecond response times and under the most severe conditions of erosion. Examples of igniter, propellant and rocket nozzle temperatures are included in this paper. Also included is heat flux measurements made by these sensors in rocket applications.

Repetitive laser ignition by optical breakdown of a LOX/H2 rocket combustion chamber with multi-injector head configuration

NASA Astrophysics Data System (ADS)

Börner, Michael; Manfletti, Chiara; Kroupa, Gerhard; Oschwald, Michael

2017-09-01

This paper reports on the repetitive laser ignition by optical breakdown within an experimental rocket combustion chamber. Ignition was performed by focusing a laser pulse generated by a miniaturized diode-pumped Nd:YAG laser system. The system, which delivers 33.2 mJ in 2.3 ns, was mounted directly to the combustion chamber. The ignition process and flame stabilization was investigated using an optical probe system monitoring the flame attachment across the 15 coaxial injector configuration. 1195 successful ignitions were performed proving the reliability of this laser ignition system and its applicability to the propellant combination LOX/hydrogen at temperatures of T_{{{H}_{ 2} }} = 120-282 K and T_{{{O}_{ 2} }} = 110-281 K.

Pegasus XL CYGNSS Second Launch Attempt

NASA Image and Video Library

2016-12-15

A pathfinder aircraft, at left, prepares for takeoff from the Skid Strip at Cape Canaveral Air Force Station in Florida. The airplane will provide photographic and video imagery of the Orbital ATK L-1011 Stargazer aircraft, in view at right, carrying a Pegasus XL Rocket with eight NASA Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft. With the aircraft flying off shore, the Pegasus rocket will be released. Five seconds later, the solid propellant engine will ignite and boost the eight hurricane observatories to orbit. The eight CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Ignition and combustion of lunar propellants

NASA Technical Reports Server (NTRS)

Burton, Rodney L.; Roberts, Ted A.; Krier, Herman

1993-01-01

The ignition and combustion of Al, Mg, and Al/Mg alloy particles in 99 percent O2/1 percent N2 mixtures is investigated at high temperatures and pressures for rocket engine applications. The 20 micron particles contain 0, 5, 10, 20, 40, 60, 80, and 100 weight percent Mg alloyed with Al, and are ignited in oxygen using the reflected shock in a shock tube near the endwall. Using this technique, the ignition delay and combustion times of the particles are measured at temperatures up to 3250 K as a function of Mg content for oxygen pressures of 8.5, 17, and 34 atm. An ignition model is developed which employs a simple lumped capacitance energy equation and temperature and pressure dependent particle and gas properties. Good agreement is achieved between the measured and predicted trends in the ignition delay times. For the particles investigated, the contribution of heterogeneous reaction to the heating of the particle is found to be significant at lower temperatures, but may be neglected as gas temperatures above 3000 K. As little as 10 percent Mg reduces the ignition delay time substantially at all pressures tested. The particle ignition delay times decrease with increasing Mg content, and this reduction becomes less pronounced as oxidizer temperature and pressure are increased.

Status on Technology Development of Optic Fiber-Coupled Laser Ignition System for Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew; Bossard, John

2003-01-01

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concept: not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio. This incentive can be translated to a convenience in the thrust chamber packaging.

Fuel/oxidizer-rich high-pressure preburners. [staged-combustion rocket engine

NASA Technical Reports Server (NTRS)

Schoenman, L.

1981-01-01

The analyses, designs, fabrication, and cold-flow acceptance testing of LOX/RP-1 preburner components required for a high-pressure staged-combustion rocket engine are discussed. Separate designs of injectors, combustion chambers, turbine simulators, and hot-gas mixing devices are provided for fuel-rich and oxidizer-rich operation. The fuel-rich design addresses the problem of non-equilibrium LOX/RP-1 combustion. The development and use of a pseudo-kinetic combustion model for predicting operating efficiency, physical properties of the combustion products, and the potential for generating solid carbon is presented. The oxygen-rich design addresses the design criteria for the prevention of metal ignition. This is accomplished by the selection of materials and the generation of well-mixed gases. The combining of unique propellant injector element designs with secondary mixing devices is predicted to be the best approach.

Hydrogen-oxygen auxiliary propulsion for the space shuttle. Volume 1: High pressure thrusters

NASA Technical Reports Server (NTRS)

1973-01-01

Technology for long life, high performing, gaseous hydrogen-gaseous oxygen rocket engines suitable for auxiliary propulsion was provided by a combined analytical and experimental program. Propellant injectors, fast response valves, igniters, and regeneratively and film-cooled thrust chambers were tested over a wide range of operating conditions. Data generated include performance, combustion efficiency, thermal characteristics film cooling effectiveness, dynamic response in pulsing, and cycle life limitations.

Logistics Complete Round Charts: Grenades, Mines, Pyrotechnics, Rockets, Rocket Motors, Demolition Material

DTIC Science & Technology

1986-06-01

Flash 3. (342 g) Electrically Comp Actuated By Blasting Machine or Battery 5/ 8 STD 37524 Simulator, Projectile M1lSA2 0.30 7549246 MIL-S-10058 Hand...8799710 Blastin Comp (0.07 oz) Blasting (Whistle) Fuze Fuze 8-14 Frictio Friction (Burst) ’ l3AI M3A1 721 Igniter 8833721 Igniter last 5-10 8799714...8799715 iBlasti Blasting Sec FuzeN Fuze M3A1 •721 Igniter 8833721 Igniter Inst 6-10 8799714 879971 Blast Blasting See Fuze Puze N13A1 587 FrictLi 8848587

Dual-Laser-Pulse Ignition

NASA Technical Reports Server (NTRS)

Trinh, Huu; Early, James W.; Thomas, Matthew E.; Bossard, John A.

2006-01-01

A dual-pulse laser (DPL) technique has been demonstrated for generating laser-induced sparks (LIS) to ignite fuels. The technique was originally intended to be applied to the ignition of rocket propellants, but may also be applicable to ignition in terrestrial settings in which electric igniters may not be suitable.

The technique for Simulation of Transient Combustion Processes in the Rocket Engine Operating with Gaseous Fuel “Hydrogen and Oxygen”

NASA Astrophysics Data System (ADS)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2017-01-01

The article describes the method for simulation of transient combustion processes in the rocket engine. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. Reactions mechanisms have been taken from several sources and verified. The method for converting ozone properties from the Shomate equation to the NASA-polynomial format was described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. Modeling difficulties with combustion model Finite Rate Chemistry, associated with a large scatter of reference data were identified and described. The way to generate the Flamelet library with CFX-RIF is described. Formulated adequate reaction mechanisms verified at a steady state have also been tested for transient simulation. The Flamelet combustion model was recognized as adequate for the transient mode. Integral parameters variation relates to the values obtained during stationary simulation. A cyclic irregularity of the temperature field, caused by precession of the vortex core, was detected in the chamber with the proposed simulation technique. Investigations of unsteady processes of rocket engines including the processes of ignition were proposed as the area for application of the described simulation technique.

Characterization of the space shuttle reaction control system engine

NASA Technical Reports Server (NTRS)

Wilson, M. S.; Stechman, R. C.; Edelman, R. B.; Fortune, O. F.; Economos, C.

1972-01-01

A computer program was developed and written in FORTRAN 5 which predicts the transient and steady state performance and heat transfer characteristics of a pulsing GO2/GH2 rocket engine. This program predicts the dynamic flow and ignition characteristics which, when combined in a quasi-steady state manner with the combustion and mixing analysis program, will provide the thrust and specific impulse of the engine as a function of time. The program also predicts the transient and steady state heat transfer characteristics of the engine using various cooling concepts. The computer program, test case, and documentation are presented. The program is applicable to any system capable of utilizing the FORTRAN 4 or FORTRAN 5 language.

Status of the Combustion Devices Injector Technology Program at the NASA MSFC

NASA Technical Reports Server (NTRS)

Jones, Gregg; Protz, Christopher; Trinh, Huu; Tucker, Kevin; Nesman, Tomas; Hulka, James

2005-01-01

To support the NASA Space Exploration Mission, an in-house program called Combustion Devices Injector Technology (CDIT) is being conducted at the NASA Marshall Space Flight Center (MSFC) for the fiscal year 2005. CDIT is focused on developing combustor technology and analysis tools to improve reliability and durability of upper-stage and in-space liquid propellant rocket engines. The three areas of focus include injector/chamber thermal compatibility, ignition, and combustion stability. In the compatibility and ignition areas, small-scale single- and multi-element hardware experiments will be conducted to demonstrate advanced technological concepts as well as to provide experimental data for validation of computational analysis tools. In addition, advanced analysis tools will be developed to eventually include 3-dimensional and multi- element effects and improve capability and validity to analyze heat transfer and ignition in large, multi-element injectors.

MEMS-Based Solid Propellant Rocket Array Thruster

NASA Astrophysics Data System (ADS)

Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

Catalytic Microtube Rocket Igniter

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Deans, Matthew C.

2011-01-01

Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Effect of Stagger on the Vibroacoustic Loads from Clustered Rockets

NASA Technical Reports Server (NTRS)

Rojo, Raymundo; Tinney, Charles E.; Ruf, Joseph H.

2016-01-01

The effect of stagger startup on the vibro-acoustic loads that form during the end- effects-regime of clustered rockets is studied using both full-scale (hot-gas) and laboratory scale (cold gas) data. Both configurations comprise three nozzles with thrust optimized parabolic contours that undergo free shock separated flow and restricted shock separated flow as well as an end-effects regime prior to flowing full. Acoustic pressure waveforms recorded at the base of the nozzle clusters are analyzed using various statistical metrics as well as time-frequency analysis. The findings reveal a significant reduction in end- effects-regime loads when engine ignition is staggered. However, regardless of stagger, both the skewness and kurtosis of the acoustic pressure time derivative elevate to the same levels during the end-effects-regime event thereby demonstrating the intermittence and impulsiveness of the acoustic waveforms that form during engine startup.

Ignition of the Pegasus rocket moments after release from the B-52 signaled acceleration of the X-43

NASA Technical Reports Server (NTRS)

2001-01-01

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A 'stack' lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing 'scramjet' engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz. The X-43A flights are the first actual flight tests of an aircraft powered by a scramjet engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). Some 90 minutes after takeoff, the Pegasus will launch from a B-52, rocketing the X-43A to Mach 7 at 95,000 feet altitude, or Mach 10 at 105,000 feet altitude. The X-43A will be powered by its revolutionary air-breathing supersonic-combustion ramjet or 'scramjet' engine. The X-43A will then fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments as it descends until it splashes into the Pacific Ocean.

Propulsion/ASME Rocket-Based Combined Cycle Activities in the Advanced Space Transportation Program Office

NASA Technical Reports Server (NTRS)

Hueter, Uwe; Turner, James

1998-01-01

NASA's Office Of Aeronautics and Space Transportation Technology (OASTT) has establish three major coals. "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville,Ala. focuses on future space transportation technologies under the "Access to Space" pillar. The Advanced Reusable Technologies (ART) Project, part of ASTP, focuses on the reusable technologies beyond those being pursued by X-33. The main activity over the past two and a half years has been on advancing the rocket-based combined cycle (RBCC) technologies. In June of last year, activities for reusable launch vehicle (RLV) airframe and propulsion technologies were initiated. These activities focus primarily on those technologies that support the year 2000 decision to determine the path this country will take for Space Shuttle and RLV. In February of this year, additional technology efforts in the reusable technologies were awarded. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion. Aerojet, Boeing-Rocketdyne and Pratt & Whitney were selected for a two-year period to design, build and ground test their RBCC engine concepts. In addition, ASTROX, Pennsylvania State University (PSU) and University of Alabama in Huntsville also conducted supporting activities. The activity included ground testing of components (e.g., injectors, thrusters, ejectors and inlets) and integrated flowpaths. An area that has caused a large amount of difficulty in the testing efforts is the means of initiating the rocket combustion process. All three of the prime contractors above were using silane (SiH4) for ignition of the thrusters. This follows from the successful use of silane in the NASP program for scramjet ignition. However, difficulties were immediately encountered when silane (an 80/20 mixture of hydrogen/silane) was used for rocket ignition.

Analysis of advanced solid rocket motor ignition phenomena

NASA Technical Reports Server (NTRS)

Foster, Winfred A., Jr.; Jenkins, Rhonald M.

1995-01-01

This report presents the results obtained from an experimental analysis of the flow field in the slots of the star grain section in the head-end of the advanced solid rocket motor during the ignition transient. This work represents an extension of the previous tests and analysis to include the effects of using a center port in conjunction with multiple canted igniter ports. The flow field measurements include oil smear data on the star slot walls, pressure and heat transfer coefficient measurements on the star slot walls and velocity measurements in the star slot.

Theoretical and Experimental Investigations of Ignition, Combustion and Expansion Processes of Hypergolic Liquid Fuel Combinations at Gas Temperatures up to 3000 K. Thesis - Rhein-Westfalia Technical Coll., 1967

NASA Technical Reports Server (NTRS)

Schulz, Harry

1987-01-01

The ignition, combustion, and expansion characteristics of hypergolic liquid propellant mixtures in small rocket engines are studied theoretically and experimentally. It is shown by using the Bray approximation procedure that the reaction H + OH + M = H2O + M (where M is the molecular mass of the gas mixture) has a strong effect on the combustion efficiency. Increases in recombination energies ranging from 30 to 65% were obtained when the rate of this reaction was increased by a factor of 10 in gas mixtures containing 90% oxygen. The effect of aluminum additions and various injection techniques on the combustion process is investigated.

Advanced research and technology program for advanced high pressure oxygen-hydrogen rocket propulsion

NASA Technical Reports Server (NTRS)

Marsik, S. J.; Morea, S. F.

1985-01-01

A research and technology program for advanced high pressure, oxygen-hydrogen rocket propulsion technology is presently being pursued by the National Aeronautics and Space Administration (NASA) to establish the basic discipline technologies, develop the analytical tools, and establish the data base necessary for an orderly evolution of the staged combustion reusable rocket engine. The need for the program is based on the premise that the USA will depend on the Shuttle and its derivative versions as its principal Earth-to-orbit transportation system for the next 20 to 30 yr. The program is focused in three principal areas of enhancement: (1) life extension, (2) performance, and (3) operations and diagnosis. Within the technological disciplines the efforts include: rotordynamics, structural dynamics, fluid and gas dynamics, materials fatigue/fracture/life, turbomachinery fluid mechanics, ignition/combustion processes, manufacturing/producibility/nondestructive evaluation methods and materials development/evaluation. An overview of the Advanced High Pressure Oxygen-Hydrogen Rocket Propulsion Technology Program Structure and Working Groups objectives are presented with highlights of several significant achievements.

Advanced research and technology programs for advanced high-pressure oxygen-hydrogen rocket propulsion

NASA Technical Reports Server (NTRS)

Marsik, S. J.; Morea, S. F.

1985-01-01

A research and technology program for advanced high pressure, oxygen-hydrogen rocket propulsion technology is presently being pursued by the National Aeronautics and Space Administration (NASA) to establish the basic discipline technologies, develop the analytical tools, and establish the data base necessary for an orderly evolution of the staged combustion reusable rocket engine. The need for the program is based on the premise that the USA will depend on the Shuttle and its derivative versions as its principal Earth-to-orbit transportation system for the next 20 to 30 yr. The program is focused in three principal areas of enhancement: (1) life extension, (2) performance, and (3) operations and diagnosis. Within the technological disciplines the efforts include: rotordynamics, structural dynamics, fluid and gas dynamics, materials fatigue/fracture/life, turbomachinery fluid mechanics, ignition/combustion processes, manufacturing/producibility/nondestructive evaluation methods and materials development/evaluation. An overview of the Advanced High Pressure Oxygen-Hydrogen Rocket Propulsion Technology Program Structure and Working Groups objectives are presented with highlights of several significant achievements.

Advanced research and technology programs for advanced high-pressure oxygen-hydrogen rocket propulsion

NASA Astrophysics Data System (ADS)

Marsik, S. J.; Morea, S. F.

1985-03-01

A research and technology program for advanced high pressure, oxygen-hydrogen rocket propulsion technology is presently being pursued by the National Aeronautics and Space Administration (NASA) to establish the basic discipline technologies, develop the analytical tools, and establish the data base necessary for an orderly evolution of the staged combustion reusable rocket engine. The need for the program is based on the premise that the USA will depend on the Shuttle and its derivative versions as its principal Earth-to-orbit transportation system for the next 20 to 30 yr. The program is focused in three principal areas of enhancement: (1) life extension, (2) performance, and (3) operations and diagnosis. Within the technological disciplines the efforts include: rotordynamics, structural dynamics, fluid and gas dynamics, materials fatigue/fracture/life, turbomachinery fluid mechanics, ignition/combustion processes, manufacturing/producibility/nondestructive evaluation methods and materials development/evaluation. An overview of the Advanced High Pressure Oxygen-Hydrogen Rocket Propulsion Technology Program Structure and Working Groups objectives are presented with highlights of several significant achievements.

Hybrid propulsion technology program: Phase 1, volume 4

NASA Technical Reports Server (NTRS)

Claflin, S. E.; Beckman, A. W.

1989-01-01

The use of a liquid oxidizer-solid fuel hybrid propellant combination in booster rocket motors appears extremely attractive due to the integration of the best features of liquid and solid propulsion systems. The hybrid rocket combines the high performance, clean exhaust, and safety of liquid propellant engines with the low cost and simplicity of solid propellant motors. Additionally, the hybrid rocket has unique advantages such as an inert fuel grain and a relative insensitivity to fuel grain and oxidizer injection anomalies. The advantages mark the hybrid rocket as a potential replacement or alternative for current and future solid propellant booster systems. The issues are addressed and recommendations are made concerning oxidizer feed systems, injectors, and ignition systems as related to hybrid rocket propulsion. Early in the program a baseline hybrid configuration was established in which liquid oxygen would be injected through ports in a solid fuel whose composition is based on hydroxyl terminated polybutadiene (HTPB). Liquid oxygen remained the recommended oxidizer and thus all of the injector concepts which were evaluated assumed only liquid would be used as the oxidizer.

Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

NASA Technical Reports Server (NTRS)

Melcher, John C., IV; Allred, Jennifer K.

2009-01-01

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to 122,000 ft (37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

Design and testing of a caseless solid-fuel integral-rocket ramjet engine for use in small tactical missiles

NASA Astrophysics Data System (ADS)

Fruge, Keith J.

1991-09-01

An investigation was conducted to determine the feasibility of a low cost, caseless, solid fuel integral rocket ramjet (IRSFRJ) that has no ejecta. Analytical design of a ramjet powered air-to-ground missile capable of being fired from a remotely piloted vehicle or helicopter was accomplished using current JANNAF and Air Force computer codes. The results showed that an IRSFRJ powered missile can exceed the velocity and range of current systems by more than a two to one ratio, without an increase in missile length and weight. A caseless IRSFRJ with a nonejecting port cover was designed and tested. The experimental results of the static tests showed that a low cost, caseless IRSFRJ with a nonejectable port cover is a viable design. Rocket ramjet transition was demonstrated and ramjet ignition was found to be insensitive to the booster tail off to air injection timing sequence.

Direct electrical arc ignition of hybrid rocket motors

NASA Astrophysics Data System (ADS)

Judson, Michael I., Jr.

Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

Mars Climate Orbiter

NASA Technical Reports Server (NTRS)

1998-01-01

The purpose of this mission is to study the climate history and the water distribution of Mars. Beautiful panoramic views of the shuttle on the launch pad, engine ignition, Rocket launch, and the separation and burnout of the Solid Rocket Boosters are shown. The footage also includes an animation of the mission. Detailed views of the path that the Orbiter traversed were shown. Once the Orbiter lands on the surface of Mars, it will dig a six to eight inch hole and collect samples from the planets' surface. The animation also included the prospective return of the Orbiter to Earth over the desert of Utah. The remote sensor on the Orbiter helps in finding the exact location of the Orbiter so that scientists may collect the sample and analyze it.

ASRM Multi-Port Igniter Flow Field Analysis

NASA Technical Reports Server (NTRS)

Kania, Lee; Dumas, Catherine; Doran, Denise

1993-01-01

The Advanced Solid Rocket Motor (ASRM) program was initiated by NASA in response to the need for a new generation rocket motor capable of providing increased thrust levels over the existing Redesigned Solid Rocket Motor (RSRM) and thus augment the lifting capacity of the space shuttle orbiter. To achieve these higher thrust levels and improve motor reliability, advanced motor design concepts were employed. In the head end of the motor, for instance, the propellent cast has been changed from the conventional annular configuration to a 'multi-slot' configuration in order to increase the burn surface area and guarantee rapid motor ignition. In addition, the igniter itself has been redesigned and currently features 12 exhaust ports in order to channel hot igniter combustion gases into the circumferential propellent slots. Due to the close proximity of the igniter ports to the propellent surfaces, new concerns over possible propellent deformation and erosive burning have arisen. The following documents the effort undertaken using computational fluid dynamics to perform a flow field analysis in the top end of the ASRM motor to determine flow field properties necessary to permit a subsequent propellent fin deformation analysis due to pressure loading and an assessment of the extent of erosive burning.

ASRM multi-port igniter flow field analysis

NASA Astrophysics Data System (ADS)

Kania, Lee; Dumas, Catherine; Doran, Denise

1993-07-01

The Advanced Solid Rocket Motor (ASRM) program was initiated by NASA in response to the need for a new generation rocket motor capable of providing increased thrust levels over the existing Redesigned Solid Rocket Motor (RSRM) and thus augment the lifting capacity of the space shuttle orbiter. To achieve these higher thrust levels and improve motor reliability, advanced motor design concepts were employed. In the head end of the motor, for instance, the propellent cast has been changed from the conventional annular configuration to a 'multi-slot' configuration in order to increase the burn surface area and guarantee rapid motor ignition. In addition, the igniter itself has been redesigned and currently features 12 exhaust ports in order to channel hot igniter combustion gases into the circumferential propellent slots. Due to the close proximity of the igniter ports to the propellent surfaces, new concerns over possible propellent deformation and erosive burning have arisen. The following documents the effort undertaken using computational fluid dynamics to perform a flow field analysis in the top end of the ASRM motor to determine flow field properties necessary to permit a subsequent propellent fin deformation analysis due to pressure loading and an assessment of the extent of erosive burning.

Ignition of the Pegasus rocket moments after release from the B-52 signaled acceleration of the X-43A/Pegasus combination over the Pacific Ocean

NASA Image and Video Library

2001-06-02

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A "stack" lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing "scramjet" engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz.

Moments after release from NASA's B-52 carrier aircraft, the X-43A/Pegasus "stack" is seen before ignition of the Pegasus rocket motor on

NASA Image and Video Library

2001-06-02

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A "stack" lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing "scramjet" engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz.

Marshall Space Flight Center Autumn 2005

NASA Technical Reports Server (NTRS)

Allen, Mike; Clar, Harry E.

2006-01-01

The East Test Area at Marshall Space Flight Center has five major test stands, each of which has two or more test positions, not counting the SSME and RD-180 engine test facilities in the West Test Area. These research and development facilities are capable of testing high pressure pumps, both fuel and oxidizer, injectors, chambers and sea-level engine assemblies, as well as simulating deep space environments in the 12, 15 and 20 foot vacuum chambers. Liquid propellant capabilities are high pressure hydrogen (liquid and gas), methane (liquid and gas), and RP-1 and high pressure LOX. Solid propellant capability includes thrust measurement and firing capability up to 1/6 scale Shuttle SRB segment. In the past six months MSFC supported multiple space access and exploration programs in the previous six months. Major programs were Space Exploration, Shuttle External Tank research, Reusable Solid Rocket Motor (RSRM) development, as well as research programs for NASA and other customers. At Test Stand 115 monopropellant ignition testing was conducted on one position. At the second position multiple ignition/variable burn time cycles were conducted on Vacuum Plasma Spatter (VPS) coated injectors. Each injector received fifty cycles; the propellants were LOX Hydrogen and the ignition source was TEA. Following completion of the monopropellant test series the stand was reconfigured to support ignition testing on a LOX Methane injector system. At TS 116 a thrust stand used to test Booster Separation Motors from the Shuttle SRB system was disassembled and moved from Chemical Systems Division s Coyote Canyon plant to MSFC. The stand was reassembled and readied for BSM testing. Also, a series of tests was run on a Pratt & Whitney Rocketdyne Low Element Density (LED) injector engine. The propellants for this engine are LOX and LH2. At TS 300 the 20 foot vacuum chamber was configured to support hydrogen testing in the Multipurpose Hydrogen Test Bed (MHTB) test article. This testing, which went 24/7 for fourteen consecutive days, demonstrated long duration storage methods intended to minimize losses of propellant in support of the Space Exploration Initiative. The facility is being converted to support similar research using liquid methane. The 12 foot chamber at TS 300 was used to create ascent profiles (both heat and altitude effects) for foam panel testing in support of the Shuttle External Tank program. At TS 500, one position was in build-up to support ATK Thiokol research into the gas dynamics associated with high pressure flow across the propellant joint in segmented solid rocket motors. The testing involves flowing high pressure gas through a 24 motor case. Initial tests will be conducted with simulated aluminum grain, followed by tests using actual propellant. The second position at TS 500 has been in build-up for testing a LOX methane thruster manufactured by KT Engineering. At the Solid Propulsion Test Area (SPTA), the first dual segment 24 solid rocket motor was fired for ATK Thiokol in support of the RSRM program. A new axial thrust measurement stand was designed and fabricated for this testing. Real Time Radiography (RTR) will be deployed to examine nozzle erosion on the next dual segment motor.

Propellant-Flow-Actuated Rocket Engine Igniter

NASA Technical Reports Server (NTRS)

Wollen, Mark

2013-01-01

A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0.001 lb/s) and pressure drops on the order of 30 to 50 kilopascal (a few psi). An analytical model has been created and tested in conjunction with a precisely calibrated reference model. The analytical model accurately captures the overall behavior of this innovation. The model is a simple "volume-orifice" concept, with each chamber considered a single temperature and pressure "node" connected to adjacent nodes, or to vent paths through flow control orifices. Mass and energy balances are applied to each node, with gas flow predicted using simple compressible flow equations.

Ignition and combustion characteristics of metallized propellants, phase 2

NASA Technical Reports Server (NTRS)

Mueller, D. C.; Turns, S. R.

1994-01-01

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.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This cutaway illustration shows the Saturn V S-IVB (third) stage with the callouts of its major components. When the S-II (second) stage of the powerful Saturn V rocket burnt out and was separated the remaining units approached orbit around the Earth. Injection into the desired orbit was attaineded as the S-IVB (third stage) was ignited and burnt. The S-IVB stage was powered by a single 200,000-pound thrust J-2 engine and had a re-start capability built in for its J-2 engine. The S-IVB restarted to speed the Apollo spacecraft to escape velocity injecting it and the astronauts into a moon trajectory.

Preventing Accidental Ignition of Upper-Stage Rocket Motors

NASA Technical Reports Server (NTRS)

Hickman, John; Morgan, Herbert; Cooper, Michael; Murbach, Marcus

2005-01-01

A report presents a proposal to reduce the risk of accidental ignition of certain upper-stage rocket motors or other high energy hazardous systems. At present, mechanically in-line initiators are used for initiation of many rocket motors and/or other high-energy hazardous systems. Electrical shorts and/or mechanical barriers, which are the basic safety devices in such systems, are typically removed as part of final arming or pad preparations while personnel are present. At this time, static discharge, test equipment malfunction, or incorrect arming techniques can cause premature firing. The proposal calls for a modular out-of-line ignition system incorporating detonating-cord elements, identified as the donor and the acceptor, separated by an air gap. In the safe configuration, the gap would be sealed with two shields, which would prevent an accidental firing of the donor from igniting the system. The shields would be removed to enable normal firing, in which shrapnel generated by the donor would reliably ignite the acceptor to continue the ordnance train. The acceptor would then ignite a through bulkhead initiator (or other similar device), which would ignite the motor or high-energy system. One shield would be remotely operated and would be moved to the armed position when a launch was imminent or conversely returned to the safe position if the launch were postponed. In the event of failure of the remotely operated shield, the other shield could be inserted manually to safe the system.

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

1997-01-01

A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

Catalytic decomposition of nitrous oxide monopropellant for hybrid motor ignition

NASA Astrophysics Data System (ADS)

Wilson, Matthew

Nitrous oxide (N2O) is an inexpensive and readily available non-toxic rocket motor oxidizer. It is the most commonly used oxidizer for hybrid bipropellant rocket systems, and several bipropellant liquid rocket designs have also used nitrous oxide. In liquid form, N2O is highly stable, but in vapor form it has the potential to decompose exothermically, releasing up to 1865 Joules per gram of vapor as it dissociates into nitrogen and oxygen. Consequently, it has long been considered as a potential "green" replacement for existing highly toxic and dangerous monopropellants. This project investigates the feasibility of using the nitrous oxide decomposition reaction as a monopropellant energy source for igniting liquid bipropellant and hybrid rockets that already use nitrous oxide as the primary oxidizer. Because nitrous oxide is such a stable propellant, the energy barrier to dissociation is quite high; normal thermal decomposition of the vapor phase does not occur until temperatures are above 800 C. The use of a ruthenium catalyst decreases the activation energy for this reaction to allow rapid decomposition below 400 C. This research investigates the design for a prototype device that channels the energy of dissociation to ignite a laboratory scale hybrid rocket motor.

Tripropellant combustion process

NASA Technical Reports Server (NTRS)

Kmiec, T. D.; Carroll, R. G.

1988-01-01

The addition of small amounts of hydrogen to the combustion of LOX/hydrocarbon propellants in large rocket booster engines has the potential to enhance the system stability. Programs being conducted to evaluate the effects of hydrogen on the combustion of LOX/hydrocarbon propellants at supercritical pressures are described. Combustion instability has been a problem during the development of large hydrocarbon fueled rocket engines. At the higher combustion chamber pressures expected for the next generation of booster engines, the effect of unstable combustion could be even more destructive. The tripropellant engine cycle takes advantage of the superior cooling characteristics of hydrogen to cool the combustion chamber and a small amount of the hydrogen coolant can be used in the combustion process to enhance the system stability. Three aspects of work that will be accomplished to evaluate tripropellant combustion are described. The first is laboratory demonstration of the benefits through the evaluation of drop size, ignition delay and burning rate. The second is analytical modeling of the combustion process using the empirical relationship determined in the laboratory. The third is a subscale demonstration in which the system stability will be evaluated. The approach for each aspect is described and the analytical models that will be used are presented.

Saving Lives With Rocket Power

NASA Technical Reports Server (NTRS)

2000-01-01

Thiokol Propulsion uses NASA's surplus rocket fuel to produce a flare that can safely destroy land mines. Through a Memorandum of Agreement between Thiokol and Marshall Space Flight Center, Thiokol uses the scrap Reusable Solid Rocket Motor (RSRM) propellant. The resulting Demining Device was developed by Thiokol with the help of DE Technologies. The Demining Device neutralizes land mines in the field without setting them off. The Demining Device flare is placed next to an uncovered land mine. Using a battery-triggered electric match, the flare is then ignited. Using the excess and now solidified rocket fuel, the flare burns a hole in the mine's case and ignites the explosive contents. Once the explosive material is burned away, the mine is disarmed and no longer dangerous.

State Machine Modeling of the Space Launch System Solid Rocket Boosters

NASA Technical Reports Server (NTRS)

Harris, Joshua A.; Patterson-Hine, Ann

2013-01-01

The Space Launch System is a Shuttle-derived heavy-lift vehicle currently in development to serve as NASA's premiere launch vehicle for space exploration. The Space Launch System is a multistage rocket with two Solid Rocket Boosters and multiple payloads, including the Multi-Purpose Crew Vehicle. Planned Space Launch System destinations include near-Earth asteroids, the Moon, Mars, and Lagrange points. The Space Launch System is a complex system with many subsystems, requiring considerable systems engineering and integration. To this end, state machine analysis offers a method to support engineering and operational e orts, identify and avert undesirable or potentially hazardous system states, and evaluate system requirements. Finite State Machines model a system as a finite number of states, with transitions between states controlled by state-based and event-based logic. State machines are a useful tool for understanding complex system behaviors and evaluating "what-if" scenarios. This work contributes to a state machine model of the Space Launch System developed at NASA Ames Research Center. The Space Launch System Solid Rocket Booster avionics and ignition subsystems are modeled using MATLAB/Stateflow software. This model is integrated into a larger model of Space Launch System avionics used for verification and validation of Space Launch System operating procedures and design requirements. This includes testing both nominal and o -nominal system states and command sequences.

Measuring the Internal Environment of Solid Rocket Motors During Ignition

NASA Technical Reports Server (NTRS)

Weisenberg, Brent; Smith, Doug; Speas, Kyle; Corliss, Adam

2003-01-01

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.

Dynamic analysis of solid propellant grains subjected to ignition pressurization loading

NASA Astrophysics Data System (ADS)

Chyuan, Shiang-Woei

2003-11-01

Traditionally, the transient analysis of solid propellant grains subjected to ignition pressurization loading was not considered, and quasi-elastic-static analysis was widely adopted for structural integrity because the analytical task gets simplified. But it does not mean that the dynamic effect is not useful and could be neglected arbitrarily, and this effect usually plays a very important role for some critical design. In order to simulate the dynamic response for solid rocket motor, a transient finite element model, accompanied by concepts of time-temperature shift principle, reduced integration and thermorheologically simple material assumption, was used. For studying the dynamic response, diverse ignition pressurization loading cases were used and investigated in the present paper. Results show that the dynamic effect is important for structural integrity of solid propellant grains under ignition pressurization loading. Comparing the effective stress of transient analysis and of quasi-elastic-static analysis, one can see that there is an obvious difference between them because of the dynamic effect. From the work of quasi-elastic-static and transient analyses, the dynamic analysis highlighted several areas of interest and a more accurate and reasonable result could be obtained for the engineer.

A Preliminary Investigation on the Destruction of Solid-Propellant Rocket Motors by Impact from Small Particles

NASA Technical Reports Server (NTRS)

Carter, David J., Jr.

1960-01-01

An investigation was conducted to determine whether solid-propellant rocket motors could be ignited and destroyed by small-particle impacts at particle velocities up to a approximately 10,940 feet per second. Spheres ranging from 1/16 to 7/32 inch in diameter were fired into simulated rocket motors containing T-22 propellant over a range of ambient pressures from sea level to 0.12 inch of mercury absolute. Simulated cases of stainless steel, aluminum alloy, and laminated Fiberglas varied in thickness from 1/50 to 1/8 inch. Within the scope of this investigation, it was found that ignition and explosive destruction of simulated steel-case rocket motors could result from impacts by steel spheres at the lowest attainable pressure.

14 CFR 23.1165 - Engine ignition systems.

Code of Federal Regulations, 2012 CFR

2012-01-01

... Controls and Accessories § 23.1165 Engine ignition systems. Link to an amendment published at 76 FR 75759... discharge of any battery used for engine ignition. (e) Each turbine engine ignition system must be... ignition systems. (f) In addition, for commuter category airplanes, each turbine engine ignition system...

Preliminary 2-D shell analysis of the space shuttle solid rocket boosters

NASA Technical Reports Server (NTRS)

Knight, Norman F., Jr.; Gillian, Ronnie E.; Nemeth, Michael P.

1987-01-01

A two-dimensional shell model of an entire solid rocket booster (SRB) has been developed using the STAGSC-1 computer code and executed on the Ames CRAY computer. The purpose of these analyses is to calculate the overall deflection and stress distributions for the SRB when subjected to mechanical loads corresponding to critical times during the launch sequence. The mechanical loading conditions for the full SRB arise from the external tank (ET) attachment points, the solid rocket motor (SRM) pressure load, and the SRB hold down posts. The ET strut loads vary with time after the Space Shuttle main engine (SSME) ignition. The SRM internal pressure varies axially by approximately 100 psi. Static analyses of the full SRB are performed using a snapshot picture of the loads. The field and factory joints are modeled by using equivalent stiffness joints instead of detailed models of the joint. As such, local joint behavior cannot be obtained from this global model.

Apparatus For Tests Of Percussion Primers

NASA Technical Reports Server (NTRS)

Bement, Laurence J.; Bailey, James W.; Schimmel, Morry L.

1991-01-01

Test apparatus and method developed to measure ignition capability of percussion primers. Closely simulates actual conditions and interfaces encountered in such applications as in munitions and rocket motors. Ignitability-testing apparatus is small bomb instrumented with pressure transducers. Sizes, shapes, and positions of bomb components and materials under test selected to obtain quantitative data on ignition.

X-1 launch from B-29 mothership

NASA Technical Reports Server (NTRS)

1947-01-01

The first of the rocket-powered research aircraft, the X-1 (originally designated the XS-1), was a bullet-shaped airplane that was built by the Bell Aircraft Company for the US Air Force and the National Advisory Committee on Aeronautics (NACA). The mission of the X-1 was to investigate the transonic speed range (speeds from just below to just above the speed of sound) and, if possible, to break the 'sound barrier'. The first of the three X-1s was glide-tested at Pinecastle Air Force Base, FL, in early 1946. The first powered flight of the X-1 was made on Dec. 9, 1946, at Edwards Air Force Base with Chalmers Goodlin, a Bell test pilot, at the controls. On Oct. 14, 1947, with USAF Captain Charles 'Chuck' Yeager as pilot, the aircraft flew faster than the speed of sound for the first time. Captain Yeager ignited the four-chambered XLR-11 rocket engines after being air-launched from under the bomb bay of a B-29 at 21,000 ft. The 6,000-lb thrust ethyl alcohol/liquid oxygen burning rockets, built by Reaction Motors, Inc., pushed him up to a speed of 700 mph in level flight. Captain Yeager was also the pilot when the X-1 reached its maximum speed of 957 mph. Another USAF pilot. Lt. Col. Frank Everest, Jr., was credited with taking the X-1 to its maximum altitude of 71,902 ft. Eighteen pilots in all flew the X-1s. The number three plane was destroyed in a fire before ever making any powered flights. A single-place monoplane, the X-1 was 31 ft long, 10 ft high, and had a wingspan of 29 ft. It weighed 4,900 lb and carried 8,200 lb of fuel. It had a flush cockpit with a side entrance and no ejection seat. This roughly 30-second video clip shows the X-1 launched from a B-29, ignition of the XLR-11 rocket engine, and the succeeding flight, including a roll. At one point, the video shows observers of the flight from the ground.

Analysis and Evaluation of German Attainments and Research in the Liquid Rocket Engine Field. Volume 4. Propellant Injectors

DTIC Science & Technology

1951-02-01

they were ob- served at a given pressure drop in "cold" testing with water or unreacted propellants. heat-transfer considerations and the location of... water as a coolant in the main chamber. The Winkler injector was used.on a test unit developing a thrust of 220 lb and an exhaust ve- locity of 6370 ft... water . Provision was made for an igniter in the center of the injector. The relatively high performance reported for this unit does not seem to be

Investigation of Non-Conventional Bio-Derived Fuels for Hybrid Rocket Motors

DTIC Science & Technology

2007-08-01

been demonstrated that a hybrid rocket system using 85% hydrogen peroxide ( HTP ) as the oxidizer and polyethylene as the solid fuel can serve as a cost...As with the tests at Surrey, they used a catalyst pack to decompose the HTP for the ignition. This type of process provides a self-ignition behavior...low regression rate as HTP and polyethylene, so it is difficult to obtain high thrust levels. MARS has the distinction of launching the first

Subscale Fast Cookoff Testing and Modeling for the Hazard Assessment of Large Rocket Motors

DTIC Science & Technology

2001-03-01

41 LIST OF TABLES Table 1 Heats of Vaporization Parameter for Two-liner Phase Transformation - Complete Liner Sublimation and/or Combined Liner...One-dimensional 2-D Two-dimensional ALE3D Arbitrary-Lagrange-Eulerian (3-D) Computer Code ALEGRA 3-D Arbitrary-Lagrange-Eulerian Computer Code for...case-liner bond areas and in the grain inner bore to explore the pre-ignition and ignition phases , as well as burning evolution in rocket motor fast

X-ray driven implosions at ignition relevant velocities on the National Ignition Facility

DOE Office of Scientific and Technical Information (OSTI.GOV)

Meezan, N. B.; MacKinnon, A. J.; Hicks, D. G.

2013-05-15

Backlit convergent ablator experiments on the National Ignition Facility [E. I. Moses et al., Phys. Plasmas 16, 041006 (2009)] are indirect drive implosions that study the inflight dynamics of an imploding capsule. Side-on, backlit radiography provides data used by the National Ignition Campaign to measure time-dependent properties of the capsule ablator including its center of mass radius, velocity, and unablated mass. Previously, Callahan [D. A. Callahan et al., Phys. Plasmas 19, 056305 (2012)] and Hicks [D. H. Hicks et al., Phys. Plasmas 19, 122702 (2012)] reported backlit convergent ablator experiments demonstrating velocities approaching those required for ignition. This paper focusesmore » on implosion performance data in the “rocket curve” plane, velocity vs. ablator mass. These rocket curve data, along with supporting numerical simulations, show that the nominal 195 μm-thick ignition capsule would reach the ignition velocity goal V = 370 km/s with low ablator mass remaining–below the goal of M = 0.25 mg. This finding led to experiments with thicker capsule ablators. A recent symmetry capsule experiment with a 20 μm thicker capsule driven by 520 TW, 1.86 MJ laser pulse (along with a companion backlit convergent ablator experiment) appears to have demonstrated V≥350 km/s with ablator mass remaining above the ignition goal.« less

14 CFR 23.1165 - Engine ignition systems.

Code of Federal Regulations, 2013 CFR

2013-01-01

... Controls and Accessories § 23.1165 Engine ignition systems. (a) Each battery ignition system must be... ignition. (e) Each turbine engine ignition system must be independent of any electrical circuit that is not... commuter category airplanes, each turbine engine ignition system must be an essential electrical load. [Doc...

STS-86 Post Flight Presentation

NASA Technical Reports Server (NTRS)

1999-01-01

The Crew of STS-86 Atlantis Shuttle, Commander James D. Wetherbee, Pilot Michael J. Bloomfield, Mission Specialists Vladimar G. Titov, Scott E. Parazynski, Jean-Loup J. M. Chretien, Wendy Lawrence, and David Wolf, narrate the footage of their mission to the Mir International Space Station. Scenes include crew suit up, walk out to the transfer vehicle, strap-in into the shuttle, start of the main engine, ignition of the rocket boosters, and separation of the solid rocket boosters. The crew of Atlantis participates in an exchange of gifts with the members of Mir, and a space walk to recover experiments outside the Mir Space Station. A beautiful panoramic view of Mir above South America is seen. Scenes also depict the closing of Mir's hatch, Atlantis' separation from Mir, and the reentry of the Atlantis Space Shuttle into the Earth's atmosphere.

Development and demonstration of flueric sounding rocket motor ignition

NASA Technical Reports Server (NTRS)

Marchese, V. P.

1974-01-01

An analytical and experimental program is described which established a flueric rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source. An evaluation was made of this concept to determine the margins of the operating range and capabilities of every component of the system. This evaluation included a determination of power supply requirements, ignitor geometry and alinement, ignitor/propellant interfacing and materials and the effects of ambient temperatures and pressure. It was demonstrated that an operator using a simple hand pump for 30 seconds could ignite BKNO3 at a standoff distance of 100 m (330 ft) with the only connection to the ignitor being a piece of plastic pneumatic tubing.

An Inexpensive Source of High Voltage

NASA Astrophysics Data System (ADS)

Saraiva, Carlos

2012-04-01

As a physics teacher I like recycling old apparatus and using them for demonstrations in my classes.1-4 In physics laboratories in schools, sources of high voltage include induction coils or electronic systems that can be bought from companies that sell lab equipment. But these sources can be very expensive. In this article, I will explain how you can use a car ignition coil as a high voltage source. Such a coil can be obtained from an old car found in a car salvage yard and used to power cathode ray tubes and discharge tubes to observe the spectra. It can also be used as a source of ignition to simulate explosive combustion that occurs in car engines, rockets, etc. You can also buy these coils in shops that sell car accessories and they are cheaper than induction coils. In Fig. 1 you can see a coil that I used.

Introduction of laser initiation for the 48-inch Advanced Solid Rocket Motor (ASRM) test motors at Marshall Space Flight Center (MSFC)

NASA Technical Reports Server (NTRS)

Zimmerman, Chris J.; Litzinger, Gerald E.

1993-01-01

The Advanced Solid Rocket Motor is a new design for the Space Shuttle Solid Rocket Booster. The new design will provide more thrust and more payload capability, as well as incorporating many design improvements in all facets of the design and manufacturing process. A 48-inch (diameter) test motor program is part of the ASRM development program. This program has multiple purposes for testing of propellent, insulation, nozzle characteristics, etc. An overview of the evolution of the 48-inch ASRM test motor ignition system which culminated with the implementation of a laser ignition system is presented. The laser system requirements, development, and operation configuration are reviewed in detail.

Geomagnetic effects caused by rocket exhaust jets

NASA Astrophysics Data System (ADS)

Lipko, Yuriy; Pashinin, Aleksandr; Khakhinov, Vitaliy; Rahmatulin, Ravil

2016-09-01

In the space experiment Radar-Progress, we have made 33 series of measurements of geomagnetic variations during ignitions of engines of Progress cargo spacecraft in low Earth orbit. We used magneto-measuring complexes, installed at observatories of the Institute of Solar-Terrestrial Physics of Siberian Branch of the Russian Academy of Sciences, and magnetotelluric equipment of a mobile complex. We assumed that engine running can cause geomagnetic disturbances in flux tubes crossed by the spacecraft. When analyzing experimental data, we took into account space weather factors: solar wind parameters, total daily mid-latitude geomagnetic activity index Kp, geomagnetic auroral electrojet index AE, global geomagnetic activity. The empirical data we obtained indicate that 18 of the 33 series showed geomagnetic variations in various time ranges.

Simultaneous dual mode combustion engine operating on spark ignition and homogenous charge compression ignition

DOEpatents

Fiveland, Scott B.; Wiggers, Timothy E.

2004-06-22

An engine particularly suited to single speed operation environments, such as stationary power generators. The engine includes a plurality of combustion cylinders operable under homogenous charge compression ignition, and at least one combustion cylinder operable on spark ignition concepts. The cylinder operable on spark ignition concepts can be convertible to operate under homogenous charge compression ignition. The engine is started using the cylinders operable under spark ignition concepts.

77 FR 42724 - Agency Information Collection Activities; Submission to OMB for Review and Approval; Comment...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-07-20

... Nonroad Spark-Ignited Engines, New Nonroad Compression-Ignited Engines, and New On-Road Heavy Duty Engines... Compression-ignited Engines, and New On-road Heavy Duty Engines (Renewal). ICR numbers: EPA ICR No. 1852.05... engines, new nonroad compression-ignited engines, and new on- road heavy duty engines. Estimated Number of...

Effect of the Thruster Configurations on a Laser Ignition Microthruster

NASA Astrophysics Data System (ADS)

Koizumi, Hiroyuki; Hamasaki, Kyoichi; Kondo, Ryo; Okada, Keisuke; Nakano, Masakatsu; Arakawa, Yoshihiro

Research and development of small spacecraft have advanced extensively throughout the world and propulsion devices suitable for the small spacecraft, microthruster, is eagerly anticipated. The authors proposed a microthruster using 1—10-mm-size solid propellant. Small pellets of solid propellant are installed in small combustion chambers and ignited by the irradiation of diode laser beam. This thruster is referred as to a laser ignition microthruster. Solid propellant enables large thrust capability and compact propulsion system. To date theories of a solid-propellant rocket have been well established. However, those theories are for a large-size solid propellant and there are a few theories and experiments for a micro-solid rocket of 1—10mm class. This causes the difficulty of the optimum design of a micro-solid rocket. In this study, we have experimentally investigated the effect of thruster configurations on a laser ignition microthruster. The examined parameters are aperture ratio of the nozzle, length of the combustion chamber, area of the nozzle throat, and divergence angle of the nozzle. Specific impulse dependences on those parameters were evaluated. It was found that large fraction of the uncombusted propellant was the main cause of the degrading performance. Decreasing the orifice diameter in the nozzle with a constant open aperture ratio was an effective method to improve this degradation.

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Nelson, Karl W.

2005-01-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bi-propellant engines used in the Apollo program. A multi-use engine. one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, bum duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Astrophysics Data System (ADS)

Brown, Kendall K.; Nelson, Karl W.

2005-02-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bipropellant engines used in the Apollo program. A multi-use engine, one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, burn duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Final Rule for Gasoline Spark-Ignition Marine Engines; Exemptions for New Nonroad Compression-Ignition Engines at or Above 37 Kilowatts and New Nonroad Spark-Ignition Engines at or Below 19 Kilowatts

EPA Pesticide Factsheets

These standards apply for outboard engines, personal watercraft engines, and jet boat engines. This rule also adds a national security exemption for Nonroad Compression-Ignition (CI) and Small SI sectors.

An Application Of High-Speed Photography To The Real Ignition Course Of Composite Propellants

NASA Astrophysics Data System (ADS)

Fusheng, Zhang; Gongshan, Cheng; Yong, Zhang; Fengchun, Li; Fanpei, Lei

1989-06-01

That the actual solid rocket motor behavior and delay time of the ignition of Ap/HTPB composite propellant ignited by high energy pyrotechics contained condensed particles have been investigated is the key of this paper. In experiments, using high speed camera, the pressure transducer, the photodiode and synchro circuit control system designed by us synchronistically observe and record all course and details of the ignition. And pressure signal, photodiode signal and high speed photography frame are corresponded one by one.

Hypergolic ignitor

NASA Technical Reports Server (NTRS)

Taylor, Eric S. (Inventor); Myers, W. Neill (Inventor); Martin, Michael A. (Inventor)

2005-01-01

An ignitor for use with the MC-1 rocket engine has a cartridge bounded by two end caps with rupture disc assemblies connected thereto. A piston assembly within the cartridge moves from one end of the cartridge during the ignition process. The inlet of the ignitor communicates with a supply taken from the discharge of the fuel pump. When the pump is initially started, the pressure differential bursts the first rupture disc to begin the movement of the piston assembly toward the discharge end. The pressurization of the cartridge causes the second rupture disc to rupture and hypergolic fluid contained within the cartridge is discharged out the ignitor outlet.

KSC-2012-4749

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. - The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Rusty Backer

KSC-2012-4761

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. - The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Kenny Allen

KSC-2012-4748

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. - The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Rusty Backer

KSC-2012-4762

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. - The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Kenny Allen

KSC-2012-4713

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Jim Grossmann

KSC-2012-4712

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – Engine ignition under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT signals liftoff of NASA's Radiation Belt Storm Probes, or RBSP, is imminent from Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Jim Grossmann

KSC-2012-4724

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Ben Smegelsky

KSC-2012-4727

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Tony Gray and Robert Murray

KSC-2012-4728

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Tony Gray and Robert Murray

KSC-2012-4729

NASA Image and Video Library

2012-08-30

CAPE CANAVERAL, Fla. – The engines ignite under the United Launch Alliance Atlas V rocket at 4:05 a.m. EDT lifting NASA's Radiation Belt Storm Probes, or RBSP, off Space Launch Complex 41 on Cape Canaveral Air Force Station in Florida. RBSP will explore changes in Earth's space environment caused by the sun -- known as "space weather" -- that can disable satellites, create power-grid failures and disrupt GPS service. The mission also will provide data on the fundamental radiation and particle acceleration processes throughout the universe. For more information on RBSP, visit http://www.nasa.gov/rbsp. Photo credit: NASA/Tony Gray and Robert Murray

XCALIBUR: a Vertical Takeoff TSTO RLV Concept with a HEDM Upperstage and a Scram-Rocket Booster

NASA Astrophysics Data System (ADS)

Bradford, J.

2002-01-01

A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15 staging condition. The upperstage is released from the booster and carries the 6,820 kg of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro- mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). The booster system is enabled main propulsion system which utilizes four RBCC engines. These engines operate in four distinct modes: air- augmented rocket (AAR), ramjet, scram-rocket, and all-rocket. The booster operates in AAR mode from takeoff to Mach 3, with ramjet mode operation from Mach 3 to Mach 6. The rocket re-ignition for scram-rocket mode occurs at Mach 6, with all-rocket mode from Mach 14 to the staging condition. The extended utilization of the scram-rocket mode greatly improves vehicle performance by providing superior vehicle acceleration when compared to the scramjet mode performance over the same flight region. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable maximum dynamic pressure. A complete vehicle system life-cycle analysis was performed in an automated, multi-disciplinary design environment. Automated disciplinary performance analysis tools include: trajectory (POST), propulsion (SCCREAM), aeroheating (TCAT II), and an Excel spreadsheet for component weight estimation. These tools were automated using `file wrappers' in Phoenix Integration's ModelCenter collaborative design environment. Performance tools utilized for the analysis, but not requiring automation included IDEAS for solid modeling and APAS for the aerodynamic analysis. The paper describes the vehicle concept and operation, discussing the types of technologies used and the nominal flight scenario. A brief discussion explaining the decision-making process for the vehicle configuration is included. For cost predictions, NAFCOM-derived cost estimating relationships were used. Economic predictions were developed using a number of codes, including CABAM (financials), AATe (operations), and GTSafetyII (safety and reliability).

Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

ERIC Educational Resources Information Center

Mattox, J. R.

2017-01-01

I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the…

Catalyst Development for Hydrogen Peroxide Rocket Engines

NASA Technical Reports Server (NTRS)

Morlan, P. W.; Wu, P.-K.; Ruttle, D. W.; Fuller, R. P.; Nejad, A. S.; Anderson, W. E.

1999-01-01

The development of various catalysts of hydrogen peroxide was conducted for the applications of liquid rocket engines. The catalyst development includes silver screen technology, solid catalyst technology, and homogeneous catalyst technology. The silver screen technology development was performed with 85% (by weight) hydrogen peroxide. The results of this investigation were used as the basis for the catalyst design of a pressure-fed liquid-fueled upper stage engine. Both silver-plated nickel 200 screens and pure silver screens were used as the active metal catalyst during the investigation, The data indicate that a high decomposition efficiency (greater than 90%) of 85% hydrogen peroxide can be achieved at a bed loading of 0.5 lbm/sq in/sec with both pure silver and silver plated screens. Samarium oxide coating, however, was found to retard the decomposition process and the catalyst bed was flooded at lower bed loading. A throughput of 200 lbm of hydrogen peroxide (1000 second run time) was tested to evaluate the catalyst aging issue and performance degradation was observed starting at approximately 400 seconds. Catalyst beds of 3.5 inch in diameter was fabricated using the same configuration for a 1,000-lbf rocket engine. High decomposition efficiency was obtained with a low pressure drop across the bed. Solid catalyst using precious metal was also developed for the decomposition of hydrogen peroxide from 85% to 98% by weight. Preliminary results show that the catalyst has a strong reactivity even after 15 minutes of peroxide decomposition. The development effort also includes the homogeneous catalyst technology. Various non-toxic catalysts were evaluated with 98% peroxide and hydrocarbon fuels. The results of open cup drop tests indicate an ignition delay around 11 ms.

75 FR 47520 - Standards of Performance for Stationary Compression Ignition and Spark Ignition Internal...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-08-06

... Ignition Internal Combustion Engines AGENCY: Environmental Protection Agency (EPA). ACTION: Extension of... for stationary compression ignition and spark ignition internal combustion engines. In this [[Page... combustion engines. After publication of the proposed rule, EPA received requests from the American Petroleum...

Combustion and Magnetohydrodynamic Processes in Advanced Pulse Detonation Rocket Engines

NASA Astrophysics Data System (ADS)

Cole, Lord Kahil

A number of promising alternative rocket propulsion concepts have been developed over the past two decades that take advantage of unsteady combustion waves in order to produce thrust. These concepts include the Pulse Detonation Rocket Engine (PDRE), in which repetitive ignition, propagation, and reflection of detonations and shocks can create a high pressure chamber from which gases may be exhausted in a controlled manner. The Pulse Detonation Rocket Induced Magnetohydrodynamic Ejector (PDRIME) is a modification of the basic PDRE concept, developed by Cambier (1998), which has the potential for performance improvements based on magnetohydrodynamic (MHD) thrust augmentation. The PDRIME has the advantage of both low combustion chamber seeding pressure, per the PDRE concept, and efficient energy distribution in the system, per the rocket-induced MHD ejector (RIME) concept of Cole, et al. (1995). In the initial part of this thesis, we explore flow and performance characteristics of different configurations of the PDRIME, assuming quasi-one-dimensional transient flow and global representations of the effects of MHD phenomena on the gas dynamics. By utilizing high-order accurate solvers, we thus are able to investigate the fundamental physical processes associated with the PDRIME and PDRE concepts and identify potentially promising operating regimes. In the second part of this investigation, the detailed coupling of detonations and electric and magnetic fields are explored. First, a one-dimensional spark-ignited detonation with complex reaction kinetics is fully evaluated and the mechanisms for the different instabilities are analyzed. It is found that complex kinetics in addition to sufficient spatial resolution are required to be able to quantify high frequency as well as low frequency detonation instability modes. Armed with this quantitative understanding, we then examine the interaction of a propagating detonation and the applied MHD, both in one-dimensional and two-dimensional transient simulations. The dynamics of the detonation are found to be affected by the application of magnetic and electric fields. We find that the regularity of one-dimensional cesium-seeded detonations can be significantly altered by reasonable applied magnetic fields (Bz ≤ 8T), but that it takes a stronger applied field (Bz > 16T) to significantly alter the cellular structure and detonation velocity of a two-dimensional detonation in the time in which these phenomena were observed. This observation is likely attributed to the additional coupling of the two-dimensional detonation with the transverse waves, which are not captured in the one-dimensional simulations. Future studies involving full ionization kinetics including collisional-radiative processes, will be used to examine these processes in further detail.

Two Phase Detonation Studies Conducted in 1971

NASA Technical Reports Server (NTRS)

Nicholls, J. A.

1972-01-01

The research covered by this third annual progress report represents a continuation of our efforts devoted to the study of detonation waves in liquid-gas systems. The motivation for the work is associated with liquid propellant rocket motor combustion instability although certainly the studies are also applicable to internal combustion engines, jet propulsion engines, safety aspects of spilled liquid fuel, coal mine explosions, and weaponry. The research has been divided into 5 phases, although all of them are intimately related. For the most part these phases are briefly summarized and the reader is referred to other publications for a more complete treatment. The exception to this is where the material herein represents the only printed information available on the particular facet of the problem. Phase A has been primarily concerned with the breakup and ignition of fuel drops by shock waves. The experimental portion of this study as well as a theoretical treatment of the ignition behavior was completed in the past year. The research is now concentrating on the passage of a shock wave over a burning drop. Phase B has been devoted to the assessment of the approximate energy release pattern in two phase detonations insofar as they affect the significant overpressures observed.

14 CFR 33.37 - Ignition system.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Ignition system. 33.37 Section 33.37... STANDARDS: AIRCRAFT ENGINES Design and Construction; Reciprocating Aircraft Engines § 33.37 Ignition system. Each spark ignition engine must have a dual ignition system with at least two spark plugs for each...

Ignition and flame stabilization of a strut-jet RBCC combustor with small rocket exhaust.

PubMed

Hu, Jichao; Chang, Juntao; Bao, Wen

2014-01-01

A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes.

Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

PubMed Central

2014-01-01

A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

14 CFR 23.1165 - Engine ignition systems.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine ignition systems. 23.1165 Section 23... Controls and Accessories § 23.1165 Engine ignition systems. (a) Each battery ignition system must be... allow continued engine operation if any battery becomes depleted. (b) The capacity of batteries and...

14 CFR 29.1165 - Engine ignition systems.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine ignition systems. 29.1165 Section 29... Engine ignition systems. (a) Each battery ignition system must be supplemented with a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any...

14 CFR 23.1165 - Engine ignition systems.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine ignition systems. 23.1165 Section 23... Controls and Accessories § 23.1165 Engine ignition systems. (a) Each battery ignition system must be... allow continued engine operation if any battery becomes depleted. (b) The capacity of batteries and...

14 CFR 29.1165 - Engine ignition systems.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine ignition systems. 29.1165 Section 29... Engine ignition systems. (a) Each battery ignition system must be supplemented with a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any...

14 CFR 23.1165 - Engine ignition systems.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Engine ignition systems. 23.1165 Section 23... Controls and Accessories § 23.1165 Engine ignition systems. (a) Each battery ignition system must be... allow continued engine operation if any battery becomes depleted. (b) The capacity of batteries and...

14 CFR 29.1165 - Engine ignition systems.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Engine ignition systems. 29.1165 Section 29... Engine ignition systems. (a) Each battery ignition system must be supplemented with a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any...

14 CFR 25.1165 - Engine ignition systems.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Engine ignition systems. 25.1165 Section 25... Engine ignition systems. (a) Each battery ignition system must be supplemented by a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any...

Forced Ignition Study Based On Wavelet Method

NASA Astrophysics Data System (ADS)

Martelli, E.; Valorani, M.; Paolucci, S.; Zikoski, Z.

2011-05-01

The control of ignition in a rocket engine is a critical problem for combustion chamber design. Therefore it is essential to fully understand the mechanism of ignition during its earliest stages. In this paper the characteristics of flame kernel formation and initial propagation in a hydrogen-argon-oxygen mixing layer are studied using 2D direct numerical simulations with detailed chemistry and transport properties. The flame kernel is initiated by adding an energy deposition source term in the energy equation. The effect of unsteady strain rate is studied by imposing a 2D turbulence velocity field, which is initialized by means of a synthetic field. An adaptive wavelet method, based on interpolating wavelets is used in this study to solve the compressible reactive Navier- Stokes equations. This method provides an alternative means to refine the computational grid points according to local demands of the physical solution. The present simulations show that in the very early instants the kernel perturbed by the turbulent field is characterized by an increased burning area and a slightly increased rad- ical formation. In addition, the calculations show that the wavelet technique yields a significant reduction in the number of degrees of freedom necessary to achieve a pre- scribed solution accuracy.

Characterization of the Ignition Over-Pressure/Sound Suppression Water in the Space Launch System Mobile Launcher Using Volume of Fluid Modeling

NASA Technical Reports Server (NTRS)

West, Jeff

2015-01-01

The Space Launch System (SLS) Vehicle consists of a Core Stage with four RS-25 engines and two Solid Rocket Boosters (SRBs). This vehicle is launched from the Launchpad using a Mobile Launcher (ML) which supports the SLS vehicle until its liftoff from the ML under its own power. The combination of the four RS-25 engines and two SRBs generate a significant Ignition Over-Pressure (IOP) and Acoustic Sound environment. One of the mitigations of these environments is the Ignition Over-Pressure/Sound Suppression (IOP/SS) subsystem installed on the ML. This system consists of six water nozzles located parallel to and 24 inches downstream of each SRB nozzle exit plane as well as 16 water nozzles located parallel to and 53 inches downstream of the RS-25 nozzle exit plane. During launch of the SLS vehicle, water is ejected through each water nozzle to reduce the intensity of the transient pressure environment imposed upon the SLS vehicle. While required for the mitigation of the transient pressure environment on the SLS vehicle, the IOP/SS subsystem interacts (possibly adversely) with other systems located on the Launch Pad. One of the other systems that the IOP/SS water is anticipated to interact with is the Hydrogen Burn-Off Igniter System (HBOI). The HBOI system's purpose is to ignite the unburned hydrogen/air mixture that develops in and around the nozzle of the RS-25 engines during engine start. Due to the close proximity of the water system to the HBOI system, the presence of the IOP/SS may degrade the effectiveness of the HBOI system. Another system that the IOP/SS water may interact with adversely is the RS-25 engine nozzles and the SRB nozzles. The adverse interaction anticipated is the wetting, to a significant degree, of the RS-25 nozzles resulting in substantial weight of ice forming and water present to a significant degree upstream of the SRB nozzle exit plane inside the nozzle itself, posing significant additional blockage of the effluent that exits the nozzle upon motor start leading to detrimental effects. The purpose of the CFD simulations were to i) characterize the location of the IOP/SS water after it is ejected from the IOP/SS nozzles, ii) characterize the interaction of the IOP/SS system with the HBOI system and iii) characterize the interaction of the IOP/SS water with the RS-25 nozzles and the SRB nozzles.

HIFiRE Flight 2 Flowpath Design Update (PREPRINT)

DTIC Science & Technology

2009-12-01

will use a sounding rocket stack and a novel second-stage ignition approach to achieve a nearly constant flight dynamic pressure over this range of...Mach numbers. The experimental payload will remain attached to the second-stage rocket motor and the experiment will occur while accelerating through...weight and drag estimates necessary for trajectory analyses to be conducted using candidate rocket motors . The preliminary trajectory analyses

SLS Test Stand Site Selection

NASA Technical Reports Server (NTRS)

Crowe, Kathryn; Williams, Michael

2015-01-01

Test site selection is a critical element of the design, development and production of a new system. With the advent of the new Space Launch System (SLS), the National Aeronautics and Space Administration (NASA) had a number of test site selection decisions that needed to be made early enough in the Program to support the planned Launch Readiness Date (LRD). This case study focuses on decisions that needed to be made in 2011 and 2012 in preparation for the April 2013 DPMC decision about where to execute the Main Propulsion Test that is commonly referred to as "Green Run." Those decisions relied upon cooperative analysis between the Program, the Test Lab and Center Operations. The SLS is a human spaceflight vehicle designed to carry a crew farther into space than humans have previously flown. The vehicle consists of four parts: the crew capsule, the upper stage, the core stage, and the first stage solid rocket boosters. The crew capsule carries the astronauts, while the upper stage, the core stage, and solid rocket boosters provide thrust for the vehicle. In other words, the stages provide the "lift" part of the lift vehicle. In conjunction with the solid rocket boosters, the core stage provides the initial "get-off-the-ground" thrust to the vehicle. The ignition of the four core stage engines and two solid rocket boosters is the first step in the launch portion of the mission. The solid rocket boosters burn out after about 2 minutes of flight, and are then jettisoned. The core stage provides thrust until the vehicle reaches a specific altitude and speed, at which point the core stage is shut off and jettisoned, and the upper stage provides vehicle thrust for subsequent mission trajectories. The integrated core stage primarily consists of a liquid oxygen tank, a liquid hydrogen tank, and the four core stage engines. For the SLS program, four RS-25 engines were selected as the four core stage engines. The RS-25 engine is the same engine that was used for Space Shuttle. The test plan for the integrated core stage was broken down into several segments: Component testing, system level testing, and element level testing. In this context, components are items such as valves, controllers, sensors, etc. Systems are items such as an entire engine, a tank, or the outer stage body. The core stage itself is considered to be an element. The rocket engines are also considered an element. At the program level, it was decided to perform a single green run test on the integrated core stage prior to shipment of it to Kennedy Space Center (KSC) for use in the EM-1 test flight of the SLS vehicle. A green run test is the first live fire of the new integrated core stage and engine elements - without boosters of course. The SLS Program had to decide where to perform SLS green run testing.

Orbit transfer rocket engine technology program

NASA Technical Reports Server (NTRS)

Gustafson, N. B.; Harmon, T. J.

1993-01-01

An advanced near term (1990's) space-based Orbit Transfer Vehicle Engine (OTVE) system was designed, and the technologies applicable to its construction, maintenance, and operations were developed under Tasks A through F of the Orbit Transfer Rocket Engine Technology Program. Task A was a reporting task. In Task B, promising OTV turbomachinery technologies were explored: two stage partial admission turbines, high velocity ratio diffusing crossovers, soft wear ring seals, advanced bearing concepts, and a rotordynamic analysis. In Task C, a ribbed combustor design was developed. Possible rib and channel geometries were chosen analytically. Rib candidates were hot air tested and laser velocimeter boundary layer analyses were conducted. A channel geometry was also chosen on the basis of laser velocimeter data. To verify the predicted heat enhancement effects, a ribbed calorimeter spool was hot fire tested. Under Task D, the optimum expander cycle engine thrust, performance and envelope were established for a set of OTV missions. Optimal nozzle contours and quick disconnects for modularity were developed. Failure Modes and Effects Analyses, maintenance and reliability studies and component study results were incorporated into the engine system. Parametric trades on engine thrust, mixture ratio, and area ratio were also generated. A control system and the health monitoring and maintenance operations necessary for a space-based engine were outlined in Task E. In addition, combustor wall thickness measuring devices and a fiberoptic shaft monitor were developed. These monitoring devices were incorporated into preflight engine readiness checkout procedures. In Task F, the Integrated Component Evaluator (I.C.E.) was used to demonstrate performance and operational characteristics of an advanced expander cycle engine system and its component technologies. Sub-system checkouts and a system blowdown were performed. Short transitions were then made into main combustor ignition and main stage operation.

NASA's Advanced solid rocket motor

NASA Technical Reports Server (NTRS)

Mitchell, Royce E.

1993-01-01

The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

Analysis and modeling of infrasound from a four-stage rocket launch.

PubMed

Blom, Philip; Marcillo, Omar; Arrowsmith, Stephen

2016-06-01

Infrasound from a four-stage sounding rocket was recorded by several arrays within 100 km of the launch pad. Propagation modeling methods have been applied to the known trajectory to predict infrasonic signals at the ground in order to identify what information might be obtained from such observations. There is good agreement between modeled and observed back azimuths, and predicted arrival times for motor ignition signals match those observed. The signal due to the high-altitude stage ignition is found to be low amplitude, despite predictions of weak attenuation. This lack of signal is possibly due to inefficient aeroacoustic coupling in the rarefied upper atmosphere.

14 CFR 25.1145 - Ignition switches.

Code of Federal Regulations, 2013 CFR

2013-01-01

... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 25.1145 - Ignition switches.

Code of Federal Regulations, 2011 CFR

2011-01-01

... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 25.1145 - Ignition switches.

Code of Federal Regulations, 2014 CFR

2014-01-01

... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 25.1145 - Ignition switches.

Code of Federal Regulations, 2010 CFR

2010-01-01

... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 25.1145 - Ignition switches.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

Performance Increase Verification for a Bipropellant Rocket Engine

NASA Technical Reports Server (NTRS)

Alexander, Leslie; Chapman, Jack; Wilson, Reed; Krismer, David; Lu, Frank; Wilson, Kim; Miller, Scott; England, Chris

2008-01-01

Component performance assessment testing for a, pressure-fed earth storable bipropellant rocket engine was successfully completed at Aerojet's Redmond test facility. The primary goal of the this development project is to increase the specific impulse of an apogee class bi-propellant engine to greater than 330 seconds with nitrogen tetroxide and monomethylhydrazine propellants and greater than 335 seconds with nitrogen tetroxide and hydrazine. The secondary goal of the project is to take greater advantage of the high temperature capabilities of iridium/rhenium chambers. In order to achieve these goals, the propellant feed pressures were increased to 400 psia, nominal, which in turn increased the chamber pressure and temperature, allowing for higher c*. The tests article used a 24-on-24 unlike doublet injector design coupled with a copper heat sink chamber to simulate a flight configuration combustion chamber. The injector is designed to produce a nominal 200 lbf of thrust with a specific impulse of 335 seconds (using hydrazine fuel). Effect of Chamber length on engine C* performance was evaluated with the use of modular, bolt-together test hardware and removable chamber inserts. Multiple short duration firings were performed to characterize injector performance across a range of thrust levels, 180 to 220 lbf, and mixture ratios, from 1.1 to 1.3. During firing, ignition transient, chamber pressure, and various temperatures were measured in order to evaluate the performance of the engine and characterize the thermal conditions. The tests successfully demonstrated the stable operation and performance potential of a full scale engine with a measured c* of XXXX ft/sec (XXXX m/s) under nominal operational conditions.

STS-95 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1998-01-01

On this first day of the STS-95 mission, the flight crew, Cmdr. Curtis L. Brown, Pilot Steven W. Lindsey, Mission Specialists Scott E. Parazynski, Stephen K. Robinson, and Pedro Duque, and Payload Specialists Chiaki Mukai and John H. Glenn, can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-87 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1997-01-01

On this first day of the STS-87 mission, the flight crew, Cmdr. Kevin R. Kregel, Pilot Steven W. Lindsey, Mission Specialists Winston E. Scott, Kalpana Chawla, and Takao Doi, and Payload Specialist Leonid K. Kadenyuk can be seen preforming pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is seen being readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-88 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1998-01-01

On this first day of the STS-88 mission, the flight crew, Commander Robert D. Cabana, Pilot Frederick W. Sturckow, and Mission Specialists Nancy J. Currie, James H. Newman, Jerry L. Ross, and Sergei Krikalev can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the "white room" for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

Performance of a Laser Ignited Multicylinder Lean Burn Natural Gas Engine

DOE PAGES

Almansour, Bader; Vasu, Subith; Gupta, Sreenath B.; ...

2017-06-06

Market demands for lower fueling costs and higher specific powers in stationary natural gas engines has engine designs trending towards higher in-cylinder pressures and leaner combustion operation. However, Ignition remains as the main limiting factor in achieving further performance improvements in these engines. Addressing this concern, while incorporating various recent advances in optics and laser technologies, laser igniters were designed and developed through numerous iterations. Final designs incorporated water-cooled, passively Q-switched, Nd:YAG micro-lasers that were optimized for stable operation under harsh engine conditions. Subsequently, the micro-lasers were installed in the individual cylinders of a lean-burn, 350 kW, inline 6-cylinder, open-chamber,more » spark ignited engine and tests were conducted. To the best of our knowledge, this is the world’s first demonstration of a laser ignited multi-cylinder natural gas engine. The engine was operated at high-load (298 kW) and rated speed (1800 rpm) conditions. Ignition timing sweeps and excess-air ratio (λ) sweeps were performed while keeping the NOx emissions below the USEPA regulated value (BSNOx < 1.34 g/kW-hr), and while maintaining ignition stability at industry acceptable values (COV_IMEP <5 %). Through such engine tests, the relative merits of (i) standard electrical ignition system, and (ii) laser ignition system were determined. In conclusion, a rigorous combustion data analysis was performed and the main reasons leading to improved performance in the case of laser ignition were identified.« less

Performance of a Laser Ignited Multicylinder Lean Burn Natural Gas Engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Almansour, Bader; Vasu, Subith; Gupta, Sreenath B.

Market demands for lower fueling costs and higher specific powers in stationary natural gas engines has engine designs trending towards higher in-cylinder pressures and leaner combustion operation. However, Ignition remains as the main limiting factor in achieving further performance improvements in these engines. Addressing this concern, while incorporating various recent advances in optics and laser technologies, laser igniters were designed and developed through numerous iterations. Final designs incorporated water-cooled, passively Q-switched, Nd:YAG micro-lasers that were optimized for stable operation under harsh engine conditions. Subsequently, the micro-lasers were installed in the individual cylinders of a lean-burn, 350 kW, inline 6-cylinder, open-chamber,more » spark ignited engine and tests were conducted. To the best of our knowledge, this is the world’s first demonstration of a laser ignited multi-cylinder natural gas engine. The engine was operated at high-load (298 kW) and rated speed (1800 rpm) conditions. Ignition timing sweeps and excess-air ratio (λ) sweeps were performed while keeping the NOx emissions below the USEPA regulated value (BSNOx < 1.34 g/kW-hr), and while maintaining ignition stability at industry acceptable values (COV_IMEP <5 %). Through such engine tests, the relative merits of (i) standard electrical ignition system, and (ii) laser ignition system were determined. In conclusion, a rigorous combustion data analysis was performed and the main reasons leading to improved performance in the case of laser ignition were identified.« less

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions. Freejet tests of a candidate flowpath for this RBCC engine were conducted at the NASA Lewis Research Center's Hypersonic Tunnel Facility between July and September 1996. This paper describes the engine flowpath and installation, outlines the primary objectives of the program, and describes the overall results of this activity. Through this program 15 full duration tests, including 13 fueled tests were made. The first major achievement was the further demonstration of the HTF capability. The facility operated at conditions up to 1950 K and 7.34 MPa, simulating approximately Mach 6.6 flight. The initial tests were unfueled and focused on verifying both facility and engine starting. During these runs additional aerodynamic appliances were incorporated onto the facility diffuser to enhance starting. Both facility and engine starting were achieved. Further, the static pressure distributions compared well with the results previously obtained in a 40% subscale flowpath study conducted in the LERC 1X1 supersonic wind tunnel (SWT), as well as the results of CFD analysis. Fueled performance results were obtained for the engine at both simulated Mach 6 (1670 K) and Mach 6.6 (1950 K) conditions. For all these tests the primary fuel was liquid JP-10 with gaseous silane (a mixture of 20% SiH4 and 80% H2 by volume) as an ignitor/pilot. These tests verified performance of this engine flowpath in a freejet mode. High combustor pressures were reached and significant changes in axial force were achieved due to combustion. Future test plans include redistributing the fuel to improve mixing, and consequently performance, at higher equivalence ratios.

Spread Across Liquids: The World's First Microgravity Combustion Experiment on a Sounding Rocket

NASA Technical Reports Server (NTRS)

1995-01-01

The Spread Across Liquids (SAL) experiment characterizes how flames spread over liquid pools in a low-gravity environment in comparison to test data at Earth's gravity and with numerical models. The modeling and experimental data provide a more complete understanding of flame spread, an area of textbook interest, and add to our knowledge about on-orbit and Earthbound fire behavior and fire hazards. The experiment was performed on a sounding rocket to obtain the necessary microgravity period. Such crewless sounding rockets provide a comparatively inexpensive means to fly very complex, and potentially hazardous, experiments and perform reflights at a very low additional cost. SAL was the first sounding-rocket-based, microgravity combustion experiment in the world. It was expected that gravity would affect ignition susceptibility and flame spread through buoyant convection in both the liquid pool and the gas above the pool. Prior to these sounding rocket tests, however, it was not clear whether the fuel would ignite readily and whether a flame would be sustained in microgravity. It also was not clear whether the flame spread rate would be faster or slower than in Earth's gravity.

14 CFR 33.37 - Ignition system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Ignition system. 33.37 Section 33.37 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Reciprocating Aircraft Engines § 33.37 Ignition system. Each spark ignition engine must have a...

Combustion and operating characteristics of spark-ignition engines

NASA Technical Reports Server (NTRS)

Heywood, J. B.; Keck, J. C.; Beretta, G. P.; Watts, P. A.

1980-01-01

The spark-ignition engine turbulent flame propagation process was investigated. Then, using a spark-ignition engine cycle simulation and combustion model, the impact of turbocharging and heat transfer variations or engine power, efficiency, and NO sub x emissions was examined.

A sustained-arc ignition system for internal combustion engines

NASA Technical Reports Server (NTRS)

Birchenough, A. G.

1977-01-01

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.

14 CFR 33.69 - Ignitions system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.69 Ignitions system. Each engine must be equipped with an ignition system for starting the engine on the ground and in flight. An...

14 CFR 33.69 - Ignitions system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.69 Ignitions system. Each engine must be equipped with an ignition system for starting the engine on the ground and in flight. An...

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 2: a Study of Low Frequency Combustion Instability in Rocket Engine Preburners Using a Heterogeneous Stirred Tank Reactor Model. Final Report M.S. Thesis - Aug. 1987

NASA Technical Reports Server (NTRS)

Bartrand, Timothy A.

1988-01-01

During the shutdown of the space shuttle main engine, oxygen flow is shut off from the fuel preburner and helium is used to push the residual oxygen into the combustion chamber. During this process a low frequency combustion instability, or chug, occurs. This chug has resulted in damage to the engine's augmented spark igniter due to backflow of the contents of the preburner combustion chamber into the oxidizer feed system. To determine possible causes and fixes for the chug, the fuel preburner was modeled as a heterogeneous stirred tank combustion chamber, a variable mass flow rate oxidizer feed system, a constant mass flow rate fuel feed system and an exit turbine. Within the combustion chamber gases were assumed perfectly mixed. To account for liquid in the combustion chamber, a uniform droplet distribution was assumed to exist in the chamber, with mean droplet diameter determined from an empirical relation. A computer program was written to integrate the resulting differential equations. Because chamber contents were assumed perfectly mixed, the fuel preburner model erroneously predicted that combustion would not take place during shutdown. The combustion rate model was modified to assume that all liquid oxygen that vaporized instantaneously combusted with fuel. Using this combustion model, the effect of engine parameters on chamber pressure oscillations during the SSME shutdown was calculated.

40 CFR 1065.230 - Raw exhaust flow meter.

Code of Federal Regulations, 2012 CFR

2012-07-01

... sample NMHC downstream of the cooling for compression-ignition engines, 2-stroke spark-ignition engines, and 4-stroke spark-ignition engines below 19 kW. (3) If cooling causes aqueous condensation, do not...

40 CFR 1065.230 - Raw exhaust flow meter.

Code of Federal Regulations, 2013 CFR

2013-07-01

... sample NMHC downstream of the cooling for compression-ignition engines, 2-stroke spark-ignition engines, and 4-stroke spark-ignition engines below 19 kW. (3) If cooling causes aqueous condensation, do not...

Laser Ignition Microthruster Experiments on KKS-1

NASA Astrophysics Data System (ADS)

Nakano, Masakatsu; Koizumi, Hiroyuki; Watanabe, Masashi; Arakawa, Yoshihiro

A laser ignition microthruster has been developed for microsatellites. Thruster performances such as impulse and ignition probability were measured, using boron potassium nitrate (B/KNO3) solid propellant ignited by a 1 W CW laser diode. The measured impulses were 60 mNs ± 15 mNs with almost 100 % ignition probability. The effect of the mixture ratios of B/KNO3 on thruster performance was also investigated, and it was shown that mixture ratios between B/KNO3/binder = 28/70/2 and 38/60/2 exhibited both high ignition probability and high impulse. Laser ignition thrusters designed and fabricated based on these data became the first non-conventional microthrusters on the Kouku Kousen Satellite No. 1 (KKS-1) microsatellite that was launched by a H2A rocket as one of six piggyback satellites in January 2009.

Experimental Investigation of Piston Heat Transfer in a Light Duty Engine Under Conventional Diesel, Homogeneous Charge Compression Ignition, and Reactivity Controlled Compression Ignition Combustion Regimes

DTIC Science & Technology

2014-01-15

in a Light Duty Engine Under Conventional Diesel, Homogeneous Charge Compression Ignition , and Reactivity Controlled Compression Ignition ...Conventional Diesel (CDC), Homogeneous Charge Compression Ignition (HCCI), and Reactivity Controlled Compression Ignition (RCCI) combustion...LTC) regimes, including reactivity controlled compression ignition (RCCI), partially premixed combustion (PPC), and homogenous charge compression

Posttest report for the Advanced Solid Rocket Motor (ASRM) igniter discharge port flow test

NASA Technical Reports Server (NTRS)

Sringer, Anthony M.

1993-01-01

The primary purpose of this test was to determine discharge coefficients for both the center axial and radial 2:1 aspect ratio exhaust ports of the ASRM multi-port igniter. In addition, both ports were tested with chamfered leading edge to assess how much improvement in discharge coefficient could potentially be achieved.

40 CFR 1065.230 - Raw exhaust flow meter.

Code of Federal Regulations, 2014 CFR

2014-07-01

... sample NMHC downstream of the cooling for compression-ignition engines, two-stroke spark-ignition engines, or four-stroke spark-ignition engines at or below 19 kW. (3) The cooling must not cause aqueous...

Analysis and modeling of infrasound from a four-stage rocket launch

DOE PAGES

Blom, Philip Stephen; Marcillo, Omar Eduardo; Arrowsmith, Stephen

2016-06-17

Infrasound from a four-stage sounding rocket was recorded by several arrays within 100 km of the launch pad. Propagation modeling methods have been applied to the known trajectory to predict infrasonic signals at the ground in order to identify what information might be obtained from such observations. There is good agreement between modeled and observed back azimuths, and predicted arrival times for motor ignition signals match those observed. The signal due to the high-altitude stage ignition is found to be low amplitude, despite predictions of weak attenuation. As a result, this lack of signal is possibly due to inefficient aeroacousticmore » coupling in the rarefied upper atmosphere.« less

Analysis and modeling of infrasound from a four-stage rocket launch

DOE Office of Scientific and Technical Information (OSTI.GOV)

Blom, Philip Stephen; Marcillo, Omar Eduardo; Arrowsmith, Stephen

Infrasound from a four-stage sounding rocket was recorded by several arrays within 100 km of the launch pad. Propagation modeling methods have been applied to the known trajectory to predict infrasonic signals at the ground in order to identify what information might be obtained from such observations. There is good agreement between modeled and observed back azimuths, and predicted arrival times for motor ignition signals match those observed. The signal due to the high-altitude stage ignition is found to be low amplitude, despite predictions of weak attenuation. As a result, this lack of signal is possibly due to inefficient aeroacousticmore » coupling in the rarefied upper atmosphere.« less

Potential of Spark Ignition Engine, 1979 Summary Source Document

DOT National Transportation Integrated Search

1980-03-01

This report provides an assessment of the potential for spark ignition engines passenger cars and light trucks. The assessment includes: tradeoffs between fuel economy and emissions; improvements in spark ignition engine efficiency; improvements in e...

LOX/Methane Main Engine Igniter Tests and Modeling

NASA Technical Reports Server (NTRS)

Breisacher, Kevin J.; Ajmani, Kumund

2008-01-01

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.

40 CFR 1065.240 - Dilution air and diluted exhaust flow meters.

Code of Federal Regulations, 2013 CFR

2013-07-01

...) AIR POLLUTION CONTROLS ENGINE-TESTING PROCEDURES Measurement Instruments Flow-Related Measurements... interval. You may use the difference between a diluted exhaust flow meter and a dilution air meter to... compression-ignition engines, 2-stroke spark-ignition engines, and 4-stroke spark-ignition engines below 19 kW...

40 CFR 1065.240 - Dilution air and diluted exhaust flow meters.

Code of Federal Regulations, 2014 CFR

2014-07-01

...) AIR POLLUTION CONTROLS ENGINE-TESTING PROCEDURES Measurement Instruments Flow-Related Measurements... interval. You may use the difference between a diluted exhaust flow meter and a dilution air meter to... compression-ignition engines, two-stroke spark-ignition engines, or four-stroke spark-ignition engines at or...

40 CFR 1065.240 - Dilution air and diluted exhaust flow meters.

Code of Federal Regulations, 2011 CFR

2011-07-01

...) AIR POLLUTION CONTROLS ENGINE-TESTING PROCEDURES Measurement Instruments Flow-Related Measurements... interval. You may use the difference between a diluted exhaust flow meter and a dilution air meter to... compression-ignition engines, 2-stroke spark-ignition engines, and 4-stroke spark-ignition engines below 19 kW...

40 CFR 1065.240 - Dilution air and diluted exhaust flow meters.

Code of Federal Regulations, 2012 CFR

2012-07-01

...) AIR POLLUTION CONTROLS ENGINE-TESTING PROCEDURES Measurement Instruments Flow-Related Measurements... interval. You may use the difference between a diluted exhaust flow meter and a dilution air meter to... compression-ignition engines, 2-stroke spark-ignition engines, and 4-stroke spark-ignition engines below 19 kW...

14 CFR 25.1165 - Engine ignition systems.

Code of Federal Regulations, 2010 CFR

2010-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw electrical energy from the same source. (c) The design of the engine ignition system must...

14 CFR 25.1165 - Engine ignition systems.

Code of Federal Regulations, 2011 CFR

2011-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw electrical energy from the same source. (c) The design of the engine ignition system must...

Three-dimensional multi-physics coupled simulation of ignition transient in a dual pulse solid rocket motor

NASA Astrophysics Data System (ADS)

Li, Yingkun; Chen, Xiong; Xu, Jinsheng; Zhou, Changsheng; Musa, Omer

2018-05-01

In this paper, numerical investigation of ignition transient in a dual pulse solid rocket motor has been conducted. An in-house code has been developed in order to solve multi-physics governing equations, including unsteady compressible flow, heat conduction and structural dynamic. The simplified numerical models for solid propellant ignition and combustion have been added. The conventional serial staggered algorithm is adopted to simulate the fluid structure interaction problems in a loosely-coupled manner. The accuracy of the coupling procedure is validated by the behavior of a cantilever panel subjected to a shock wave. Then, the detailed flow field development, flame propagation characteristics, pressure evolution in the combustion chamber, and the structural response of metal diaphragm are analyzed carefully. The burst-time and burst-pressure of the metal diaphragm are also obtained. The individual effects of the igniter's mass flow rate, metal diaphragm thickness and diameter on the ignition transient have been systemically compared. The numerical results show that the evolution of the flow field in the combustion chamber, the temperature distribution on the propellant surface and the pressure loading on the metal diaphragm surface present a strong three-dimensional behavior during the initial ignition stage. The rupture of metal diaphragm is not only related to the magnitude of pressure loading on the diaphragm surface, but also to the history of pressure loading. The metal diaphragm thickness and diameter have a significant effect on the burst-time and burst-pressure of metal diaphragm.

Analysis of the laser ignition of methane/oxygen mixtures in a sub-scale rocket combustion chamber

NASA Astrophysics Data System (ADS)

Wohlhüter, Michael; Zhukov, Victor P.; Sender, Joachim; Schlechtriem, Stefan

2017-06-01

The laser ignition of methane/oxygen mixtures in a sub-scale rocket combustion chamber has been investigated numerically and experimentally. The ignition test case used in the present paper was generated during the In-Space Propulsion project (ISP-1), a project focused on the operation of propulsion systems in space, the handling of long idle periods between operations, and multiple reignitions under space conditions. Regarding the definition of the numerical simulation and the suitable domain for the current model, 2D and 3D simulations have been performed. Analysis shows that the usage of a 2D geometry is not suitable for this type of simulation, as the reduction of the geometry to a 2D domain significantly changes the conditions at the time of ignition and subsequently the flame development. The comparison of the numerical and experimental results shows a strong discrepancy in the pressure evolution and the combustion chamber pressure peak following the laser spark. The detailed analysis of the optical Schlieren and OH data leads to the conclusion that the pressure measurement system was not able to capture the strong pressure increase and the peak value in the combustion chamber during ignition. Although the timing in flame development following the laser spark is not captured appropriately, the 3D simulations reproduce the general ignition phenomena observed in the optical measurement systems, such as pressure evolution and injector flow characteristics.

Shakedown and Preliminary Calibration Tests for the Fuel Engine Evaluation System Using the KM914A Sachs Rotary Combustion Engine.

DTIC Science & Technology

1981-12-01

obtained recommendations are made to improve the system. FEES was designed to handle spark ignition and compression ignition research engines of...Thermometer T W OF Temperature Web Bulb Sling Psychrometer % Relative Humidity Psychrometric chart mm Hg Vapor Pressure Vapor Pressure chart - Correction...results obtained recommendations are made to improve the system. FEES was designed to handle spark ignition and compression ignition research engines of

STS-85 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1997-01-01

On this first day of the STS-85 mission, the flight crew, Cmdr. Curtis L. Brown, Jr., Pilot Kent V. Rominger, Payload Cmdr. N. Jan Davis (Ph.D.), Mission Specialists Robert L. Curbeam, Jr., and Stephen K. Robinson (Ph.D.), and Payload Specialist Bjarni V. Tryggvason can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew can be seen being readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-96 FD Highlights and Crew Activities Report: Flight Day 01

NASA Technical Reports Server (NTRS)

1999-01-01

On this first day of the STS-96 Discovery mission, the flight crew, Commander Kent V. Rominger, Pilot Rick D. Husband, and Mission Specialists Ellen Ochoa, Tamara E. Jernigan, Daniel T. Barry, Julie Payette, and Valery Ivanovich Tokarev are seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-89 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1998-01-01

On this first day of the STS-89 mission, the flight crew, Cmdr. Terrence W. Wilcutt, Pilot Frank Edwards, and Mission Specialists Michael P. Anderson, James F. Reilly, Bonnie J. Dunbar, Salizhan Shakirovich Sharipov, David A. Wolf and Andrew S.W. Thomas, can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-86 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1997-01-01

On this first day of the STS-86 mission, the flight crew, Cmdr. James D. Wetherbee, Jr., Pilot Michael J. Bloomfield, Mission Specialists Scott E. Parazynski, Jean-Loup Chretien, Vladimir G. Titov, Wendy B. Lawrence and David A. Wolf can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also included are various panoramic views of the shuttle on the pad. The crew can be seen being readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-91 Flight Day 1 Highlights and Crew Activities Report

NASA Technical Reports Server (NTRS)

1998-01-01

On this first day of the STS-91 mission, the flight crew, Cmdr. Charles J. Precourt, Pilot Dominic L. Pudwill Gorie, and Mission Specialists Franklin R. Chang-Diaz, Janet Lynn Kavandi, Wendy B. Lawrence, Valery Victorovitch Ryumin and Andrew S. W. Thomas, can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-83 Day 01

NASA Technical Reports Server (NTRS)

1997-01-01

On this first day of the STS-83 mission, the flight crew, Cmdr. James D. Halsell Jr., Pilot Susan L. Still, Payload Cmdr. Janice E. Voss, Mission Specialists Michael L. Gernhardt and Donald A. Thomas, and Payload Specialists Gregory T. Linteris and Roger K. Crouch can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew can be seen being readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-81 Flight Day 1

NASA Technical Reports Server (NTRS)

1997-01-01

This first day of the STS-81 mission begins with the flight crew, Cmdr. Michael A. Baker, Pilot Brent W. Jett, Mission Specialists, John M. Grunsfeld, Marsha S. Ivins, Peter J.K. Wisoff, and Jerry M. Linenger, performing pre-launch activities such as eating the traditional breakfast, being suited-up, and riding out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including the countdown, engine ignition, and launch. The film ends with the separation of the Solid Rocket Boosters (SRB) from the shuttle.

Laser-induced breakdown ignition in a gas fed two-stroke engine

NASA Astrophysics Data System (ADS)

Loktionov, E. Y.; Pasechnikov, N. A.; Telekh, V. D.

2018-01-01

Laser-induced ignition for internal combustion engines is investigated intensively after demonstration of a compact ‘laser plug’ possibility. Laser spark benefits as compared to traditional spark plugs are higher compression rate, and possibility of almost any fuel ignition, so lean mixtures burning with lower temperatures could reduce harmful exhausts (NO x , CH, etc). No need in electrode and possibility for multi-point, linear or circular ignition can make combustion even more effective. Laser induced combustion wave appears faster and is more stable in time, than electric one, so can be used for ramjets, chemical thrusters, and gas turbines. To the best of our knowledge, we have performed laser spark ignition of a gas fed two-stroke engine for the first time. Combustion temperature and pressure, exhaust composition, ignition timing were investigated at laser and compared to a regular electric spark ignition in a two-stroke model engine. Presented results show possibility for improvement of two-stroke engines performance, in terms of rotation rate increase and NO x emission reduction. Such compact engines using locally mined fuel could be highly demanded in remote Arctic areas.

Ignition study of a petrol/CNG single cylinder engine

NASA Astrophysics Data System (ADS)

Khan, N.; Saleem, Z.; Mirza, A. A.

2005-11-01

Benefits of laser ignition over the electrical ignition system for Compressed Natural Gas (CNG) engines have fuelled automobile industry and led to an extensive research on basic characteristics to switch over to the emerging technologies. This study was undertaken to determine the electrical and physical characteristics of the electric spark ignition of single cylinder petrol/CNG engine to determine minimum ignition requirements and timeline of ignition events to use in subsequent laser ignition study. This communication briefly reviews the ongoing research activities and reports the results of this experimental study. The premixed petrol and CNG mixtures were tested for variation of current and voltage characteristics of the spark with speed of engine. The current magnitude of discharge circuit was found to vary linearly over a wide range of speed but the stroke to stroke fire time was found to vary nonlinearly. The DC voltage profiles were observed to fluctuate randomly during ignition process and staying constant in rest of the combustion cycle. Fire to fire peaks of current amplitudes fluctuated up to 10% of the peak values at constant speed but increased almost linearly with increase in speed. Technical barriers of laser ignition related to threshold minimum ignition energy, inter-pulse durations and firing sequence are discussed. Present findings provide a basic initiative and background information for designing suitable timeline algorithms for laser ignited leaner direct injected CNG engines.

The starting transient of solid propellant rocket motors with high internal gas velocities. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Peretz, A.; Caveny, L. H.; Kuo, K. K.; Summerfield, M.

1973-01-01

A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

KSC-2013-4198

NASA Image and Video Library

2013-11-20

VAN HORN, Texas – Blue Origin test fires a powerful new hydrogen- and oxygen-fueled American rocket engine at the company's West Texas facility. During the test, the BE-3 engine fired at full power for more than two minutes to simulate a launch, then paused for about four minutes, mimicking a coast through space before it re-ignited for a brief final burn. The last phase of the test covered the work the engine could perform in landing the booster back softly on Earth. Blue Origin, a partner of NASA’s Commercial Crew Program, or CCP, is developing its Orbital Launch Vehicle, which could eventually be used to launch the company's Space Vehicle into orbit to transport crew and cargo to low-Earth orbit. CCP is aiding in the innovation and development of American-led commercial capabilities for crew transportation and rescue services to and from the station and other low-Earth orbit destinations by the end of 2017. For information about CCP, visit www.nasa.gov/commercialcrew. Photo credit: NASA/Lauren Harnett

KSC-2013-4124

NASA Image and Video Library

2013-11-20

VAN HORN, Texas – Blue Origin test fires a powerful new hydrogen- and oxygen-fueled American rocket engine at the company's West Texas facility. During the test, the BE-3 engine fired at full power for more than two minutes to simulate a launch, then paused for about four minutes, mimicking a coast through space before it re-ignited for a brief final burn. The last phase of the test covered the work the engine could perform in landing the booster back softly on Earth. Blue Origin, a partner of NASA’s Commercial Crew Program, or CCP, is developing its Orbital Launch Vehicle, which could eventually be used to launch the company's Space Vehicle into orbit to transport crew and cargo to low-Earth orbit. CCP is aiding in the innovation and development of American-led commercial capabilities for crew transportation and rescue services to and from the station and other low-Earth orbit destinations by the end of 2017. For information about CCP, visit www.nasa.gov/commercialcrew. Photo credit: Blue Origin

Fuel Effects on Ignition and Their Impact on Advanced Combustion Engines (Poster)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Taylor, J.; Li, H.; Neill, S.

The objective of this report is to develop a pathway to use easily measured ignition properties as metrics for characterizing fuels in advanced combustion engine research--correlate IQT{trademark} measured parameters with engine data. In HCCL engines, ignition timing depends on the reaction rates throughout compression stroke: need to understand sensitivity to T, P, and [O{sub 2}]; need to rank fuels based on more than one set of conditions; and need to understand how fuel composition (molecular species) affect ignition properties.

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Zakany, James S.

1996-01-01

A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face based on the heat flux estimates from the temperature measurements.

A Comparative Study of Cycle Variability of Laser Plug Ignition vs Classical Spark Plug Ignition in Combustion Engines

NASA Astrophysics Data System (ADS)

Done, Bogdan

2017-10-01

Over the past 30 years numerous studies and laboratory experiments have researched the use of laser energy to ignite gas and fuel-air mixtures. The actual implementation of this laser application has still to be fully achieved in a commercial automotive application. Laser Plug Ignition as a replacement for Spark Plug Ignition in the internal combustion engines of automotive vehicles, offers several potential benefits such as extending lean burn capability, reducing the cyclic variability between combustion cycles and decreasing the total amount of ignition costs, and implicitly weight and energy requirements. The paper presents preliminary results of cycle variability study carried on a SI Engine equipped with laser Plug Ignition system. Versus classic ignition system, the use of the laser Plug Ignition system assures the reduction of the combustion process variability, reflected in the lower values of the coefficient of variability evaluated for indicated mean effective pressure, maximum pressure, maximum pressure angle and maximum pressure rise rate. The laser plug ignition system was mounted on an experimental spark ignition engine and tested at the regime of 90% load and 2800 rev/min, at dosage of λ=1.1. Compared to conventional spark plug, laser ignition assures the efficiency at lean dosage.

Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

NASA Astrophysics Data System (ADS)

Mattox, J. R.

2017-01-01

I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the release of combustion products, i.e., a "whoosh rocket." My recommendation is that the standard fuel for pedagogical whoosh demonstrations be isopropanol, and the recommended vessel is the 3.8-L high-density polyethylene (HDPE) bottle.

Proceedings of the Workshop on the Chemical Suppression of Rocket Afterburning and of Gun Muzzle Flash

DTIC Science & Technology

1987-03-01

We report here the first results of this gun simulator used in the study of muzzle flash. The test setup used is shown in Figure 18. Pressure ports...experiments. For the first tests , the exploding wires mentioned above ignited the gas mixture. Later, "soft" ignition by means of a single tungsten...wire, placed axially in the chamber, was also tested . The voltage pulse applied across this hot wire is shown in Figure 19. This "soft" ignition

Variable valve timing in a homogenous charge compression ignition engine

DOEpatents

Lawrence, Keith E.; Faletti, James J.; Funke, Steven J.; Maloney, Ronald P.

2004-08-03

The present invention relates generally to the field of homogenous charge compression ignition engines, in which fuel is injected when the cylinder piston is relatively close to the bottom dead center position for its compression stroke. The fuel mixes with air in the cylinder during the compression stroke to create a relatively lean homogeneous mixture that preferably ignites when the piston is relatively close to the top dead center position. However, if the ignition event occurs either earlier or later than desired, lowered performance, engine misfire, or even engine damage, can result. The present invention utilizes internal exhaust gas recirculation and/or compression ratio control to control the timing of ignition events and combustion duration in homogeneous charge compression ignition engines. Thus, at least one electro-hydraulic assist actuator is provided that is capable of mechanically engaging at least one cam actuated intake and/or exhaust valve.

Distributed ignition method and apparatus for a combustion engine

DOEpatents

Willi, Martin L.; Bailey, Brett M.; Fiveland, Scott B.; Gong, Weidong

2006-03-07

A method and apparatus for operating an internal combustion engine is provided. The method comprises the steps of introducing a primary fuel into a main combustion chamber of the engine, introducing a pilot fuel into the main combustion chamber of the engine, determining an operating load of the engine, determining a desired spark plug ignition timing based on the engine operating load, and igniting the primary fuel and pilot fuel with a spark plug at the desired spark plug ignition timing. The method is characterized in that the octane number of the pilot fuel is lower than the octane number of the primary fuel.

Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

NASA Astrophysics Data System (ADS)

Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

2016-10-01

This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

77 FR 40879 - Agency Information Collection Activities; Submission to OMB for Review and Approval; Comment...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-07-11

... Compression Ignition Internal Combustion Engines (Renewal) AGENCY: Environmental Protection Agency (EPA....regulations.gov . Title: NSPS for Stationary Source Compression Ignition Internal Combustion Engines (Renewal... Performance Standards (NSPS) for Stationary Source Compression Ignition Internal Combustion Engines (40 CFR...

Final Rule for Phase 2 Emission Standards for New Nonroad Spark-Ignition Handheld Engines At or Below 19 Kilowatts and Minor Amendments to Emission Requirements Applicable to Small Spark-Ignition Engines and Marine Spark-Ignition Engines

EPA Pesticide Factsheets

Rule summary, rule history, CFR citations and additional resources concerning emissions standards for engines principally used in handheld lawn and garden equipment such as trimmers, leaf blowers, and chainsaws.

High-speed schlieren imaging of rocket exhaust plumes

NASA Astrophysics Data System (ADS)

Coultas-McKenney, Caralyn; Winter, Kyle; Hargather, Michael

2016-11-01

Experiments are conducted to examine the exhaust of a variety of rocket engines. The rocket engines are mounted in a schlieren system to allow high-speed imaging of the engine exhaust during startup, steady state, and shutdown. A variety of rocket engines are explored including a research-scale liquid rocket engine, consumer/amateur solid rocket motors, and water bottle rockets. Comparisons of the exhaust characteristics, thrust and cost for this range of rockets is presented. The variety of nozzle designs, target functions, and propellant type provides unique variations in the schlieren imaging.

Fundamental Studies of Ignition Process in Large Natural Gas Engines Using Laser Spark Ignition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Azer Yalin; Bryan Willson

Past research has shown that laser ignition provides a potential means to reduce emissions and improve engine efficiency of gas-fired engines to meet longer-term DOE ARES (Advanced Reciprocating Engine Systems) targets. Despite the potential advantages of laser ignition, the technology is not seeing practical or commercial use. A major impediment in this regard has been the 'open-path' beam delivery used in much of the past research. This mode of delivery is not considered industrially practical owing to safety factors, as well as susceptibility to vibrations, thermal effects etc. The overall goal of our project has been to develop technologies andmore » approaches for practical laser ignition systems. To this end, we are pursuing fiber optically coupled laser ignition system and multiplexing methods for multiple cylinder engine operation. This report summarizes our progress in this regard. A partial summary of our progress includes: development of a figure of merit to guide fiber selection, identification of hollow-core fibers as a potential means of fiber delivery, demonstration of bench-top sparking through hollow-core fibers, single-cylinder engine operation with fiber delivered laser ignition, demonstration of bench-top multiplexing, dual-cylinder engine operation via multiplexed fiber delivered laser ignition, and sparking with fiber lasers. To the best of our knowledge, each of these accomplishments was a first.« less

Laser ignition - Spark plug development and application in reciprocating engines

NASA Astrophysics Data System (ADS)

Pavel, Nicolaie; Bärwinkel, Mark; Heinz, Peter; Brüggemann, Dieter; Dearden, Geoff; Croitoru, Gabriela; Grigore, Oana Valeria

2018-03-01

Combustion is one of the most dominant energy conversion processes used in all areas of human life, but global concerns over exhaust gas pollution and greenhouse gas emission have stimulated further development of the process. Lean combustion and exhaust gas recirculation are approaches to improve the efficiency and to reduce pollutant emissions; however, such measures impede reliable ignition when applied to conventional ignition systems. Therefore, alternative ignition systems are a focus of scientific research. Amongst others, laser induced ignition seems an attractive method to improve the combustion process. In comparison with conventional ignition by electric spark plugs, laser ignition offers a number of potential benefits. Those most often discussed are: no quenching of the combustion flame kernel; the ability to deliver (laser) energy to any location of interest in the combustion chamber; the possibility of delivering the beam simultaneously to different positions, and the temporal control of ignition. If these advantages can be exploited in practice, the engine efficiency may be improved and reliable operation at lean air-fuel mixtures can be achieved, making feasible savings in fuel consumption and reduction in emission of exhaust gasses. Therefore, laser ignition can enable important new approaches to address global concerns about the environmental impact of continued use of reciprocating engines in vehicles and power plants, with the aim of diminishing pollutant levels in the atmosphere. The technology can also support increased use of electrification in powered transport, through its application to ignition of hybrid (electric-gas) engines, and the efficient combustion of advanced fuels. In this work, we review the progress made over the last years in laser ignition research, in particular that aimed towards realizing laser sources (or laser spark plugs) with dimensions and properties suitable for operating directly on an engine. The main envisaged solutions for positioning of the laser spark plug, i.e. placing it apart from or directly on the engine, are introduced. The path taken from the first solution proposed, to build a compact laser suitable for ignition, to the practical realization of a laser spark plug is described. Results obtained by ignition of automobile test engines, with laser devices that resemble classical spark plugs, are specifically discussed. It is emphasized that technological advances have brought this method of laser ignition close to the application and installation in automobiles powered by gasoline engines. Achievements made in the laser ignition of natural gas engines are outlined, as well as the utilization of laser ignition in other applications. Scientific and technical advances have allowed realization of laser devices with multiple (up to four) beam outputs, but many other important aspects (such as integration, thermal endurance or vibration strength) are still to be solved. Recent results of multi-beam ignition of a single-cylinder engine in a test bench set-up are encouraging and have led to increased research interest in this direction. A fundamental understanding of the processes involved in laser ignition is crucial in order to exploit the technology's full potential. Therefore, several measurement techniques, primarily optical types, used to characterize the laser ignition process are reviewed in this work.

Some aspects of the CI engine modification aimed at operation on LPG with the application of spark ignition

NASA Astrophysics Data System (ADS)

Kaparuk, J.; Luft, S.; Skrzek, T.; Wojtyniak, M.

2016-09-01

A lot of investigation on modification of the compression ignition engine aimed at operation on LPG with the application of spark ignition has been carried out in the Laboratory of Vehicles and Combustion Engines at Kazimierz Pulaski University of Technology and Humanities in Radom. This paper presents results of investigation on establishment of the proper ignition advance angle in the modified engine. Within the framework of this investigation it was assessed the effect of this regulation on basic engine operating parameters, exhaust emission as well as basic combustion parameters.

Laser ignition of engines: a realistic option!

NASA Astrophysics Data System (ADS)

Weinrotter, M.; Srivastava, D. K.; Iskra, K.; Graf, J.; Kopecek, H.; Klausner, J.; Herdin, G.; Wintner, E.

2006-01-01

Due to the demands of the market to increase efficiencies and power densities of gas engines, existing ignition schemes are gradually reaching their limits. These limitations initially triggered the development of laser ignition as an effective alternative, first only for gas engines and now for a much wider range of internal combustion engines revealing a number of immediate advantages like no electrode erosion or flame kernel quenching. Furthermore and most noteworthy, already the very first engine tests about 5 years ago had resulted in a drastic reduction of NO x emissions. Within this broad range investigation, laser plasmas were generated by ns Nd-laser pulses and characterized by emission and Schlieren diagnostic methods. High-pressure chamber experiments with lean hydrogen-methane-air mixtures were successfully performed and allowed the determination of essential parameters like minimum pulse energies at different ignition pressures and temperatures as well as at variable fuel air compositions. Multipoint ignition was studied for different ignition point locations. In this way, relevant parameters were acquired allowing to estimate future laser ignition systems. Finally, a prototype diode-pumped passively Q-switched Nd:YAG laser was tested successfully at a gasoline engine allowing to monitor the essential operation characteristics. It is expected that laser ignition involving such novel solid-state lasers will allow much lower maintenance efforts.

Spark ignition timing control system for internal combustion engine with feature of suppression of jerking during engine acceleration

DOE Office of Scientific and Technical Information (OSTI.GOV)

Tomisawa, N.

1989-07-04

This patent describes a spark ignition timing control system for an internal combustion engine, it comprises: sensor means monitoring preselected parameters for producing a sensor signal; first means for deriving a spark ignition timing on the basis of data contained in the sensor signal; second means for detecting an engine acceleration demand for producing an accelerating condition indicative signal; and third means, responsive to the accelerating condition indicative signal, for modifying the spark ignition timing derived by the first means after expiration of a first predetermined period of time of occurence of the accelerating condition indicative signal, in such amore » manner that the spark ignition timing is advanced and retarded for suppressing cycle-to-cycle fluctuation of engine speed and for smoothly increasing engine speed.« less

Experimental Study of Ballistic-Missile Base Heating with Operating Rocket

NASA Technical Reports Server (NTRS)

Nettle, J. Cary

1958-01-01

A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor. Temperature as high as 1200 F was measured in the base region because of the ignition of burnable rocket gases. combustibles that are dumped into the base by accessories seriously aggravate the base-burning temperature rise.

Research on measurement of aviation magneto ignition strength and balance

NASA Astrophysics Data System (ADS)

Gao, Feng; He, Zhixiang; Zhang, Dingpeng

2017-12-01

Aviation magneto ignition system failure accounted for two-thirds of the total fault aviation piston engine and above. At present the method used for this failure diagnosis is often depended on the visual inspections in the civil aviation maintenance field. Due to human factors, the visual inspections cannot provide ignition intensity value and ignition equilibrium deviation value among the different spark plugs in the different cylinder of aviation piston engine. So air magneto ignition strength and balance testing has become an aviation piston engine maintenance technical problem needed to resolve. In this paper, the ultraviolet sensor with detection wavelength of 185~260nm and driving voltage of 320V DC is used as the core of ultraviolet detection to detect the ignition intensity of Aviation magneto ignition system and the balance deviation of the ignition intensity of each cylinder. The experimental results show that the rotational speed within the range 0 to 3500 RPM test error less than 0.34%, ignition strength analysis and calculation error is less than 0.13%, and measured the visual inspection is hard to distinguish between high voltage wire leakage failure of deviation value of 200 pulse ignition strength balance/Sec. The method to detect aviation piston engine maintenance of magneto ignition system fault has a certain reference value.

Ignition kinetics of boron in primary combustion products of propellant based on its unique characteristics

NASA Astrophysics Data System (ADS)

Ao, Wen; Wang, Yang; Wu, Shixi

2017-07-01

Study on the boron-based primary combustion products can bridge the gap between primary combustion and secondary combustion in solid rocket ramjets. To clarify the initial state and ignition characteristics of boron particles in the after-burning chamber of solid rocket ramjets, the elemental, composition and morphology of the primary combustion products collected under gas generator chamber pressure of 0.2 MPa and 6 MPa were investigated by energy dispersive (EDS), X-ray photoelectron spectroscopy (XPS) and scanning electron microscopy with energy dispersive (SEM-EDS) individually. The ignition times of boron particles among the primary combustion products were determined using a high temperature tube furnace system. The BD model was adopted for numerical verification. The numerical solution procedure of boron ignition model in a real afterburner chamber was modified. The results show that the sum of B, C, O elements in the primary combustion products reaches approximately 90%. The primary combustion products are mainly consisted of B, C, and B2O3. Images of the primary combustion products present highly agglomeration, indicating an oxidation of boron surface. Numerous spherical carbon particles with a diameter around 100 nm are observed in the products. Three features of the boron in the primary combustion products are obtained, compared to virgin boron. First most of the boron lumps are covered by carbon particles on the surface. Second the mean particle size is five times larger than that of virgin boron. Third the overall initial oxide layer covered on boron surface increases its thickness by above 0.1 μm. The ignition time of boron in the primary combustion products reaches 20-30 ms under 1673-1873 K, which is quite different from virgin boron of 4 ms. Numerical calculation results show the key reason leading to such a long ignition time is the variation of the initial oxide layer thickness. In conclusion, the physicochemical properties of boron particles are found to differ with virgin boron after primary combustion process. The accurate evaluation of the initial oxide layer thickness and initial particle radius is a crucial procedure before the numerical calculation of boron ignition kinetics. Results of our study are expected to provide better insight in the simulation of solid rocket ramjets working process.

14 CFR 33.69 - Ignitions system.

Code of Federal Regulations, 2013 CFR

2013-01-01

... STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.69 Ignitions system. Each..., except that only one igniter is required for fuel burning augmentation systems. [Amdt. 33-6, 39 FR 35466... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Ignitions system. 33.69 Section 33.69...

14 CFR 33.69 - Ignitions system.

Code of Federal Regulations, 2014 CFR

2014-01-01

... STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.69 Ignitions system. Each..., except that only one igniter is required for fuel burning augmentation systems. [Amdt. 33-6, 39 FR 35466... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Ignitions system. 33.69 Section 33.69...

14 CFR 33.69 - Ignitions system.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.69 Ignitions system. Each..., except that only one igniter is required for fuel burning augmentation systems. [Amdt. 33-6, 39 FR 35466... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Ignitions system. 33.69 Section 33.69...

New Technology Sparks Smoother Engines and Cleaner Air

NASA Technical Reports Server (NTRS)

2001-01-01

Automotive Resources, Inc. (ARI) has developed a new device for igniting fuel in engines-the SmartPlug.TM SmartPlug is a self-contained ignition system that may be retrofitted to existing spark-ignition and compression-ignition engines. The SmartPlug needs as little as six watts of power for warm-up, and requires no electricity at all when the engine is running. Unlike traditional spark plugs, once the SmartPlug ignites the engine, and the engine heats up, the power supply for the plug is no longer necessary. In the utility industry, SmartPlugs can be used in tractors, portable generators, compressors, and pumps. In addition to general-purpose applications, such as lawn mowers and chainsaws, SmartPlugs can also be used in the recreational, marine, aviation, and automotive industries. Unlike traditional ignition systems, the SmartPlug system requires no distributor, coil points, or moving parts. SmartPlugs are non-fouling, with a faster and cleaner burn than traditional spark plugs. They prevent detonation and are not sensitive to moisture, allowing them to be used on a variety of engines. Other advantages include no electrical noise, no high voltage, exceptionally high altitude capabilities, and better cold-start statistics than those of standard spark ignition systems. Future applications for the SmartPlug are being evaluated by manufacturers in the snowmobile industry.

Flight set 360L001 (STS-26) igniter, post flight

NASA Technical Reports Server (NTRS)

Hale, Elgie J.

1990-01-01

Space Transportation System (STS-26) was launched from KSC pad 39B. Two of the Redesigned Solid Rocket Motors (RSRM) were part of the launch system and are designated RSRM-1A and RSRM-1B. Each of these motors were initiated by an 1U75164 igniter. The post flight condition of the igniters and associated components is documented. The overall performance of the igniter components was excellent. No damage or heat affected areas were noted. The sealing elements of the igniter functioned as expected with no evidence of erosion or blowby. The thermal protection system protected all areas adequately. No excessive erosion was noted. Corrosion was found in the special bolt holes in the igniter chamber. The corrosion will not affect refurbishment of the chamber. Beginning with flight 5 grease has been added to the chamber holes to prevent this erosion.

Calculations of the Performance of a Compression-Ignition Engine-Compressor Turbine Combination I : Performance of a Highly Supercharged Compression-Ignition Engine

NASA Technical Reports Server (NTRS)

Sanders, J. C.; Mendelson, Alexander

1945-01-01

Small high-speed single-cylinder compression-ignition engines were tested to determine their performance characteristics under high supercharging. Calculations were made on the energy available in the exhaust gas of the compression-ignition engines. The maximum power at any given maximum cylinder pressure was obtained when the compression pressure was equal to the maximum cylinder pressure. Constant-pressure combustion was found possible at an engine speed of 2200 rpm. Exhaust pressures and temperatures were determined from an analysis of indicator cards. The analysis showed that, at rich mixtures with the exhaust back pressure equal to the inlet-air pressure, there is excess energy available for driving a turbine over that required for supercharging. The presence of this excess energy indicates that a highly supercharged compression-ignition engine might be desirable as a compressor and combustion chamber for a turbine.

Numerical study of chemically reacting viscous flow relevant to pulsed detonation engines

NASA Astrophysics Data System (ADS)

Yi, Tae-Hyeong

2005-11-01

A computational fluid dynamics code for two-dimensional, multi-species, laminar Navier-Stokes equations is developed to simulate a recently proposed engine concept for a pulsed detonation based propulsion system and to investigate the feasibility of the engine of the concept. The governing equations that include transport phenomena such as viscosity, thermal conduction and diffusion are coupled with chemical reactions. The gas is assumed to be thermally perfect and in chemically non-equilibrium. The stiffness due to coupling the fluid dynamics and the chemical kinetics is properly taken care of by using a time-operator splitting method and a variable coefficient ordinary differential equation solver. A second-order Roe scheme with a minmod limiter is explicitly used for space descretization, while a second-order, two-step Runge-Kutta method is used for time descretization. In space integration, a finite volume method and a cell-centered scheme are employed. The first-order derivatives in the equations of transport properties are discretized by a central differencing with Green's theorem. Detailed chemistry is involved in this study. Two chemical reaction mechanisms are extracted from GRI-Mech, which are forty elementary reactions with thirteen species for a hydrogen-air mixture and twenty-seven reactions with eight species for a hydrogen-oxygen mixture. The code is ported to a high-performance parallel machine with Message-Passing Interface. Code validation is performed with chemical kinetic modeling for a stoichiometric hydrogen-air mixture, an one-dimensional detonation tube, a two-dimensional, inviscid flow over a wedge and a viscous flow over a flat plate. Detonation is initiated using a numerically simulated arc-ignition or shock-induced ignition system. Various freestream conditions are utilized to study the propagation of the detonation in the proposed concept of the engine. Investigation of the detonation propagation is performed for a pulsed detonation rocket and a supersonic combustion chamber. For a pulsed detonation rocket case, the detonation tube is embedded in a mixing chamber where an initiator is added to the main detonation chamber. Propagating detonation waves in a supersonic combustion chamber is investigated for one- and two-dimensional cases. The detonation initiated by an arc and a shock wave is studied in the inviscid and viscous flow, respectively. Various features including a detonation-shock interaction, a detonation diffraction, a base flow and a vortex are observed.

Future space transport

NASA Technical Reports Server (NTRS)

Grishin, S. D.; Chekalin, S. V.

1984-01-01

Prospects for the mastery of space and the basic problems which must be solved in developing systems for both manned and cargo spacecraft are examined. The achievements and flaws of rocket boosters are discussed as well as the use of reusable spacecraft. The need for orbiting satellite solar power plants and related astrionics for active control of large space structures for space stations and colonies in an age of space industrialization is demonstrated. Various forms of spacecraft propulsion are described including liquid propellant rocket engines, nuclear reactors, thermonuclear rocket engines, electrorocket engines, electromagnetic engines, magnetic gas dynamic generators, electromagnetic mass accelerators (rail guns), laser rocket engines, pulse nuclear rocket engines, ramjet thermonuclear rocket engines, and photon rockets. The possibilities of interstellar flight are assessed.

Thinshell symmetry surrogates for the National Ignition Facility: A rocket equation analysis

NASA Astrophysics Data System (ADS)

Amendt, Peter; Shestakov, A. I.; Landen, O. L.; Bradley, D. K.; Pollaine, S. M.; Suter, L. J.; Turner, R. E.

2001-06-01

Several techniques for inferring the degree of flux symmetry in indirectly driven cylindrical hohlraums have been developed over the past several years for eventual application to the National Ignition Facility (NIF) [Paisner et al., Laser Focus World 30, 75 (1994)]. These methods use various ignition capsule surrogates, including non-cryogenic imploded capsules [Hauer et al., Phys. Plasmas 2, 2488 (1995)], backlit aerogel foamballs [Amendt et al., Rev. Sci. Instrum. 66, 785 (1995)], reemission balls [Delamater, Magelssen, and Hauer, Phys. Rev. E 53, 5240 (1996)], and backlit thinshells [Pollaine et al., Phys. Plasmas 8, 2357 (2001)]. Recent attention has focussed on the backlit thinshells as a promising means for detecting higher-order Legendre flux asymmetries, e.g., P6 and P8, which are predicted to be important sources of target performance degradation on the NIF for levels greater than 1% [Haan et al., Phys. Plasmas 2, 2490 (1995)]. A key property of backlit thinshells is the strong amplification of modal flux asymmetry imprinting with shell convergence. A simple single-parameter analytic description based on a rocket model is presented which explores the degree of linearity of the shell response to an imposed flux asymmetry. Convergence and mass ablation effects introduce a modest level of nonlinearity in the shell response. The effect of target fabrication irregularities on shell distortion is assessed with the rocket model and particular sensitivity to shell thickness variations is shown. The model can be used to relate an observed or simulated backlit implosion trajectory to an ablation pressure asymmetry history. Ascertaining this history is an important element for readily establishing the degree of surrogacy of a symmetry target for a NIF ignition capsule.

78 FR 48826 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-12

... could cause a fuel leak near an ignition source (e.g., hot brakes or engine exhaust nozzle... could cause a fuel leak near an ignition source (e.g., hot brakes or engine nozzle), consequently... ignition source (e.g., hot brakes or engine nozzle), consequently leading to a fuel-fed fire. (f...

Analysis of liquid-propellant rocket engines designed by F. A. Tsander

NASA Technical Reports Server (NTRS)

Dushkin, L. S.; Moshkin, Y. K.

1977-01-01

The development of the oxygen-gasoline OR-2 engines and the oxygen-alcohol GIRD-10 rocket engine is described. A result of Tsander's rocket research was an engineering method for propellant calculation of oxygen-propellant rocket engines that determined the basic parameters of the engine and the structural elements.

A Study of Flame Physics and Solid Propellant Rocket Physics

DTIC Science & Technology

2007-10-01

and ellipsoids, and the packing of pellets relevant to igniter modeling. Other topics are the instabilities of smolder waves, premixed flame...instabilities in narrow tubes, and flames supported by a spinning porous plug burner . Much of this work has been reported in the high-quality archival...perchlorate in fuel binder, the combustion of model propellant packs of ellipses and ellipsoids, and the packing of pellets relevant to igniter modeling

Ares I-X Post Flight Ignition Overpressure Review

NASA Technical Reports Server (NTRS)

Alvord, David A.

2010-01-01

Ignition Overpressure (IOP) is an unsteady fluid flow and acoustic phenomena caused by the rapid expansion of gas from the rocket nozzle within a ducted launching space resulting in an initially higher amplitude pressure wave. This wave is potentially dangerous to the structural integrity of the vehicle. An in-depth look at the IOP environments resulting from the Ares I-X Solid Rocket Booster configuration showed high correlation between the pre-flight predictions and post-flight analysis results. Correlation between the chamber pressure and IOP transients showed successful acoustic mitigation, containing the strongest IOP waves below the Mobile Launch Pad deck. The flight data allowed subsequent verification and validation of Ares I-X unsteady fluid ducted launcher predictions, computational fluid dynamic models, and strong correlation with historical Shuttle data.

Development Status of Reusable Rocket Engine

NASA Astrophysics Data System (ADS)

Yoshida, Makoto; Takada, Satoshi; Naruo, Yoshihiro; Niu, Kenichi

A 30-kN rocket engine, a pilot engine, is being developed in Japan. Development of this pilot engine has been initiated in relation to a reusable sounding rocket, which is also being developed in Japan. This rocket takes off vertically, reaches an altitude of 100 km, lands vertically at the launch site, and is launched again within several days. Due to advantage of reusability, successful development of this rocket will mean that observation missions can be carried out more frequently and economically. In order to realize this rocket concept, the engines installed on the rocket should be characterized by reusability, long life, deep throttling and health monitoring, features which have not yet been established in Japanese rocket engines. To solve the engineering factors entitled by those features, a new design methodology, advanced engine simulations and engineering testing are being focused on in the pilot engine development stage. Especially in engineering testing, limit condition data is acquired to facilitate development of new diagnostic techniques, which can be applied by utilizing the mobility of small-size hardware. In this paper, the development status of the pilot engine is described, including fundamental design and engineering tests of the turbopump bearing and seal, turbine rig, injector and combustion chamber, and operation and maintenance concepts for one hundred flights by a reusable rocket are examined.

Hybrid boosters for future launch vehicles

NASA Astrophysics Data System (ADS)

Dargies, E.; Lo, R. E.

1987-10-01

Hybrid rocket propulsion systems furnish the advantages of much higher safety levels, due both to shut-down capability in case of ignition failure to one unit and the potential choice of nontoxic propellant combinations, such as LOX/polyethylene; they nevertheless yield performance levels comparable or superior to those of solid rocket boosters. Attention is presently given to the results of DFVLR analytical model studies of hybrid propulsion systems, with attention to solid fuel grain geometrical design and propellant grain surface ablation rate. The safety of hybrid rockets recommends them for use by manned spacecraft.

Miniature Rocket Motor for Aircraft Stall/Spin Recovery

NASA Technical Reports Server (NTRS)

Lucy, M. H.

1985-01-01

Design accommodates different thrust levels and burn times with minimum weight. Different thrust levels achieved by substituting other propellants of different diameter and burn-rate characteristics. Different burn times achieved by simply changing length of grain/tube assembly. Grain bond material also acts as insulator for fiberglass tube. Rocket motor attached to aircraft model and ignited from radio-controlled 4.8-volt power source. Device provides more than twice energy available in previous designs at only 60 percent of weight. Rocket motor used to identify energy requirements for aircraft stall/spin recovery positive propulsion system.

Aqueous Ethanol Ignition and Engine Studies, Phase I

DOT National Transportation Integrated Search

2010-09-01

Our objectives were to design a micro-dilution tunnel for monitoring engine emissions, measure ignition temperature and heat release from ethanol-water-air mixtures on platinum, and initiate a computational fluid dynamics model of a catalytic igniter...

Overview of rocket engine control

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.; Musgrave, Jeffrey L.

1991-01-01

The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.

Autoignition Chemistry of Surrogate Fuel Components in an Engine Environment

DTIC Science & Technology

2015-08-21

compression ratio (CR) on the auto - ignition of decane. Crank angle resolved cylinder pressure data was acquired and analyzed using an engine heat...schematic shown in Fig. 1, consists of a modified CFR (Cooperative Fuel Research) engine coupled to a dynamometer. In practical compression 2 ignition ...engines, auto - ignition occurs in the premixed spray envelope that forms during the fuel injection process. To focus on this regime without the

KSC-07pd2592

NASA Image and Video Library

2007-09-27

KENNEDY SPACE CENTER, FLA. -- Nearly enveloped by the smoke after ignition, the Delta II rocket carrying NASA's Dawn spacecraft rises from the smoke and fire on the launch pad to begin its 1.7-billion-mile journey through the inner solar system to study a pair of asteroids. Liftoff was at 7:34 a.m. EDT from Pad 17-B at Cape Canaveral Air Force Station. Dawn is the ninth mission in NASA's Discovery Program. The spacecraft will be the first to orbit two planetary bodies, asteroid Vesta and dwarf planet Ceres, during a single mission. Vesta and Ceres lie in the asteroid belt between Mars and Jupiter. It is also NASA's first purely scientific mission powered by three solar electric ion propulsion engines. Photo credit: NASA/Sandra Joseph & Rafael Hernandez

STS-82 Flight Day 01 Highlights

NASA Technical Reports Server (NTRS)

1997-01-01

The first day of the STS-82 mission begins with the crew, Commander Kenneth D. Bowersox, Pilot Scott J. Horowitz, Payload Commander Mark C. Lee, and Mission Specialists Gregory J. Harbaugh, Steven L. Smith, Joseph R. Tanner, and Steven A. Hawley performing pre-launch activities such as eating the traditional breakfast, being suited up, and riding out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch, and arm retraction, launch activities are shown including the countdown, engine ignition, launch, shuttle roll maneuver, and then the separation of the Solid Rocket Boosters (SRB) from the shuttle. Once in orbit the cargo bay doors are seen opening.

STS-99 Crew Activities Report/Flight Day 1 Highlights

NASA Technical Reports Server (NTRS)

2000-01-01

Live footage shows the crew, Commander Kevin R. Kregel, Pilot Dominic L. Pudwill Gorie, and Mission Specialists Janet L. Kavandi, Janice E. Voss, Mamoru Mohri and Gerhard P.J. Thiele, seated in the dining room with the traditional cake. The crew is seen performing various pre-launch activities including suit-up, walk out to the Astro-van, and strap-in into the vehicle. Also seen are the retractions of the orbiter access arm and the gaseous oxygen mint hood, main engine start, booster ignition, liftoff, and separation of the solid rocket boosters. The Red Team (first of the dual shift crew) includes Kregel, Kavandi, and Thiele, who are shown conducting jet thruster firings, activating radar instruments, and deploying the boom (mass).

75 FR 32611 - Standards of Performance for Stationary Compression Ignition and Spark Ignition Internal...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-06-08

... implement more stringent standards for stationary compression ignition engines with displacement greater... engines with displacement at or above 30 liters per cylinder to align more closely with recent standards.... Standards for New Engines With Displacement Greater Than or Equal to 10 l/cyl and Less Than 30 l/cyl B...

Design and Testing of a Liquid Nitrous Oxide and Ethanol Fueled Rocket Engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Youngblood, Stewart

A small-scale, bi-propellant, liquid fueled rocket engine and supporting test infrastructure were designed and constructed at the Energetic Materials Research and Testing Center (EMRTC). This facility was used to evaluate liquid nitrous oxide and ethanol as potential rocket propellants. Thrust and pressure measurements along with high-speed digital imaging of the rocket exhaust plume were made. This experimental data was used for validation of a computational model developed of the rocket engine tested. The developed computational model was utilized to analyze rocket engine performance across a range of operating pressures, fuel-oxidizer mixture ratios, and outlet nozzle configurations. A comparative study ofmore » the modeling of a liquid rocket engine was performed using NASA CEA and Cantera, an opensource equilibrium code capable of being interfaced with MATLAB. One goal of this modeling was to demonstrate the ability of Cantera to accurately model the basic chemical equilibrium, thermodynamics, and transport properties for varied fuel and oxidizer operating conditions. Once validated for basic equilibrium, an expanded MATLAB code, referencing Cantera, was advanced beyond CEAs capabilities to predict rocket engine performance as a function of supplied propellant flow rate and rocket engine nozzle dimensions. Cantera was found to comparable favorably to CEA for making equilibrium calculations, supporting its use as an alternative to CEA. The developed rocket engine performs as predicted, demonstrating the developedMATLAB rocket engine model was successful in predicting real world rocket engine performance. Finally, nitrous oxide and ethanol were shown to perform well as rocket propellants, with specific impulses experimentally recorded in the range of 250 to 260 seconds.« less

Detection of combustion start in the controlled auto ignition engine by wavelet transform of the engine block vibration signal

NASA Astrophysics Data System (ADS)

Kim, Seonguk; Min, Kyoungdoug

2008-08-01

The CAI (controlled auto ignition) engine ignites fuel and air mixture by trapping high temperature burnt gas using a negative valve overlap. Due to auto ignition in CAI combustion, efficiency improvements and low level NOx emission can be obtained. Meanwhile, the CAI combustion regime is restricted and control parameters are limited. The start of combustion data in the compressed ignition engine are most critical for controlling the overall combustion. In this research, the engine block vibration signal is transformed by the Meyer wavelet to analyze CAI combustion more easily and accurately. Signal acquisition of the engine block vibration is a more suitable method for practical use than measurement of in-cylinder pressure. A new method for detecting combustion start in CAI engines through wavelet transformation of the engine block vibration signal was developed and results indicate that it is accurate enough to analyze the start of combustion. Experimental results show that wavelet transformation of engine block vibration can track the start of combustion in each cycle. From this newly developed method, the start of combustion data in CAI engines can be detected more easily and used as input data for controlling CAI combustion.

Three-dimensional numerical and experimental studies on transient ignition of hybrid rocket motor

NASA Astrophysics Data System (ADS)

Tian, Hui; Yu, Ruipeng; Zhu, Hao; Wu, Junfeng; Cai, Guobiao

2017-11-01

This paper presents transient simulations and experimental studies of the ignition process of the hybrid rocket motors (HRMs) using 90% hydrogen peroxide (HP) as the oxidizer and polymethyl methacrylate (PMMA) and Polyethylene (PE) as fuels. A fluid-solid coupling numerically method is established based on the conserved form of the three-dimensional unsteady Navier-Stokes (N-S) equations, considering gas fluid with chemical reactions and heat transfer between the fluid and solid region. Experiments are subsequently conducted using high-speed camera to record the ignition process. The flame propagation, chamber pressurizing process and average fuel regression rate of the numerical simulation results show good agreement with the experimental ones, which demonstrates the validity of the simulations in this study. The results also indicate that the flame propagation time is mainly affected by fluid dynamics and it increases with an increasing grain port area. The chamber pressurizing process begins when the flame propagation completes in the grain port. Furthermore, the chamber pressurizing time is about 4 times longer than the time of flame propagation.

Simulation of Acoustics for Ares I Scale Model Acoustic Tests

NASA Technical Reports Server (NTRS)

Putnam, Gabriel; Strutzenberg, Louise L.

2011-01-01

The Ares I Scale Model Acoustics Test (ASMAT) is a series of live-fire tests of scaled rocket motors meant to simulate the conditions of the Ares I launch configuration. These tests have provided a well documented set of high fidelity acoustic measurements useful for validation including data taken over a range of test conditions and containing phenomena like Ignition Over-Pressure and water suppression of acoustics. To take advantage of this data, a digital representation of the ASMAT test setup has been constructed and test firings of the motor have been simulated using the Loci/CHEM computational fluid dynamics software. Results from ASMAT simulations with the rocket in both held down and elevated configurations, as well as with and without water suppression have been compared to acoustic data collected from similar live-fire tests. Results of acoustic comparisons have shown good correlation with the amplitude and temporal shape of pressure features and reasonable spectral accuracy up to approximately 1000 Hz. Major plume and acoustic features have been well captured including the plume shock structure, the igniter pulse transient, and the ignition overpressure.

Hot Water Propulsion for Horizontal Rocket Assisted Take-Off Systems for Future Reusable Launch Vehicles

NASA Astrophysics Data System (ADS)

Pilz, N.; Adirim, H.; Lo, R.; Schildknecht, A.

2004-10-01

Among other concepts, reusable space transportation systems that comprise winged reusable launch vehicles (RLV) with horizontal take-off and horizontal landing (HTHL) are under worldwide investigation, e.g. the respective concepts within ESA's FESTIP-Study (Future European Space Transportation Integration Program) or the HOPPER concept by EADS-ST. The payload of these RLVs could be significantly increased by means of a ground-based take-off assistance system that would accelerate the vehicle along a horizontal track until it reaches the desired speed to ignite its onboard engines for leaving the ground and launching into orbit. This paper illustrates the advantages of horizontal take-off for winged RLVs and provides an overview of launch-assist options for HTHL RLVs. It presents hot water propulsion for ground-based take-off assistance systems for future RLVs as an attractive choice besides magnetic levitation and acceleration (maglev) technology. Finally, preliminary design concepts are presented for a rocket assisted take-off system (RATOS) with hot water propulsion followed by an analysis of its improvement potential.

Powdered Magnesium-Carbon Dioxide Rocket Combustion Technology for In Situ Mars Propulsion

NASA Technical Reports Server (NTRS)

Foote, J. P.; Litchford, R. J.

2007-01-01

Powdered magnesium (Mg) carbon dioxide (CO2) combustion is examined as a potential in situ propellant combination for Mars propulsion. Although this particular combination has relatively low performance in comparison to traditional bipropellants, it remains attractive as a potential basis for future martian mobility systems, since it could be partially or wholly manufactured from indigenous planetary resources. As a means of achieving high mobility during long-duration Mars exploration missions, the poorer performing in situ combination can, in fact, become a superior alternative to conventional storable propellants, which would need to be entirely transported from Earth. Thus, the engineering aspects of powdered metal combustion devices are discussed including transport/injection of compacted powder, ignition, combustion efficiency, combustion stability, dilution effects, lean burn limits, and slag formation issues. It is suggested that these technological issues could be effectively addressed through a multiphase research and development effort beginning with basic feasibility tests using an existing dump configured atmospheric pressure burner. Follow-on phases would involve the development and testing of a pressurized research combustor and technology demonstration tests of a prototypical rocket configuration.

STS-96 Post Flight Presentation

NASA Technical Reports Server (NTRS)

1999-01-01

The Crew of STS-96 Discovery Shuttle, Commander Kent V. Rominger, Pilot Rick D. Husband, Mission Specialists Ellen Ochoa, Tamara E. Jernigan, Daniel T. Barry, Julie Payette, and Valery Ivanovich Tokarev, are shown narrating the mission highlights. Scenes include walk out to the transfer vehicle, and launch of the shuttle. Also presented are scenes of the start of the main engine, ignition of the solid rocket boosters, and the separation of the solid rocket boosters. Footage of Payette preparing the on-board camera equipment, while Barry and Jernigan perform routine checks of the equipment is seen. Also presented are various pictures of the shuttle in its orbit, the docking of the shuttle with the Mir International Space Station, and crewmembers during their space walk. Beautiful panoramic views of the Great Lake, Houston, and a combined view of Italy and Turkey are seen. The crew of Discovery is shown performing a juice ball experiment, tumbling, undocking, performing transfer operations, and deploying the STARSHINE educational satellite. The film ends with the reentry of the Discovery Space Shuttle into the Earth's atmosphere.

Developing hybrid near-space technologies for affordable access to suborbital space

NASA Astrophysics Data System (ADS)

Badders, Brian David

High power rockets and high altitude balloons are two near-space technologies that could be combined in order to provide access to the mesosphere and, eventually, suborbital space. This "rockoon" technology has been used by several large budget space programs before being abandoned in favor of even more expensive, albeit more accurate, ground launch systems. With the increased development of nano-satellites and atmospheric sensors, combined with rising interest in global atmospheric data, there is an increase in desire for affordable access to extreme altitudes that does not necessarily require the precision of ground launches. Development of hybrid near-space technologies for access to over 200k ft. on a small budget brings many challenges within engineering, systems integration, cost analysis, market analysis, and business planning. This research includes the design and simulation testing of all the systems needed for a safe and reusable launch system, the cost analysis for initial production, the development of a business plan, and the development of a marketing plan. This project has both engineering and scientific significance in that it can prove the space readiness of new technologies, raise their technology readiness levels (TRLs), expedite the development process, and also provide new data to the scientific community. It also has the ability to stimulate university involvement in the aerospace industry and help to inspire the next generation of workers in the space sector. Previous development of high altitude balloon/high power rocket hybrid systems have been undertaken by government funded military programs or large aerospace corporations with varying degrees of success. However, there has yet to be a successful flight with this type of system which provides access to the upper mesosphere in a university setting. This project will aim to design and analyze a viable system while testing the engineering process under challenging budgetary constraints. The technical, engineering, and systems integration challenges that will be investigated are rocket design, launch platform design, communications, ignition systems, recovery systems, and stabilization methods. This will be done using rocket performance simulation software, computer-aided design software, and computational fluid dynamic analysis software. The business planning is also an important part of this research. Through detailed market analysis, the needs for the proposed product/services being developed will be assessed. Through the combination of detailed cost analysis and the market needs, the economic viability of this launch system will be determined.

Development and Testing of a Green Monopropellant Ignition System

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Merkley, Daniel P.; Eilers, Shannon D.; Judson, Michael I.; Taylor, Terry L.

2013-01-01

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.

Space Shuttle SRM Ignition System. [Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Bolieau, C. W.; Baker, J. S.; Folkman, S. L.

1978-01-01

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.

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

NASA Technical Reports Server (NTRS)

Kleinhenz, Julie; Sarmiento, Charles; Marshall, William

2012-01-01

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.

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

NASA Technical Reports Server (NTRS)

Kleinhenz, Julie; Sarmiento, Charles; Marshall, William

2012-01-01

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.

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Astrophysics Data System (ADS)

Bickford, R. L.; Madzsar, G.

1990-07-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Technical Reports Server (NTRS)

Bickford, R. L.; Madzsar, G.

1990-01-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

Potential of Spark Ignition Engine : Engine Design Concepts

DOT National Transportation Integrated Search

1980-03-01

This report provides a review and assessment of potential improvements in fuel economy for a selected number of spark ignition engine design technologies for passenger cars and light trucks. The engine design technologies examined include: : a) optim...

Alternative Automobile Engines

ERIC Educational Resources Information Center

Wilson, David Gordon

1978-01-01

Requirements for cleaner and more efficient engines have stimulated a search for alternatives to the conventional spark-ignition engine. So far, the defects of the alternative engines are clearer than the virtues. The following engines are compared: spark ignition, diesel, vapor-cycle, Stirling, and gas turbine. (Author/MA)

14 CFR 27.1145 - Ignition switches.

Code of Federal Regulations, 2014 CFR

2014-01-01

... master ignition control. (b) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 27.1145 Ignition...

14 CFR 27.1145 - Ignition switches.

Code of Federal Regulations, 2012 CFR

2012-01-01

... master ignition control. (b) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 27.1145 Ignition...

14 CFR 27.1145 - Ignition switches.

Code of Federal Regulations, 2010 CFR

2010-01-01

... master ignition control. (b) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 27.1145 Ignition...

14 CFR 27.1145 - Ignition switches.

Code of Federal Regulations, 2011 CFR

2011-01-01

... master ignition control. (b) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 27.1145 Ignition...

14 CFR 27.1145 - Ignition switches.

Code of Federal Regulations, 2013 CFR

2013-01-01

... master ignition control. (b) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 27.1145 Ignition...

Prechamber equipped laser ignition for improved performance in natural gas engines

DOE PAGES

Almansour, Bader; Vasu, Subith; Gupta, Sreenath B.; ...

2017-04-25

Lean-burn operation of stationary natural gas engines offers lower NO x emissions and improved efficiency. A proven pathway to extend lean-burn operation has been to use laser ignition instead of standard spark ignition. However, under lean conditions, flame speed reduces thereby offsetting any efficiency gains resulting from the higher ratio of specific heats, γ. The reduced flame speeds, in turn, can be compensated with the use of a prechamber to result in volumetric ignition, and thereby lead to faster combustion. In this study, the optimal geometry of PCLI was identified through several tests in a single-cylinder engine as a compromisemore » between autoignition, NO x and soot formation within the prechamber. Subsequently, tests were conducted in a single-cylinder natural gas engine comparing the performance of three ignition systems: standard electrical spark ignition (SI), single-point laser ignition (LI), and prechamber equipped laser ignition (PCLI). Out of the three, the performance of PCLI was far superior compared to the other two. Efficiency gain of 2.1% points could be achieved while complying with EPA regulation (BSNO x < 1.34 kW-hr) and the industry standard for ignition stability (COV_IMEP < 5%). Finally, test results and data analysis are presented identifying the combustion mechanisms leading to the improved performance.« less

Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

NASA Technical Reports Server (NTRS)

Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

1999-01-01

In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

Rocket propulsion by thermonuclear micro-bombs ignited with intense relativistic electron beams.

NASA Technical Reports Server (NTRS)

Winterberg, F.

1971-01-01

Discussion of a method for the ignition of a thermonuclear microbomb by means of an intense relativistic electron beam with regard to its potential application to rocket propulsion. With such a system, exhaust velocities up to 1000 km/sec, corresponding to a specific impulse of 100,000 sec, seem to be within the realm of possibility. The rocket is propelled by a chain of thermonuclear microbombs exploded in a concave magnetic mirror produced by superconducting field coils. The magnetic pressure of the field reflects the fireball generated by the explosion. For the large capacitor bank required to generate the intense relativistic electron beam, a desirable lightweight design may be possible through use of ferroelectric materials. Because of the high cost of the T-D and He 3-D thermonuclear material, the system has to be optimized by minimizing the T-D and He 3-D consumption by a proper TD and He 3-D fuel to hydrogen propellant mass ratio, leading to a larger total system mass than would be absolutely necessary.

Grease-Resistant O Rings for Joints in Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Harvey, Albert R.; Feldman, Harold

2003-01-01

There is a continuing effort to develop improved O rings for sealing joints in solid-fuel rocket motors. Following an approach based on the lessons learned in the explosion of the space shuttle Challenger, investigators have been seeking O-ring materials that exhibit adequate resilience for effective sealing over a broad temperature range: What are desired are O rings that expand far and fast enough to maintain seals, even when metal sealing surfaces at a joint move slightly away from each other shortly after ignition and the motor was exposed to cold weather before ignition. Other qualities desired of the improved O rings include adequate resistance to ablation by hot rocket gases and resistance to swelling when exposed to hydrocarbon-based greases used to protect some motor components against corrosion. Five rubber formulations two based on a fluorosilicone polymer and three based on copolymers of epichlorohydrin with ethylene oxide were tested as candidate O-ring materials. Of these, one of the epichlorohydrin/ethylene oxide formulations was found to offer the closest to the desired combination of properties and was selected for further evaluation.

Study of solid rocket motor for space shuttle booster, volume 2, book 1

NASA Technical Reports Server (NTRS)

1972-01-01

The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

14 CFR 33.37 - Ignition system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Ignition system. 33.37 Section 33.37 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Reciprocating Aircraft Engines § 33.37 Ignition system...

Efficiency of the rocket engines with a supersonic afterburner

NASA Astrophysics Data System (ADS)

Sergienko, A. A.

1992-08-01

The paper is concerned with the problem of regenerative cooling of the liquid-propellant rocket engine combustion chamber at high pressures of the working fluid. It is shown that high combustion product pressures can be achieved in the liquid-propellant rocket engine with a supersonic afterburner than in a liquid-propellant rocket engine with a conventional subsonic combustion chamber for the same allowable heat flux density. However, the liquid-propellant rocket engine with a supersonic afterburner becomes more economical than the conventional engine only at generator gas temperatures of 1700 K and higher.

Reactivity-controlled compression ignition drive cycle emissions and fuel economy estimations using vehicle system simulations

DOE PAGES

Curran, Scott J.; Gao, Zhiming; Wagner, Robert M.

2014-12-22

In-cylinder blending of gasoline and diesel to achieve reactivity-controlled compression ignition has been shown to reduce NO X and soot emissions while maintaining or improving brake thermal efficiency as compared with conventional diesel combustion. The reactivity-controlled compression ignition concept has an advantage over many advanced combustion strategies in that the fuel reactivity can be tailored to the engine speed and load, allowing stable low-temperature combustion to be extended over more of the light-duty drive cycle load range. In this paper, a multi-mode reactivity-controlled compression ignition strategy is employed where the engine switches from reactivity-controlled compression ignition to conventional diesel combustionmore » when speed and load demand are outside of the experimentally determined reactivity-controlled compression ignition range. The potential for reactivity-controlled compression ignition to reduce drive cycle fuel economy and emissions is not clearly understood and is explored here by simulating the fuel economy and emissions for a multi-mode reactivity-controlled compression ignition–enabled vehicle operating over a variety of US drive cycles using experimental engine maps for multi-mode reactivity-controlled compression ignition, conventional diesel combustion, and a 2009 port-fuel injected gasoline engine. Drive cycle simulations are completed assuming a conventional mid-size passenger vehicle with an automatic transmission. Multi-mode reactivity-controlled compression ignition fuel economy simulation results are compared with the same vehicle powered by a representative 2009 port-fuel injected gasoline engine over multiple drive cycles. Finally, engine-out drive cycle emissions are compared with conventional diesel combustion, and observations regarding relative gasoline and diesel tank sizes needed for the various drive cycles are also summarized.« less

Metallized Gelled Propellants: Oxygen/RP-1/aluminum Rocket Combustion Experiments

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Zakany, James S.

1995-01-01

A series of combustion experiments were conducted to measure the specific impulse, Cstar-, and specific-impulse efficiencies of a rocket engine using metallized gelled liquid propellants. These experiments used a small 20- to 40-1bf (89- to 178-N) thrust, modular engine consisting of an injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum and gaseous oxygen was the oxidizer. Ten different injectors were used during the testing: 6 for the baseline 02/RP-1 tests and 4 for the gelled fuel tests which covered a wide range of mixture ratios. At the peak of the Isp versus oxidizer-to-fuel ratio (O/F) data, a range of 93 to 99% Cstar efficiency was reached with ungelled 02/RP-1. A Cstar efficiency range of 75 to 99% was obtained with gelled RP-l (0-wt% RP-1/Al) while the metallized 5-wt% RP-1/Al delivered a Cstar efficiency of 94 to 99% at the peak Isp in the O/F range tested. An 88 to 99% Cstar efficiency was obtained at the peak Isp of the gelled RP1/Al with 55-wt% Al. Specific impulse efficiencies for the 55-wt% RP-1/Al of 67%-83% were obtained at a 2.4:1 expansion ratio. Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0- and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling. These experiments represent a first step in characterizing the performance of and operational issues with gelled RP-1 fuels.

KSC-04PD-1080

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. -- A remote camera captures ignition of the Delta II rocket carrying the Gravity Probe B spacecraft from Space Launch Complex 2 on Vandenberg AFB, Calif., at 9:57:24 a.m. PDT.

KSC-2013-4164

NASA Image and Video Library

2013-11-20

VAN HORN, Texas – Blue Origin’s test stand, back right, is framed by a wind mill at the company’s West Texas facility. The company used this test stand to fire its powerful new hydrogen- and oxygen-fueled American rocket engine, the BE-3. The engine fired at full power for more than two minutes to simulate a launch, then paused for about four minutes, mimicking a coast through space before it re-ignited for a brief final burn. The last phase of the test covered the work the engine could perform in landing the booster back softly on Earth. Blue Origin, a partner of NASA’s Commercial Crew Program, or CCP, is developing its Orbital Launch Vehicle, which could eventually be used to launch the company's Space Vehicle into orbit to transport crew and cargo to low-Earth orbit. CCP is aiding in the innovation and development of American-led commercial capabilities for crew transportation and rescue services to and from the station and other low-Earth orbit destinations by the end of 2017. For information about CCP, visit www.nasa.gov/commercialcrew. Photo credit: NASA/Lauren Harnett

Virtual engine management simulator for educational purposes

NASA Astrophysics Data System (ADS)

Drosescu, R.

2017-10-01

This simulator was conceived as a software program capable of generating complex control signals, identical to those in the electronic management systems of modern spark ignition or diesel engines. Speed in rpm and engine load percentage defined by throttle opening angle represent the input variables in the simulation program and are graphically entered by two-meter instruments from the simulator central block diagram. The output signals are divided into four categories: synchronization and position of each cylinder, spark pulses for spark ignition engines, injection pulses and, signals for generating the knock window for each cylinder in the case of a spark ignition engine. The simulation program runs in real-time so each signal evolution reflects the real behavior on a physically thermal engine. In this way, the generated signals (ignition or injection pulses) can be used with additionally drivers to control an engine on the test bench.

Hydrogen as an Auxiliary Fuel in Compression-Ignition Engines

NASA Technical Reports Server (NTRS)

Gerrish, Harold C; Foster, H

1936-01-01

An investigation was made to determine whether a sufficient amount of hydrogen could be efficiently burned in a compression-ignition engine to compensate for the increase of lift of an airship due to the consumption of the fuel oil. The performance of a single-cylinder four-stroke-cycle compression-ignition engine operating on fuel oil alone was compared with its performance when various quantities of hydrogen were inducted with the inlet air. Engine-performance data, indicator cards, and exhaust-gas samples were obtained for each change in engine-operating conditions.

An experimental and modeling study investigating the ignition delay in a military diesel engine running hexadecane (cetane) fuel

DOE PAGES

Cowart, Jim S.; Fischer, Warren P.; Hamilton, Leonard J.; ...

2013-02-01

In an effort aimed at predicting the combustion behavior of a new fuel in a conventional diesel engine, cetane (n-hexadecane) fuel was used in a military engine across the entire speed–load operating range. The ignition delay was characterized for this fuel at each operating condition. A chemical ignition delay was also predicted across the speed–load range using a detailed chemical kinetic mechanism with a constant pressure reactor model. At each operating condition, the measured in-cylinder pressure and predicted temperature at the start of injection were applied to the detailed n-hexadecane kinetic mechanism, and the chemical ignition delay was predicted withoutmore » any kinetic mechanism calibration. The modeling results show that fuel–air parcels developed from the diesel spray with an equivalence ratio of 4 are the first to ignite. The chemical ignition delay results also showed decreasing igntion delays with increasing engine load and speed, just as the experimental data revealed. At lower engine speeds and loads, the kinetic modeling results show the characteristic two-stage negative temperature coefficient behavior of hydrocarbon fuels. However, at high engine speeds and loads, the reactions do not display negative temperature coefficient behavior, as the reactions proceed directly into high-temperature pathways due to higher temperatures and pressure at injection. A moderate difference between the total and chemical ignition delays was then characterized as a phyical delay period that scales inversely with engine speed. This physical delay time is representative of the diesel spray development time and is seen to become a minority fraction of the total igntion delay at higher engine speeds. In addition, the approach used in this study suggests that the ignition delay and thus start of combustion may be predicted with reasonable accuracy using kinetic modeling to determine the chemical igntion delay. Then, in conjunction with the physical delay time (experimental or modeling based), a new fuel’s acceptability in a conventional engine could be assessed by determining that the total ignition delay is not too short or too long.« less

Infrared Imagery of Solid Rocket Exhaust Plumes

NASA Technical Reports Server (NTRS)

Moran, Robert P.; Houston, Janice D.

2011-01-01

The Ares I Scale Model Acoustic Test program consisted of a series of 18 solid rocket motor static firings, simulating the liftoff conditions of the Ares I five-segment Reusable Solid Rocket Motor Vehicle. Primary test objectives included acquiring acoustic and pressure data which will be used to validate analytical models for the prediction of Ares 1 liftoff acoustics and ignition overpressure environments. The test article consisted of a 5% scale Ares I vehicle and launch tower mounted on the Mobile Launch Pad. The testing also incorporated several Water Sound Suppression Systems. Infrared imagery was employed during the solid rocket testing to support the validation or improvement of analytical models, and identify corollaries between rocket plume size or shape and the accompanying measured level of noise suppression obtained by water sound suppression systems.

14 CFR 25.1165 - Engine ignition systems.

Code of Federal Regulations, 2013 CFR

2013-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... that draw electrical energy from the same source. (c) The design of the engine ignition system must...

14 CFR 25.1165 - Engine ignition systems.

Code of Federal Regulations, 2012 CFR

2012-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... that draw electrical energy from the same source. (c) The design of the engine ignition system must...

Benefit from NASA

NASA Image and Video Library

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for the new flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes

Land Mines Removal

NASA Technical Reports Server (NTRS)

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for new the flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes.

Land Mines Removal

NASA Technical Reports Server (NTRS)

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for the new flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes

Ignition of Liquid Fuel Spray and Simulated Solid Rocket Fuel by Photoignition of Carbon Nanotube Utilizing a Camera Flash

DTIC Science & Technology

2011-12-01

10,11 There has been a recent report on the photoignition of graphene oxide for fuel ignition applications.12 In this report, we will describe the...slide Aluminum foil Glass petri dish Xe flash Camera Sample Black spray paint Figure 2- Schematic and photographs of the experimental setup...Gilje, Sergey Dubin, Alireza Badakhshan, Jabari Farrar, Stephen. A. Danczyk, Richard B. Kaner, “Photothermal Deoxygenation of Graphene Oxide for

Liquid Rocket Engine Testing Overview

NASA Technical Reports Server (NTRS)

Rahman, Shamim

2005-01-01

Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

Exhaust gas recirculation in a homogeneous charge compression ignition engine

DOEpatents

Duffy, Kevin P [Metamora, IL; Kieser, Andrew J [Morton, IL; Rodman, Anthony [Chillicothe, IL; Liechty, Michael P [Chillicothe, IL; Hergart, Carl-Anders [Peoria, IL; Hardy, William L [Peoria, IL

2008-05-27

A homogeneous charge compression ignition engine operates by injecting liquid fuel directly in a combustion chamber, and mixing the fuel with recirculated exhaust and fresh air through an auto ignition condition of the fuel. The engine includes at least one turbocharger for extracting energy from the engine exhaust and using that energy to boost intake pressure of recirculated exhaust gas and fresh air. Elevated proportions of exhaust gas recirculated to the engine are attained by throttling the fresh air inlet supply. These elevated exhaust gas recirculation rates allow the HCCI engine to be operated at higher speeds and loads rendering the HCCI engine a more viable alternative to a conventional diesel engine.

Homogenous charge compression ignition engine having a cylinder including a high compression space

DOEpatents

Agama, Jorge R.; Fiveland, Scott B.; Maloney, Ronald P.; Faletti, James J.; Clarke, John M.

2003-12-30

The present invention relates generally to the field of homogeneous charge compression engines. In these engines, fuel is injected upstream or directly into the cylinder when the power piston is relatively close to its bottom dead center position. The fuel mixes with air in the cylinder as the power piston advances to create a relatively lean homogeneous mixture that preferably ignites when the power piston is relatively close to the top dead center position. However, if the ignition event occurs either earlier or later than desired, lowered performance, engine misfire, or even engine damage, can result. Thus, the present invention divides the homogeneous charge between a controlled volume higher compression space and a lower compression space to better control the start of ignition.

Exhaust Emissions Measured Under Real Traffic Conditions from Vehicles Fitted with Spark Ignition and Compression Ignition Engines

NASA Astrophysics Data System (ADS)

Merkisz, Jerzy; Lijewski, Piotr; Fuć, Paweł

2011-06-01

The tests performed under real traffic conditions provide invaluable information on the relations between the engine parameters, vehicle parameters and traffic conditions (traffic congestion) on one side and the exhaust emissions on the other. The paper presents the result of road tests obtained in an urban and extra-urban cycles for vehicles fitted with different engines, spark ignition engine and compression ignition engine. For the tests a portable emission analyzer SEMTECH DS. by SENSORS was used. This analyzer provides online measurement of the concentrations of exhaust emission components on a vehicle in motion under real traffic conditions. The tests were performed in city traffic. A comparative analysis has been presented of the obtained results for vehicles with individual powertrains.

14 CFR 29.1165 - Engine ignition systems.

Code of Federal Regulations, 2012 CFR

2012-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... that draw from the same source. (c) The design of the engine ignition system must account for— (1) The...

14 CFR 29.1165 - Engine ignition systems.

Code of Federal Regulations, 2013 CFR

2013-01-01

... automatically available as an alternate source of electrical energy to allow continued engine operation if any... that draw from the same source. (c) The design of the engine ignition system must account for— (1) The...

Some Effects of Injection Advance Angle, Engine-Jacket Temperature, and Speed on Combustion in a Compression-Ignition Engine

NASA Technical Reports Server (NTRS)

Rothrock, A M; Waldron, C D

1936-01-01

An optical indicator and a high-speed motion-picture camera capable of operating at the rate of 2,000 frames per second were used to record simultaneously the pressure development and the flame formation in the combustion chamber of the NACA combustion apparatus. Tests were made at engine speeds of 570 and 1,500 r.p.m. The engine-jacket temperature was varied from 100 degrees to 300 degrees F. And the injection advance angle from 13 degrees after top center to 120 degrees before top center. The results show that the course of the combustion is largely controlled by the temperature and pressure of the air in the chamber from the time the fuel is injected until the time at which combustion starts and by the ignition lag. The conclusion is presented that in a compression-ignition engine with a quiescent combustion chamber the ignition lag should be the longest that can be used without excessive rates of pressure rise; any further shortening of the ignition lag decreased the effective combustion of the engine.

Fluidized-Solid-Fuel Injection Process

NASA Technical Reports Server (NTRS)

Taylor, William

1992-01-01

Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

J-2X Powerpack hot-fire test

NASA Image and Video Library

2008-01-31

The first hot-fire test of the J-2X power pack 1A gas generator was performed Jan. 31 on the A-1 Test Stand at Stennis Space Center. Initial indications are that all test objectives were met. The test was designed as a 3.42-second helium spin start with gas generator ignition and it went the full scheduled duration. Test conductors reported a smooth start with normal shutdown and described the event as a 'good test.' The test was part of the early component testing for the new J-2X engine being built by NASA to power the Ares I and Ares V rockets that will carry humans back to the moon and on to Mars. It was performed as one in a series of 12 scheduled tests. Those tests began last November at Stennis, but the January 31 event represented the first hot-fire test. The Stennis tests are a critical step in the successful development of the J-2X engine.

Liquid Rocket Engine Testing

DTIC Science & Technology

2016-10-21

Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

Compression Ignition Engines – Revolutionary Technology That has Civilized Frontiers all Over the Globe from the Industrial Revolution into the Twenty-First Century

DOE Office of Scientific and Technical Information (OSTI.GOV)

Ciatti, Stephen A.

The history, present and future of the compression ignition engine is a fascinating story that spans over 100 years, from the time of Rudolf Diesel to the highly regulated and computerized engines of the 21st Century. The development of these engines provided inexpensive, reliable and high power density machines to allow transportation, construction and farming to be more productive with less human effort than in any previous period of human history. The concept that fuels could be consumed efficiently and effectively with only the ignition of pressurized and heated air was a significant departure from the previous coal-burning architecture ofmore » the 1800s. Today, the compression ignition engine is undergoing yet another revolution. The equipment that provides transport, builds roads and infrastructure, and harvests the food we eat needs to meet more stringent requirements than ever before. How successfully 21st Century engineers are able to make compression ignition engine technology meet these demands will be of major influence in assisting developing nations (with over 50% of the world’s population) achieve the economic and environmental goals they seek.« less

X-1 aircraft in flight

NASA Technical Reports Server (NTRS)

1949-01-01

The first of the rocket-powered research aircraft, the X-1 (originally designated the XS-1), was a bullet-shaped airplane that was built by the Bell Aircraft Company for the US Air Force and the National Advisory Committee for Aeronautics (NACA). The mission of the X-1 was to investigate the transonic speed range (speeds from just below to just above the speed of sound) and, if possible, to break the 'sound barrier'. The first of the three X-1s was glide-tested at Pinecastle Field, FL, in early 1946. The first powered flight of the X-1 was made on Dec. 9, 1946, at Muroc Army Air Field (later redesignated Edwards Air Force Base) with Chalmers Goodlin, a Bell test pilot,at the controls. On Oct. 14, 1947, with USAF Captain Charles 'Chuck' Yeager as pilot, the aircraft flew faster than the speed of sound for the first time. Captain Yeager ignited the four-chambered XLR-11 rocket engines after being air-launched from under the bomb bay of a B-29 at 21,000 ft. The 6,000-lb thrust ethyl alcohol/liquid oxygen burning rockets, built by Reaction Motors, Inc., pushed him up to a speed of 700 mph in level flight. Captain Yeager was also the pilot when the X-1 reached its maximum speed of 957 mph. Another USAF pilot. Lt. Col. Frank Everest, Jr., was credited with taking the X-1 to its maximum altitude of 71,902 ft. Eighteen pilots in all flew the X-1s. The number three plane was destroyed in a fire before evermaking any powered flights. A single-place monoplane, the X-1 was 31 ft long, 10 ft high, and had a wingspan of 29 ft. It weighed 4,900 lb and carried 8,200 lb of fuel. It had a flush cockpit with a side entrance and no ejection seat. The following movie runs about 20 seconds, and shows several air-to-air views of X-1 Number 2 and its modified B-50 mothership. It begins with different angles of the X-1 in-flight while mated to the B-50's bomb bay, and ends showing the air-launch. The X-1 drops below the B-50, then accelerates away as the rockets ignite.

Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

NASA Astrophysics Data System (ADS)

Marshall, Joel H.

A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

14 CFR 29.1145 - Ignition switches.

Code of Federal Regulations, 2010 CFR

2010-01-01

... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 29.1145 Ignition switches. (a) Ignition switches must control each ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 29.1145 - Ignition switches.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 29.1145 Ignition switches. (a) Ignition switches must control each ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 29.1145 - Ignition switches.

Code of Federal Regulations, 2013 CFR

2013-01-01

... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 29.1145 Ignition switches. (a) Ignition switches must control each ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 29.1145 - Ignition switches.

Code of Federal Regulations, 2011 CFR

2011-01-01

... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 29.1145 Ignition switches. (a) Ignition switches must control each ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

14 CFR 29.1145 - Ignition switches.

Code of Federal Regulations, 2014 CFR

2014-01-01

... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Powerplant Controls and Accessories § 29.1145 Ignition switches. (a) Ignition switches must control each ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c...

Space engine safety system

NASA Technical Reports Server (NTRS)

Maul, William A.; Meyer, Claudia M.

1991-01-01

A rocket engine safety system was designed to initiate control procedures to minimize damage to the engine or vehicle or test stand in the event of an engine failure. The features and the implementation issues associated with rocket engine safety systems are discussed, as well as the specific concerns of safety systems applied to a space-based engine and long duration space missions. Examples of safety system features and architectures are given, based on recent safety monitoring investigations conducted for the Space Shuttle Main Engine and for future liquid rocket engines. Also, the general design and implementation process for rocket engine safety systems is presented.

KSC-2009-1973

NASA Image and Video Library

2009-03-06

CAPE CANAVERAL, Fla. – On Launch Pad 17-B at Cape Canaveral Air Force Station in Florida, the engines on United Launch Alliance's Delta II rocket carrying NASA's Kepler spacecraft ignite. Liftoff was on time at 10:49 p.m. EST. Kepler is a spaceborne telescope designed to search the nearby region of our galaxy for Earth-size planets orbiting in the habitable zone of stars like our sun. The habitable zone is the region around a star where temperatures permit water to be liquid on a planet's surface. The challenge for Kepler is to look at a large number of stars in order to statistically estimate the total number of Earth-size planets orbiting sun-like stars in the habitable zone. Kepler will survey more than 100,000 stars in our galaxy. Photo credit: NASA/Tony Gray

STS-88 Mission Highlights Resources Tape. Tape A

NASA Technical Reports Server (NTRS)

1999-01-01

The STS-88 flight crew, Commander Robert D. Cabana, Pilot Frederick W. Sturckow, and Mission Specialists Nancy J. Currie, James H. Newman, Jerry L. Ross, and Sergei Krikalev present a video overview of their space flight. This is the first of three videos which show the highlights of the Endeavour mission. Important visual images include pre-launch activities such as the eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also included are various panoramic views of the shuttle on the pad. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters. Once on-orbit crew members are seen delivering and connecting the UNITY Connecting Module to the ZARYA Control Module.

STS-82 Post Flight Presentation

NASA Technical Reports Server (NTRS)

1997-01-01

The STS-82 crew, Commander Kenneth D. Bowersox, Pilot Scott J. Horowitz, Payload Commander Mark C. Lee, and Mission Specialists Gregory J. Harbaugh, Steven L. Smith, Joseph R. Tanner, and Steven A. Hawley present a video and still picture overview of their mission. Included in the presentation are the following: the pre-launch activities such as eating the traditional breakfast, being suited up, and riding out to the launch pad, various panoramic views of the shuttle on the pad, the countdown, engine ignition, launch, shuttle roll maneuver, separation of the Solid Rocket Boosters (SRB) from the shuttle, survey of the payload bay with the Shuttle's 50-foot remote manipulator system (RMS), the successful retrieve of the Hubble Space Telescope (HST), EVAs to repair HST, release of HST, and the shuttle's landing.

40 CFR 94.9 - Compliance with emission standards.

Code of Federal Regulations, 2011 CFR

2011-07-01

... engineering analysis of information equivalent to such in-use data, such as data from research engines or... (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.9 Compliance with emission standards. (a) The general...

40 CFR 94.9 - Compliance with emission standards.

Code of Federal Regulations, 2010 CFR

2010-07-01

... engineering analysis of information equivalent to such in-use data, such as data from research engines or... (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.9 Compliance with emission standards. (a) The general...

40 CFR 94.9 - Compliance with emission standards.

Code of Federal Regulations, 2014 CFR

2014-07-01

... engineering analysis of information equivalent to such in-use data, such as data from research engines or... (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.9 Compliance with emission standards. (a) The general...

40 CFR 94.9 - Compliance with emission standards.

Code of Federal Regulations, 2012 CFR

2012-07-01

... engineering analysis of information equivalent to such in-use data, such as data from research engines or... (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.9 Compliance with emission standards. (a) The general...

40 CFR 94.9 - Compliance with emission standards.

Code of Federal Regulations, 2013 CFR

2013-07-01

... engineering analysis of information equivalent to such in-use data, such as data from research engines or... (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.9 Compliance with emission standards. (a) The general...

Developments in REDES: The rocket engine design expert system

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) is being developed at the NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP, a nozzle design program named RAO, a regenerative cooling channel performance evaluation code named RTE, and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Developments in REDES: The Rocket Engine Design Expert System

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Friction of Compression-ignition Engines

NASA Technical Reports Server (NTRS)

Moore, Charles S; Collins, John H , Jr

1936-01-01

The cost in mean effective pressure of generating air flow in the combustion chambers of single-cylinder compression-ignition engines was determined for the prechamber and the displaced-piston types of combustion chamber. For each type a wide range of air-flow quantities, speeds, and boost pressures was investigated. Supplementary tests were made to determine the effect of lubricating-oil temperature, cooling-water temperature, and compression ratio on the friction mean effective pressure of the single-cylinder test engine. Friction curves are included for two 9-cylinder, radial, compression-ignition aircraft engines. The results indicate that generating the optimum forced air flow increased the motoring losses approximately 5 pounds per square inch mean effective pressure regardless of chamber type or engine speed. With a given type of chamber, the rate of increase in friction mean effective pressure with engine speed is independent of the air-flow speed. The effect of boost pressure on the friction cannot be predicted because the friction was decreased, unchanged, or increased depending on the combustion-chamber type and design details. High compression ratio accounts for approximately 5 pounds per square inch mean effective pressure of the friction of these single-cylinder compression-ignition engines. The single-cylinder test engines used in this investigation had a much higher friction mean effective pressure than conventional aircraft engines or than the 9-cylinder, radial, compression-ignition engines tested so that performance should be compared on an indicated basis.

A Comparison of Ignition Characteristics of Diesel Fuels as Determined in Engines and in a Constant-volume Bomb

NASA Technical Reports Server (NTRS)

Selden, Robert F

1939-01-01

Ignition-lag data have been obtained for seven fuels injected into heated, compressed air under conditions simulating those in a compression-ignition engine. The results of the bomb tests have been compared with similar engine data, and the differences between the two sets of results are explained in terms of the response of each fuel to variations in air density and temperature.

Development of Mechanics in Support of Rocket Technology in Ukraine

NASA Astrophysics Data System (ADS)

Prisnyakov, Vladimir

2003-06-01

The paper analyzes the advances of mechanics made in Ukraine in resolving various problems of space and rocket technology such as dynamics and strength of rockets and rocket engines, rockets of different purpose, electric rocket engines, and nonstationary processes in various systems of rockets accompanied by phase transitions of working media. Achievements in research on the effect of vibrations and gravitational fields on the behavior of space-rocket systems are also addressed. Results obtained in investigating the reliability and structural strength durability conditions for nuclear installations, solid- and liquid-propellant engines, and heat pipes are presented

Rocket Engine Oscillation Diagnostics

NASA Technical Reports Server (NTRS)

Nesman, Tom; Turner, James E. (Technical Monitor)

2002-01-01

Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

Credit BG. View shows the north and west facades of ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. View shows the north and west facades of the building as seen when looking east southeast (124°). Igniters for solid rocket motors were built and tested here. This building was rated for a maximum of 20 pounds (9.1 Kg) of class 1.1 materials and four personnel. Note the lightning rods on roof corners and the exterior electrical system - Jet Propulsion Laboratory Edwards Facility, Igniter Laboratory, Edwards Air Force Base, Boron, Kern County, CA

Lightning flashes triggered in altitude by the rocket and wire technique

NASA Technical Reports Server (NTRS)

Laroche, P.; Bondiou, A.; Berard, A. Eybert; Barret, L.; Berlandis, J. P.; Terrier, G.; Jafferis, W.

1989-01-01

Electrical measurements were conducted in 1987 and 1988 on streamer and leader discharges occurring during the first stages of a triggered flash. This paper describes the pulsing phenomenon observed at positive leader onset (typical pulsing rate 25 microns), and it is shown that the same process happened in the case of the ignition of a flash triggered in altitude; with a wire several hundred meters long, positive leader propagates alone for several ms before the ignition of the downward negative stepped leader.

Engine Valve Actuation For Combustion Enhancement

DOEpatents

Reitz, Rolf Deneys; Rutland, Christopher J.; Jhavar, Rahul

2004-05-18

A combustion chamber valve, such as an intake valve or an exhaust valve, is briefly opened during the compression and/or power strokes of a 4-stroke combustion cycle in an internal combustion engine (in particular, a diesel or CI engine). The brief opening may (1) enhance mixing withing the combustion chamber, allowing more complete oxidation of particulates to decrease engine emissions; and/or may (2) delay ignition until a more desirable time, potentially allowing a means of timing ignition in otherwise difficult-to-control conditions, e.g., in HCCI (Homogeneous Charge Compression Ignition) conditions.

Engine valve actuation for combustion enhancement

DOEpatents

Reitz, Rolf Deneys [Madison, WI; Rutland, Christopher J [Madison, WI; Jhavar, Rahul [Madison, WI

2008-03-04

A combustion chamber valve, such as an intake valve or an exhaust valve, is briefly opened during the compression and/or power strokes of a 4-strokes combustion cycle in an internal combustion engine (in particular, a diesel or CI engine). The brief opening may (1) enhance mixing withing the combustion chamber, allowing more complete oxidation of particulates to decrease engine emissions; and/or may (2) delay ignition until a more desirable time, potentially allowing a means of timing ignition in otherwise difficult-to-control conditions, e.g., in HCCI (Homogeneous Charge Compression Ignition) conditions.

Ignition and combustion: Low compression ratio, high output diesel

NASA Technical Reports Server (NTRS)

1981-01-01

The feasibility of converting a spark ignition aircraft engine GTSI0-520 to compression ignition without increasing the peak combustion pressure of 1100 lbs/sq.in. was determined. The final contemplated utilized intake air heating at idle and light load and a compression ratio of about 10:1 with a small amount of fumigation (the addition of about 15% fuel into the combustion air before the cylinder). The engine used was a modification of a Continental-Teledyne gasoline engine cylinder from the GTSI0-520 supercharged aircraft engine.

Crank angle detecting system for engines

DOE Office of Scientific and Technical Information (OSTI.GOV)

Yuzawa, H.; Nishiyama, M.; Nakamura, K.

1988-05-31

An ignition system for a multi-cylinder internal combustion engine is described comprising: (a) engine cylinders in which spark plugs are installed respectively, (b) indicating means disposed so as to synchronize with an engine crankshaft and formed with a large number of slits and a small number of slits, the large number of slits being provided for indicating crankshaft angular positions and the small number of slits being provided for indicating predetermined piston strokes and wherein the small number of slits have mutually different widths from each other to distinguish between piston strokes of at least the groups of cylinders; (c)more » sensing means for sensing crankshaft angular positions in cooperation with the large number of slits of the indicating means and outputting a crank angle signal representing the crankshaft angular position and for sensing the predetermined piston strokes in cooperation with the small number of slits and outputting different width piston stroke signals corresponding to the different width slits; (d) discriminating means for identifying each cylinder group and outputting cylinder group identification signals on the basis of the different width stroke signals derived from the sensing means; (e) ignition timing determining means for generating an ignition timing signal on the basis of the crank angle signal; (f) ignition coil controlling means for generating ignition coil current signals corresponding to the cylinder group identification signals; and (g) ignition timing controlling means for generating cylinder group ignition signals in response to the ignition coil current signals and ignition timing signal so that the spark plugs of each cylinder group are ignited at a proper time.« less

78 FR 50412 - California State Nonroad Engine Pollution Control Standards; Amendments to Spark Ignition Marine...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-19

... Engine Pollution Control Standards; Amendments to Spark Ignition Marine Engine and Boat Regulations... Marine Engine and Boat Regulations (2008 Marine SI Amendments or 2008 Amendments). CARB requested EPA... the 2008 Marine SI Amendments. DATES: EPA has tentatively scheduled a public hearing concerning CARB's...

Dual Expander Cycle Rocket Engine with an Intermediate, Closed-cycle Heat Exchanger

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2008-01-01

A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger thermally coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

Astronautics

NASA Technical Reports Server (NTRS)

1977-01-01

Principles of rocket engineering, flight dynamics, and trajectories are discussed in this summary of Soviet rocket development and technology. Topics include rocket engine design, propellants, propulsive efficiency, and capabilities required for orbital launch. The design of the RD 107, 108, 119, and 214 rocket engines and their uses in various satellite launches are described. NASA's Saturn 5 and Atlas Agena launch vehicles are used to illustrate the requirements of multistage rockets.

Self-ignition of S.I. engine model fuels: A shock tube investigation at high pressure

DOE Office of Scientific and Technical Information (OSTI.GOV)

Fieweger, K.; Blumenthal, R.; Adomeit, G.

1997-06-01

The self-ignition of several spark-ignition (SI) engine fuels (iso-octane, methanol, methyl tert-butyl ether and three different mixtures of iso-octane and n-heptane), mixed with air, was investigated experimentally under relevant engine conditions by the shock tube technique. Typical modes of the self-ignition process were registered cinematographically. For temperatures relevant to piston engine combustion, the self-ignition process always starts as an inhomogeneous, deflagrative mild ignition. This instant is defined by the ignition delay time, {tau}{sub defl}. The deflagration process in most cases is followed by a secondary explosion (DDT). This transition defines a second ignition delay time, {tau}{sub DDT}, which is amore » suitable approximation for the chemical ignition delay time, if the change of the thermodynamic conditions of the unburned test gas due to deflagration is taken into account. For iso-octane at p = 40 bar, a NTC (negative temperature coefficient), behavior connected with a two step (cool flame) self-ignition at low temperatures was observed. This process was very pronounced for rich and less pronounced for stoichiometric mixtures. The results of the {tau}{sub DDT} delays of the stoichiometric mixtures were shortened by the primary deflagration process in the temperature range between 800 and 1,000 K. Various mixtures of iso-octane and n-heptane were investigated. The results show a strong influence of the n-heptane fraction in the mixture, both on the ignition delay time and on the mode of self-ignition. The self-ignition of methanol and MTBE (methyl tert-butyl ether) is characterized by a very pronounced initial deflagration. For temperatures below 900 K (methanol: 800 K), no secondary explosion occurs. Taking into account the pressure increase due to deflagration, the measured delays {tau}{sub DDT} of the secondary explosion are shortened by up to one order of magnitude.« less

Control of Propellant Lead/Lag to the LAE in the AXAF Propulsion System

NASA Technical Reports Server (NTRS)

Casillas, A. R.; Eninger, J.; Joseph, G.; Kenney, J.; Trinidad, M.

1998-01-01

Control of the rate at which hypergolic propellants are supplied to a rocket engine prior to ignition is critically important. Potentially damaging explosions may result from excessive lead of either propellant into the combustion chamber. Because the injector fill process is governed by the engine as well as the propellant feed system design, proper management of this issue must take both into consideration. This was recognized early in the development of TRW's Advanced Columbium-Liquid Apogee Engine (LAE), which was flight-qualified in 1996 to maneuver the Advanced X-Ray Astrophysics Facility (AXAF) spacecraft into orbit. The LAE runs on hydrazine and nitrogen tetroxide (MON-3) at a nominal mixture ratio of 1.0. This paper describes the comprehensive test program conducted to ensure reliable startup operation of the LAE in the AYAF propulsion system. The most significant factors affecting chamber fuel lead were found to be: (1) engine location, (2) propellant saturation level, (3) amount of undissolved gas in the lines, and (4) off- nominal tank pressures. Hot-fire tests at a chamber fuel lead range over and above that expected for the LAEs in AXAF demonstrated extremely tolerant behavior of the engine. AY-AF is scheduled for launch on NASA's STS-93 in December 1998.

Reusable rocket engine optical condition monitoring

NASA Technical Reports Server (NTRS)

Wyett, L.; Maram, J.; Barkhoudarian, S.; Reinert, J.

1987-01-01

Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

40 CFR 91.1001 - Applicability.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM MARINE SPARK-IGNITION ENGINES Exclusion and Exemption of Marine SI Engines § 91.1001 Applicability. The requirements of this subpart K are applicable to all marine spark-ignition propulsion engines...

Passenger Car Spark Ignition Data Base : Volume 1. Executive Summary.

DOT National Transportation Integrated Search

1979-12-01

Test data was obtained from spark ignition production and preproduction engines at the engine and vehicle level. The engines were applicable for vehicles 2000 to 3000 pounds in weight. The data obtained provided trade-offs between fuel economy, power...

Investigation of the fundamentals of low-energy nanosecond pulse ignition: Final CRADA Report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wallner, Thomas; Scarcelli, Riccardo; Zhang, Anqi

A detailed investigation of the fundamentals of low-energy nanosecond pulse ignition was performed with the objective to overcome the barrier presented by limited knowledge and characterization of nonequilibrium plasma ignition for realistic internal combustion engine applications (be it in the automotive or power generation field) and shed light on the mechanisms which improve the performance of the advanced TPS ignition system compared to conventional state-of-the-art hardware. Three main tasks of the research included experimental evaluation on a single-cylinder automotive gasoline engine, experimental evaluation on a single-cylinder stationary natural gas engine and energy quantification using x-ray diagnostics.

A reduced mechanism for biodiesel surrogates with low temperature chemistry for compression ignition engine applications

NASA Astrophysics Data System (ADS)

Luo, Zhaoyu; Plomer, Max; Lu, Tianfeng; Som, Sibendu; Longman, Douglas E.

2012-04-01

Biodiesel is a promising alternative fuel for compression ignition (CI) engines. It is a renewable energy source that can be used in these engines without significant alteration in design. The detailed chemical kinetics of biodiesel is however highly complex. In the present study, a skeletal mechanism with 123 species and 394 reactions for a tri-component biodiesel surrogate, which consists of methyl decanoate, methyl 9-decanoate and n-heptane was developed for simulations of 3-D turbulent spray combustion under engine-like conditions. The reduction was based on an improved directed relation graph (DRG) method that is particularly suitable for mechanisms with many isomers, followed by isomer lumping and DRG-aided sensitivity analysis (DRGASA). The reduction was performed for pressures from 1 to 100 atm and equivalence ratios from 0.5 to 2 for both extinction and ignition applications. The initial temperatures for ignition were from 700 to 1800 K. The wide parameter range ensures the applicability of the skeletal mechanism under engine-like conditions. As such the skeletal mechanism is applicable for ignition at both low and high temperatures. Compared with the detailed mechanism that consists of 3299 species and 10806 reactions, the skeletal mechanism features a significant reduction in size while still retaining good accuracy and comprehensiveness. The validations of ignition delay time, flame lift-off length and important species profiles were also performed in 3-D engine simulations and compared with the experimental data from Sandia National Laboratories under CI engine conditions.

Relation of Hydrogen and Methane to Carbon Monoxide in Exhaust Gases from Internal-Combustion Engines

NASA Technical Reports Server (NTRS)

Gerrish, Harold C; Tessmann, Arthur M

1935-01-01

The relation of hydrogen and methane to carbon monoxide in the exhaust gases from internal-combustion engines operating on standard-grade aviation gasoline, fighting-grade aviation gasoline, hydrogenated safety fuel, laboratory diesel fuel, and auto diesel fuel was determined by analysis of the exhaust gases. Two liquid-cooled single-cylinder spark-ignition, one 9-cylinder radial air-cooled spark-ignition, and two liquid-cooled single-cylinder compression-ignition engines were used.

Ignition process in Diesel engines

NASA Technical Reports Server (NTRS)

Wentzel, W

1936-01-01

This report analyzes the heating and vaporization process of fuel droplets in a compression-ignition engine on the basis of the theory of similitude - according to which, the period for heating and complete vaporization of the average size fuel drop is only a fraction of the actually observed ignition lag. The result is that ignition takes place in the fuel vapor air mixture rather than on the surface of the drop. The theoretical result is in accord with the experimental observations by Rothrock and Waldron. The combustion shock occurring at lower terminal compression temperature, especially in the combustion of coal-tar oil, is attributable to a simultaneous igniting of a larger fuel-vapor volume formed prior to ignition.

Effects of Fuel Laminar Flame Speed Compared to Engine Tumble Ratio, Ignition Energy, and Injection Strategy on Lean and EGR Dilute Spark Ignition Combustion

DOE PAGES

Kolodziej, Christopher P.; Pamminger, Michael; Sevik, James; ...

2017-03-28

Previously we show that fuels with higher laminar flame speed also have increased tolerance to EGR dilution. In this work, the effects of fuel laminar flame speed on both lean and EGR dilute spark ignition combustion stability were examined. Fuels blends of pure components (iso-octane, n-heptane, toluene, ethanol, and methanol) were derived at two levels of laminar flame speed. Each fuel blend was tested in a single-cylinder spark-ignition engine under both lean-out and EGR dilution sweeps until the coefficient of variance of indicated mean effective pressure increased above thresholds of 3% and 5%. The relative importance of fuel laminar flamemore » speed to changes to engine design parameters (spark ignition energy, tumble ratio, and port vs. direct injection) was also assessed. Our results showed that fuel laminar flame speed can have as big an effect on lean or EGR dilute engine operation as engine design parameters, with the largest effects seen during EGR dilute operation and when changes were made to cylinder charge motion.« less

Effects of Fuel Laminar Flame Speed Compared to Engine Tumble Ratio, Ignition Energy, and Injection Strategy on Lean and EGR Dilute Spark Ignition Combustion

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kolodziej, Christopher P.; Pamminger, Michael; Sevik, James

Previously we show that fuels with higher laminar flame speed also have increased tolerance to EGR dilution. In this work, the effects of fuel laminar flame speed on both lean and EGR dilute spark ignition combustion stability were examined. Fuels blends of pure components (iso-octane, n-heptane, toluene, ethanol, and methanol) were derived at two levels of laminar flame speed. Each fuel blend was tested in a single-cylinder spark-ignition engine under both lean-out and EGR dilution sweeps until the coefficient of variance of indicated mean effective pressure increased above thresholds of 3% and 5%. The relative importance of fuel laminar flamemore » speed to changes to engine design parameters (spark ignition energy, tumble ratio, and port vs. direct injection) was also assessed. Our results showed that fuel laminar flame speed can have as big an effect on lean or EGR dilute engine operation as engine design parameters, with the largest effects seen during EGR dilute operation and when changes were made to cylinder charge motion.« less

Compression-ignition engine performance with undoped and doped fuel oils and alcohol mixtures

NASA Technical Reports Server (NTRS)

Moore, Charles S; Foster, Hampton H

1939-01-01

Several fuel oils, doped fuel oils, and mixtures of alcohol and fuel oil were tested in a high-speed, single-cylinder, compression-ignition engine to determine power output, fuel consumption, and ignition and combustion characteristics. Fuel oils or doped fuel oils of high octane number had shorter ignition lags, lower rates of pressure rise, and gave smoother engine operation than fuel oils or doped fuel oils of low octane number. Higher engine rotative speeds and boost pressures resulted in smoother engine operation and permitted the use of fuel oils of relatively low octane number. Although the addition of a dope to a fuel oil decreased the ignition lag and the rate of pressure rise, the ensuing rate of combustion was somewhat slower than for the undoped fuel oil so that the effectiveness of combustion was practically unchanged. Alcohol used as an auxiliary fuel, either as a mixture or by separate injection, increased the rates of pressure rise and induced roughness. In general, the power output decreased as the proportion of alcohol increased and, below maximum power, varied with the heating value of the total fuel charge.

Pegasus XL CYGNSS Second Launch Attempt, Drop & Launch of Rocket

NASA Image and Video Library

2016-12-15

The Orbital ATK Pegasus XL rocket carrying NASA's Cyclone Global Navigation Satellite System, or CYGNSS, spacecraft is released and the first stage ignites at 8:37 a.m. EST. The rocket was released from the Orbital ATK L-1011 Stargazer aircraft flying over the Atlantic Ocean offshore from Daytona Beach, Florida following takeoff from the Skid Strip at Cape Canaveral Air Force Station. This image was taken from a NASA F-18 chase plane provided by Armstrong Flight Research Center in California. The CYGNSS satellites will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes.

Experimental research and design planning in the field of liquid-propellant rocket engines conducted between 1934 - 1944 by the followers of F. A. Tsander

NASA Technical Reports Server (NTRS)

Dushkin, L. S.

1977-01-01

The development of the following Liquid-Propellant Rocket Engines (LPRE) is reviewed: (1) an alcohol-oxygen single-firing LPRE for use in wingless and winged rockets, (2) a similar multifiring LPRE for use in rocket gliders, (3) a combined solid-liquid propellant rocket engine, and (4) an aircraft LPRE operating on nitric acid and kerosene.

Rocket-Based Combined Cycle Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

2001-01-01

The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ OF THE X-15 ENGINE TEST COMPLEX. Looking northeast. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

Passenger Car Spark Ignition Data Base : Volume 2. Discussion and Results.

DOT National Transportation Integrated Search

1979-12-01

Test data was obtained from spark ignition production and preproduction engines at the engine and vehicle level. The engines were applicable for vehicles 2000 to 3000 pounds in weight. The data obtained provided trade-offs between fuel economy, power...

40 CFR 94.12 - Interim provisions.

Code of Federal Regulations, 2014 CFR

2014-07-01

... differences in testing will not affect NOX emission rates. (g) Flexibility for engines over 560kW... EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.12 Interim provisions. This section contains provisions that apply...

40 CFR 94.12 - Interim provisions.

Code of Federal Regulations, 2010 CFR

2010-07-01

... differences in testing will not affect NOX emission rates. (g) Flexibility for engines over 560kW...) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.12 Interim provisions. This section contains provisions...

40 CFR 94.12 - Interim provisions.

Code of Federal Regulations, 2013 CFR

2013-07-01

... differences in testing will not affect NOX emission rates. (g) Flexibility for engines over 560kW...) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.12 Interim provisions. This section contains provisions...

40 CFR 94.12 - Interim provisions.

Code of Federal Regulations, 2012 CFR

2012-07-01

... differences in testing will not affect NOX emission rates. (g) Flexibility for engines over 560kW...) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.12 Interim provisions. This section contains provisions...

40 CFR 94.12 - Interim provisions.

Code of Federal Regulations, 2011 CFR

2011-07-01

... differences in testing will not affect NOX emission rates. (g) Flexibility for engines over 560kW...) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.12 Interim provisions. This section contains provisions...

Passenger Car Spark Ignition Data Base : Volume 3. Miscellaneous Data. Part 1.

DOT National Transportation Integrated Search

1979-12-01

Test data was obtained from spark ignition production and preproduction engines at the engine and vehicle level. The engines were applicable for vehicles 2000 to 3000 pounds in weight. The data obtained provided trade-offs between fuel economy, power...

Passenger Car Spark Ignition Data Base : Volume 3. Miscellaneous Data. Part 2.

DOT National Transportation Integrated Search

1979-12-01

Test data was obtained from spark ignition production and preproduction engines at the engine and vehicle level. The engines were applicable for vehicles 2000 to 3000 pounds in weight. The data obtained provided trade-offs between fuel economy, power...

On the assessment of performance and emissions characteristics of a SI engine provided with a laser ignition system

NASA Astrophysics Data System (ADS)

Birtas, A.; Boicea, N.; Draghici, F.; Chiriac, R.; Croitoru, G.; Dinca, M.; Dascalu, T.; Pavel, N.

2017-10-01

Performance and exhaust emissions of spark ignition engines are strongly dependent on the development of the combustion process. Controlling this process in order to improve the performance and to reduce emissions by ensuring rapid and robust combustion depends on how ignition stage is achieved. An ignition system that seems to be able for providing such an enhanced combustion process is that based on plasma generation using a Q-switched solid state laser that delivers pulses with high peak power (of MW-order level). The laser-spark devices used in the present investigations were realized using compact diffusion-bonded Nd:YAG/Cr4+:YAG ceramic media. The laser igniter was designed, integrated and built to resemble a classical spark plug and therefore it could be mounted directly on the cylinder head of a passenger car engine. In this study are reported the results obtained using such ignition system provided for a K7M 710 engine currently produced by Renault-Dacia, where the standard calibrations were changed towards the lean mixtures combustion zone. Results regarding the performance, the exhaust emissions and the combustion characteristics in optimized spark timing conditions, which demonstrate the potential of such an innovative ignition system, are presented.

Mobile Source Emissions Regulatory Compliance Data Inventory

EPA Pesticide Factsheets

The Mobile Source Emissions Regulatory Compliance Data Inventory data asset contains measured summary compliance information on light-duty, heavy-duty, and non-road engine manufacturers by model, as well as fee payment data required by Title II of the 1990 Amendments to the Clean Air Act, to certify engines for sale in the U.S. and collect compliance certification fees. Data submitted by manufacturers falls into 12 industries: Heavy Duty Compression Ignition, Marine Spark Ignition, Heavy Duty Spark Ignition, Marine Compression Ignition, Snowmobile, Motorcycle & ATV, Non-Road Compression Ignition, Non-Road Small Spark Ignition, Light-Duty, Evaporative Components, Non-Road Large Spark Ignition, and Locomotive. Title II also requires the collection of fees from manufacturers submitting for compliance certification. Manufacturers submit data on an annual basis, to document engine model changes for certification. Manufacturers also submit compliance information on already certified in-use vehicles randomly selected by the EPA (1) year into their life and (4) years into their life to ensure that emissions systems continue to function appropriately over time.The EPA performs targeted confirmatory tests on approximately 15% of vehicles submitted for certification. Confirmatory data on engines is associated with its corresponding submission data to verify the accuracy of manufacturer submission beyond standard business rules.Section 209 of the 1990 Amendments to the Clea

40 CFR Appendix I to Part 92 - Emission Related Locomotive and Engine Parameters and Specifications

Code of Federal Regulations, 2011 CFR

2011-07-01

... injection—non-compression ignition engines. a. Control parameters and calibrations. b. Idle mixture. c. Fuel...(s). i. Injector timing calibration. 4. Fuel injection—compression ignition engines. a. Control... restriction. III. Fuel System. 1. General. a. Engine idle speed. 2. Carburetion. a. Air-fuel flow calibration...

40 CFR Appendix I to Part 92 - Emission Related Locomotive and Engine Parameters and Specifications

Code of Federal Regulations, 2014 CFR

2014-07-01

... injection—non-compression ignition engines. a. Control parameters and calibrations. b. Idle mixture. c. Fuel...(s). i. Injector timing calibration. 4. Fuel injection—compression ignition engines. a. Control... restriction. III. Fuel System. 1. General. a. Engine idle speed. 2. Carburetion. a. Air-fuel flow calibration...

40 CFR Appendix I to Part 92 - Emission Related Locomotive and Engine Parameters and Specifications

Code of Federal Regulations, 2013 CFR

2013-07-01

... injection—non-compression ignition engines. a. Control parameters and calibrations. b. Idle mixture. c. Fuel...(s). i. Injector timing calibration. 4. Fuel injection—compression ignition engines. a. Control... restriction. III. Fuel System. 1. General. a. Engine idle speed. 2. Carburetion. a. Air-fuel flow calibration...

40 CFR Appendix I to Part 92 - Emission Related Locomotive and Engine Parameters and Specifications

Code of Federal Regulations, 2012 CFR

2012-07-01

... injection—non-compression ignition engines. a. Control parameters and calibrations. b. Idle mixture. c. Fuel...(s). i. Injector timing calibration. 4. Fuel injection—compression ignition engines. a. Control... restriction. III. Fuel System. 1. General. a. Engine idle speed. 2. Carburetion. a. Air-fuel flow calibration...

Researches on Preliminary Chemical Reactions in Spark-Ignition Engines

NASA Technical Reports Server (NTRS)

Muehlner, E.

1943-01-01

Chemical reactions can demonstrably occur in a fuel-air mixture compressed in the working cylinder of an Otto-cycle (spark ignition) internal-combustion engine even before the charge is ignited by the flame proceeding from the sparking plug. These are the so-called "prelinminary reactions" ("pre-flame" combustion or oxidation), and an exact knowledge of their characteristic development is of great importance for a correct appreciation of the phenomena of engine-knock (detonation), and consequently for its avoidance. Such reactions can be studied either in a working engine cylinder or in a combustion bomb. The first method necessitates a complicated experimental technique, while the second has the disadvantage of enabling only a single reaction to be studied at one time. Consequently, a new series of experiments was inaugurated, conducted in a motored (externally-driven) experimental engine of mixture-compression type, without ignition, the resulting preliminary reactions being detectable and measurable thermometrically.

Large Eddy Simulations of Transverse Combustion Instability in a Multi-Element Injector

DTIC Science & Technology

2016-07-27

plagued the development of liquid rocket engines and remains a large riskin the development and acquisition of new liquid rocket engines. Combustion...simulations to better understand the physics that can lead combustion instability in liquid rocket engines. Simulations of this type are able to...instabilities found in liquid rocket engines are transverse. The motivating of the experiment behind the current work is to subject the CVRC injector

Rocketdyne RBCC Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratckin, G.; Goldman, A.; Ortwerth, P.; Weisberg, S.

1999-01-01

Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

The Production and Evolution of Atomic Oxygen in the Afterglow of Streamer Discharge in Atmospheric Pressure Fuel/Air Mixtures

DTIC Science & Technology

2013-07-02

in streamer discharge afterglow in a variety of fueVair mixtures in order to account for the 0 reaction pathways in transient plasma ignition. It is... plasma ignition (TPI), the use of streamers for ignition in combustion engines, holds great promise for improving performance. TPI has been tested...standard spark gap or arc ignition methods [1-4]. These improvements to combustion allow increasing power and efficiency in existing engines such as

Coal-Fired Rocket Engine

NASA Technical Reports Server (NTRS)

Anderson, Floyd A.

1987-01-01

Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

NASA Astrophysics Data System (ADS)

Sutton, George P.

The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

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

NASA Technical Reports Server (NTRS)

Kleinhenz, Julie; Sarmiento, Charles; Marshall, William

2012-01-01

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.

Russian Rocket Engine Test

NASA Technical Reports Server (NTRS)

1998-01-01

NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

KSC-2011-6891

NASA Image and Video Library

2011-09-10

CAPE CANAVERAL, Fla. – At ignition, flames from the engines begin liftoff of the United Launch Alliance Delta II Heavy rocket carrying NASA’s twin Gravity Recovery and Interior Laboratory (GRAIL) mission off Space Launch Complex 17B on Cape Canaveral Air Force Station. The spacecraft launched at 9:08:52 a.m. EDT Sept. 10. GRAIL-A will separate from the second stage of the rocket at about one hour, 21 minutes after liftoff, followed by GRAIL-B at 90 minutes after launch. The spacecraft are embarking on a three-month journey to reach the moon. GRAIL will fly twin spacecraft in tandem around the moon to precisely measure and map variations in the moon's gravitational field. The mission will provide the most accurate global gravity field to date for any planet, including Earth. This detailed information will reveal differences in the density of the moon's crust and mantle and will help answer fundamental questions about the moon's internal structure, thermal evolution, and history of collisions with asteroids. The aim is to map the moon's gravity field so completely that future moon vehicles can safely navigate anywhere on the moon’s surface. For more information, visit http://www.nasa.gov/grail. Photo credit: NASA/Sandra Joseph and Don Kight

KSC-2011-6889

NASA Image and Video Library

2011-09-10

CAPE CANAVERAL, Fla. – Overlooking the Central Florida coast, engine ignition begins liftoff of the United Launch Alliance Delta II Heavy rocket carrying NASA’s twin Gravity Recovery and Interior Laboratory (GRAIL) mission off Space Launch Complex 17B on Cape Canaveral Air Force Station. The spacecraft launched at 9:08:52 a.m. EDT Sept. 10. GRAIL-A will separate from the second stage of the rocket at about one hour, 21 minutes after liftoff, followed by GRAIL-B at 90 minutes after launch. The spacecraft are embarking on a three-month journey to reach the moon. GRAIL will fly twin spacecraft in tandem around the moon to precisely measure and map variations in the moon's gravitational field. The mission will provide the most accurate global gravity field to date for any planet, including Earth. This detailed information will reveal differences in the density of the moon's crust and mantle and will help answer fundamental questions about the moon's internal structure, thermal evolution, and history of collisions with asteroids. The aim is to map the moon's gravity field so completely that future moon vehicles can safely navigate anywhere on the moon’s surface. For more information, visit http://www.nasa.gov/grail. Photo credit: NASA/Sandra Joseph and Don Kight

Impact of Formaldehyde Addition on Auto-Ignition in Internal-Combustion Engines

NASA Astrophysics Data System (ADS)

Kuwahara, Kazunari; Ando, Hiromitsu; Furutani, Masahiro; Ohta, Yasuhiko

By employing a direct-injection diesel engine equipped with a common-rail type of injection system, by adding formaldehyde (CH2O) to the intake air, and by changing the fuel-injection timing, the compression ratio and the intake-air temperature, a mechanism for CH2O as a fuel additive to affect auto-ignition was discussed. Unlike an HCCI type of engine, the diesel engine can expose an air-fuel mixture only to a limited range of the in-cylinder temperature before the ignition, and can separate low- and high-temperature parts of the mechanism. When low-temperature oxidation starts at a temperature above 900K, there are cases that the CH2O advances the ignition timing. Below 900K, to the contrary, it always retards the timing. It is because, above 900K, a part of the CH2O changes into CO together with H2O2 as an ignition promoter. Below 900K, on the other hand, the CH2O itself acts as an OH radical scavenger against cool-flame reaction, from the beginning of low-temperature oxidation. Then, the engine was modified for its extraordinary function as a gasoline-knocking generator, in order that an effect of CH2O on knocking could be discussed. The CH2O retards the onset of auto-ignition of an end gas. Judging from a large degree of the retardation, the ignition is probably triggered below 900K.

78 FR 77671 - Information Collection Request Submitted to OMB for Review and Approval; Comment Request; NSPS...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-12-24

... Internal Combustion Engines (Renewal) AGENCY: Environmental Protection Agency (EPA). ACTION: Notice... for Stationary Spark Ignition Internal Combustion Engines (40 CFR Part 60, Subpart JJJJ) (Renewal... operators of stationary spark ignition internal combustion engines. Respondent's obligation to respond...

Turbojet-engine Starting and Acceleration

NASA Technical Reports Server (NTRS)

Mc Cafferty, R. J.; Straight, D. M.

1956-01-01

From considerations of safety and reliability in performance of gas-turbine aircraft, it is clear that engine starting and acceleration are of utmost importance. For this reason extensive efforts have been devoted to the investigation of the factors involved in the starting and acceleration of engines. In chapter III it is shown that certain basic combustion requirements must be met before ignition can occur; consequently, the design and operation of an engine must be tailored to provide these basic requirements in the combustion zone of the engine, particularly in the vicinity of the ignition source. It is pointed out in chapter III that ignition by electrical discharges is aided by high pressure, high temperature, low gas velocity and turbulence, gaseous fuel-air mixture, proper mixture strength, and-an optimum spark. duration. The simultaneous achievement of all these requirements in an actual turbojet-engine combustor is obviously impossible, yet any attempt to satisfy as many requirements as possible will result in lower ignition energies, lower-weight ignition systems, and greater reliability. These factors together with size and cost considerations determine the acceptability of the final ignition system. It is further shown in chapter III that the problem of wall quenching affects engine starting. For example, the dimensions of the volume to be burned must be larger than the quenching distance at the lowest pressure and the most adverse fuel-air ratio encountered. This fact affects the design of cross-fire tubes between adjacent combustion chambers in a tubular-combustor turbojet engine. Only two chambers in these engines contain spark plugs; therefore, the flame must propagate through small connecting tubes between the chambers. The quenching studies indicate that if the cross-fire tubes are too narrow the flame will not propagate from one chamber to another. In order to better understand the role of the basic factors in actual engine operation, many investigations have been conducted in single combustors from gas-turbine engines and in full-scale engines in altitude tanks and in flight. The purpose of the present chapter is to discuss the results of such studies and, where possible, to interpret these results qualitatively in terms of the basic requirements reported in chapter III. The discussion parallels the three phases of turbojet engine starting: (1) Ignition of the fuel-air mixture (2) Propagation of flame throughout the combustion zone (3) Acceleration of the engine to operating speed.

Numerical Analysis of the Interaction between Thermo-Fluid Dynamics and Auto-Ignition Reaction in Spark Ignition Engines

NASA Astrophysics Data System (ADS)

Saijyo, Katsuya; Nishiwaki, Kazuie; Yoshihara, Yoshinobu

The CFD simulations were performed integrating the low-temperature oxidation reaction. Analyses were made with respect to the first auto-ignition location in the case of a premixed-charge compression auto-ignition in a laminar flow field and in the case of the auto-ignition in an end gas during an S. I. Engine combustion process. In the latter simulation, the spatially-filtered transport equations were solved to express fluctuating temperatures in a turbulent flow in consideration of strong non-linearity to temperature in the reaction equations. It is suggested that the first auto-ignition location does not always occur at higher-temperature locations and that the difference in the locations of the first auto-ignition depends on the time period during which the local end gas temperature passes through the region of shorter ignition delay, including the NTC region.

Fuel mixture stratification as a method for improving homogeneous charge compression ignition engine operation

DOEpatents

Dec, John E [Livermore, CA; Sjoberg, Carl-Magnus G [Livermore, CA

2006-10-31

A method for slowing the heat-release rate in homogeneous charge compression ignition ("HCCI") engines that allows operation without excessive knock at higher engine loads than are possible with conventional HCCI. This method comprises injecting a fuel charge in a manner that creates a stratified fuel charge in the engine cylinder to provide a range of fuel concentrations in the in-cylinder gases (typically with enough oxygen for complete combustion) using a fuel with two-stage ignition fuel having appropriate cool-flame chemistry so that regions of different fuel concentrations autoignite sequentially.

Contactless electric igniter for vehicle to lower exhaust emission and fuel consumption.

PubMed

Shen, Chih-Lung; Su, Jye-Chau

2014-01-01

An electric igniter for engine/hybrid vehicles is presented. The igniter comprises a flyback converter, a voltage-stacked capacitor, a PIC-based controller, a differential voltage detector, and an ignition coil, of which structure is non-contact type. Since the electric igniter adopts a capacitor to accumulate energy for engine ignition instead of traditional contacttype approach, it enhances the igniting performance of a spark plug effectively. As a result, combustion efficiency is promoted, fuel consumption is saved, and exhaust emission is reduced. The igniter not only is good for fuel efficiency but also can reduce HC and CO emission significantly, which therefore is an environmentally friendly product. The control core of the igniter is implemented on a single chip, which lowers discrete component count, reduces system volume, and increases reliability. In addition, the ignition timing can be programmed so that a timing regulator can be removed from the proposed system, simplifying its structure. To verify the feasibility and functionality of the igniter, key waveforms are measured and real-car experiments are performed as well.

Contactless Electric Igniter for Vehicle to Lower Exhaust Emission and Fuel Consumption

PubMed Central

Su, Jye-Chau

2014-01-01

An electric igniter for engine/hybrid vehicles is presented. The igniter comprises a flyback converter, a voltage-stacked capacitor, a PIC-based controller, a differential voltage detector, and an ignition coil, of which structure is non-contact type. Since the electric igniter adopts a capacitor to accumulate energy for engine ignition instead of traditional contacttype approach, it enhances the igniting performance of a spark plug effectively. As a result, combustion efficiency is promoted, fuel consumption is saved, and exhaust emission is reduced. The igniter not only is good for fuel efficiency but also can reduce HC and CO emission significantly, which therefore is an environmentally friendly product. The control core of the igniter is implemented on a single chip, which lowers discrete component count, reduces system volume, and increases reliability. In addition, the ignition timing can be programmed so that a timing regulator can be removed from the proposed system, simplifying its structure. To verify the feasibility and functionality of the igniter, key waveforms are measured and real-car experiments are performed as well. PMID:24672372

Measuring Model Rocket Engine Thrust Curves

ERIC Educational Resources Information Center

Penn, Kim; Slaton, William V.

2010-01-01

This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

Numerical simulation and validation of SI-CAI hybrid combustion in a CAI/HCCI gasoline engine

NASA Astrophysics Data System (ADS)

Wang, Xinyan; Xie, Hui; Xie, Liyan; Zhang, Lianfang; Li, Le; Chen, Tao; Zhao, Hua

2013-02-01

SI-CAI hybrid combustion, also known as spark-assisted compression ignition (SACI), is a promising concept to extend the operating range of CAI (Controlled Auto-Ignition) and achieve the smooth transition between spark ignition (SI) and CAI in the gasoline engine. In this study, a SI-CAI hybrid combustion model (HCM) has been constructed on the basis of the 3-Zones Extended Coherent Flame Model (ECFM3Z). An ignition model is included to initiate the ECFM3Z calculation and induce the flame propagation. In order to precisely depict the subsequent auto-ignition process of the unburned fuel and air mixture independently after the initiation of flame propagation, the tabulated chemistry concept is adopted to describe the auto-ignition chemistry. The methodology for extracting tabulated parameters from the chemical kinetics calculations is developed so that both cool flame reactions and main auto-ignition combustion can be well captured under a wider range of thermodynamic conditions. The SI-CAI hybrid combustion model (HCM) is then applied in the three-dimensional computational fluid dynamics (3-D CFD) engine simulation. The simulation results are compared with the experimental data obtained from a single cylinder VVA engine. The detailed analysis of the simulations demonstrates that the SI-CAI hybrid combustion process is characterised with the early flame propagation and subsequent multi-site auto-ignition around the main flame front, which is consistent with the optical results reported by other researchers. Besides, the systematic study of the in-cylinder condition reveals the influence mechanism of the early flame propagation on the subsequent auto-ignition.

Microwave-Assisted Ignition for Improved Internal Combustion Engine Efficiency

NASA Astrophysics Data System (ADS)

DeFilippo, Anthony Cesar

The ever-present need for reducing greenhouse gas emissions associated with transportation motivates this investigation of a novel ignition technology for internal combustion engine applications. Advanced engines can achieve higher efficiencies and reduced emissions by operating in regimes with diluted fuel-air mixtures and higher compression ratios, but the range of stable engine operation is constrained by combustion initiation and flame propagation when dilution levels are high. An advanced ignition technology that reliably extends the operating range of internal combustion engines will aid practical implementation of the next generation of high-efficiency engines. This dissertation contributes to next-generation ignition technology advancement by experimentally analyzing a prototype technology as well as developing a numerical model for the chemical processes governing microwave-assisted ignition. The microwave-assisted spark plug under development by Imagineering, Inc. of Japan has previously been shown to expand the stable operating range of gasoline-fueled engines through plasma-assisted combustion, but the factors limiting its operation were not well characterized. The present experimental study has two main goals. The first goal is to investigate the capability of the microwave-assisted spark plug towards expanding the stable operating range of wet-ethanol-fueled engines. The stability range is investigated by examining the coefficient of variation of indicated mean effective pressure as a metric for instability, and indicated specific ethanol consumption as a metric for efficiency. The second goal is to examine the factors affecting the extent to which microwaves enhance ignition processes. The factors impacting microwave enhancement of ignition processes are individually examined, using flame development behavior as a key metric in determining microwave effectiveness. Further development of practical combustion applications implementing microwave-assisted spark technology will benefit from predictive models which include the plasma processes governing the observed combustion enhancement. This dissertation documents the development of a chemical kinetic mechanism for the plasma-assisted combustion processes relevant to microwave-assisted spark ignition. The mechanism includes an existing mechanism for gas-phase methane oxidation, supplemented with electron impact reactions, cation and anion chemical reactions, and reactions involving vibrationally-excited and electronically-excited species. Calculations using the presently-developed numerical model explain experimentally-observed trends, highlighting the relative importance of pressure, temperature, and mixture composition in determining the effectiveness of microwave-assisted ignition enhancement.

A modified Pegasus rocket drops steadily away after release from NASA's B-52B, before accelerating the X-43A over the Pacific Ocean on March 27, 2004

NASA Image and Video Library

2004-03-27

The second X-43A hypersonic research aircraft and its modified Pegasus booster rocket drop away from NASA's B-52B launch aircraft over the Pacific Ocean on March 27, 2004. The mission originated from the NASA Dryden Flight Research Center at Edwards Air Force Base, Calif. Moments later the Pegasus booster ignited to accelerate the X-43A to its intended speed of Mach 7.

A modified Pegasus rocket drops away after release from NASA's B-52B before accelerating the X-43A over a Pacific Ocean test range on Nov. 16, 2004

NASA Image and Video Library

2004-11-16

The third X-43A hypersonic research aircraft and its modified Pegasus booster rocket drop away from NASA's B-52B launch aircraft over the Pacific Ocean on November 16, 2004. The mission originated from the NASA Dryden Flight Research Center at Edwards Air Force Base, California. Moments later the Pegasus booster ignited to accelerate the X-43A to its intended speed of Mach 10.

DESIGN OF A HIGH COMPRESSION, DIRECT INJECTION, SPARK-IGNITION, METHANOL FUELED RESEARCH ENGINE WITH AN INTEGRAL INJECTOR-IGNITION SOURCE INSERT, SAE PAPER 2001-01-3651

EPA Science Inventory

A stratified charge research engine and test stand were designed and built for this work. The primary goal of this project was to evaluate the feasibility of using a removal integral injector ignition source insert which allows a convenient method of charging the relative locat...

Ignition Study on a Rotary-valved Air-breathing Pulse Detonation Engine

NASA Astrophysics Data System (ADS)

Wu, Yuwen; Han, Qixiang; Shen, Yujia; Zhao, Wei

2017-05-01

In the present study, the ignition effect on detonation initiation was investigated in the air-breathing pulse detonation engine. Two kinds of fuel injection and ignition methods were applied. For one method, fuel and air was pre-mixed outside the PDE and then injected into the detonation tube. The droplet sizes of mixtures were measured. An annular cavity was used as the ignition section. For the other method, fuel-air mixtures were mixed inside the PDE, and a pre-combustor was utilized as the ignition source. At firing frequency of 20 Hz, transition to detonation was obtained. Experimental results indicated that the ignition position and initial flame acceleration had important effects on the deflagration-to-detonation transition.

40 CFR 89.124 - Record retention, maintenance, and submission.

Code of Federal Regulations, 2010 CFR

2010-07-01

... construction, including a general description of the origin and buildup of the engine, steps taken to ensure... PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM NEW AND IN-USE NONROAD COMPRESSION-IGNITION ENGINES Emission... manufacturer of any nonroad compression-ignition engine must maintain the following adequately organized...

Nitrous Oxide/Paraffin Hybrid Rocket Engines

NASA Technical Reports Server (NTRS)

Zubrin, Robert; Snyder, Gary

2010-01-01

Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

Performance potential of gas-core and fusion rockets - A mission applications survey.

NASA Technical Reports Server (NTRS)

Fishbach, L. H.; Willis, E. A., Jr.

1971-01-01

This paper reports an evaluation of the performance potential of five nuclear rocket engines for four mission classes. These engines are: the regeneratively cooled gas-core nuclear rocket; the light bulb gas-core nuclear rocket; the space-radiator cooled gas-core nuclear rocket; the fusion rocket; and an advanced solid-core nuclear rocket which is included for comparison. The missions considered are: earth-to-orbit launch; near-earth space missions; close interplanetary missions; and distant interplanetary missions. For each of these missions, the capabilities of each rocket engine type are compared in terms of payload ratio for the earth launch mission or by the initial vehicle mass in earth orbit for space missions (a measure of initial cost). Other factors which might determine the engine choice are discussed. It is shown that a 60 day manned round trip to Mars is conceivable.-

A study of ignition phenomena of bulk metals by radiant heating

NASA Technical Reports Server (NTRS)

Branch, Melvin C.; Abbud-Madrid, A.; Feiereisen, T. J.; Daily, J. W.

1993-01-01

Early research on combustion of metals was motivated by the knowledge of the large heat release and corresponding high temperatures associated with metal-oxygen reactions. The advent of space flight brought about an increased interest in the ignition and combustion of metallic particles as additives in solid rocket propellants. More recently, attention has been given to the flammability properties of bulk, structural metals due to the number of accidental explosions of metal components in high-pressure oxygen systems. The following work represents a preliminary study that is part of a broader research effort aimed at providing further insight into the phenomena of bulk metal combustion by looking at the effects of gravity on the ignition behavior of metals. The scope of this preliminary experimental study includes the use of a non-coherent, continuous radiation ignition source, the measurement of temperature profiles of a variety of metals and a qualitative observation of the ignition phenomena at normal gravity. The specific objectives of the investigation include: (1) a feasibility study of the use of a continuous radiation source for metal ignition; (2) testing and characterization of the ignition behavior of a variety of metals; and (3) building a preliminary experimental database on ignition of metals under normal gravity conditions.

DataRocket: Interactive Visualisation of Data Structures

NASA Astrophysics Data System (ADS)

Parkes, Steve; Ramsay, Craig

2010-08-01

CodeRocket is a software engineering tool that provides cognitive support to the software engineer for reasoning about a method or procedure and for documenting the resulting code [1]. DataRocket is a software engineering tool designed to support visualisation and reasoning about program data structures. DataRocket is part of the CodeRocket family of software tools developed by Rapid Quality Systems [2] a spin-out company from the Space Technology Centre at the University of Dundee. CodeRocket and DataRocket integrate seamlessly with existing architectural design and coding tools and provide extensive documentation with little or no effort on behalf of the software engineer. Comprehensive, abstract, detailed design documentation is available early on in a project so that it can be used for design reviews with project managers and non expert stakeholders. Code and documentation remain fully synchronised even when changes are implemented in the code without reference to the existing documentation. At the end of a project the press of a button suffices to produce the detailed design document. Existing legacy code can be easily imported into CodeRocket and DataRocket to reverse engineer detailed design documentation making legacy code more manageable and adding substantially to its value. This paper introduces CodeRocket. It then explains the rationale for DataRocket and describes the key features of this new tool. Finally the major benefits of DataRocket for different stakeholders are considered.

Robust Rocket Engine Concept

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.

1995-01-01

The potential for a revolutionary step in the durability of reusable rocket engines is made possible by the combination of several emerging technologies. The recent creation and analytical demonstration of life extending (or damage mitigating) control technology enables rapid rocket engine transients with minimum fatigue and creep damage. This technology has been further enhanced by the formulation of very simple but conservative continuum damage models. These new ideas when combined with recent advances in multidisciplinary optimization provide the potential for a large (revolutionary) step in reusable rocket engine durability. This concept has been named the robust rocket engine concept (RREC) and is the basic contribution of this paper. The concept also includes consideration of design innovations to minimize critical point damage.

40 CFR 1045.1 - Does this part apply for my products?

Code of Federal Regulations, 2011 CFR

2011-07-01

... POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Overview... exhaust emissions apply to new, spark-ignition propulsion marine engines beginning with the 2010 model...

40 CFR 1045.1 - Does this part apply for my products?

Code of Federal Regulations, 2012 CFR

2012-07-01

... POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Overview... exhaust emissions apply to new, spark-ignition propulsion marine engines beginning with the 2010 model...

40 CFR 1045.1 - Does this part apply for my products?

Code of Federal Regulations, 2013 CFR

2013-07-01

... POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Overview... exhaust emissions apply to new, spark-ignition propulsion marine engines beginning with the 2010 model...

40 CFR 1045.1 - Does this part apply for my products?

Code of Federal Regulations, 2014 CFR

2014-07-01

... POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Overview... exhaust emissions apply to new, spark-ignition propulsion marine engines beginning with the 2010 model...

40 CFR 1045.1 - Does this part apply for my products?

Code of Federal Regulations, 2010 CFR

2010-07-01

... POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Overview... exhaust emissions apply to new, spark-ignition propulsion marine engines beginning with the 2010 model...

Synthesis and testing of hypergolic ionic liquids for chemical propulsion

NASA Astrophysics Data System (ADS)

Stovbun, S. V.; Shchegolikhin, A. N.; Usachev, S. V.; Khomik, S. V.; Medvedev, S. P.

2017-06-01

Synthesis of new highly energetic ionic liquids (ILs) is described, and their hypergolic ignition properties are tested. The synthesized ILs combine the advantages of conventional rocket propellants with the energy characteristics of acetylene derivatives. To this end, N-alkylated imidazoles (alkyl = ethyl, butyl) have been synthesized and alkylated with propargyl bromide. The desired ionic liquids have been produced by metathesis using Ag dicyanamide. Modified hypergolic drop tests with white fuming nitric acid have been performed for N-ethyl (IL-1) and N-butyl propargylimidazolium (IL-2) ionic liquids. In the modified drop tests, high-speed shadowgraph imaging is used to visualize the process, and the temperature rise due to ignition is monitored with a two-color photodetector. It is shown that the ignition delay is shorter for IL-1 as compared to IL-2. The ignition of IL-1 occurs in two stages, whereas the combustion of IL-2 proceeds smoothly without secondary flashes.

Effect of Boron Clusters on the Ignition Reaction of HNO3 and Dicynanamide-Based Ionic Liquids.

PubMed

Schmidt, Michael W; Gordon, Mark S

2017-10-19

Many ionic liquids containing the dicynamide anion (DCA - , formula N(CN) 2 - ) exhibit hypergolic ignition when exposed to the common oxidizer nitric acid. However, the ignition delay is often about 10 times longer than the desired 5 ms for rocket applications, so that improvements are desired. Experiments in the past decade have suggested both a mechanism for the early reaction steps and also that additives such as decaborane can reduce the ignition delay. The mechanisms for reactions of nitric acid with both DCA - and protonated DCAH are considered here, using accurate wave function methods. Complexation of DCA - or DCAH with borane clusters B 10 H 14 or B 9 H 14 - is found to modify these mechanisms slightly by changing the nature of some of the intermediate saddle points and by small reductions in the reaction barriers.

Russian Rocket Engine Test

NASA Technical Reports Server (NTRS)

1998-01-01

NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust. The test was the first test ever anywhere outside Russia of a Russian designed and built engine.

Concentric traveling ionospheric disturbances triggered by the launch of a SpaceX Falcon 9 rocket

NASA Astrophysics Data System (ADS)

Lin, Charles C. H.; Shen, Ming-Hsueh; Chou, Min-Yang; Chen, Chia-Hung; Yue, Jia; Chen, Po-Cheng; Matsumura, Mitsuru

2017-08-01

We report the first observation of concentric traveling ionospheric disturbances (CTIDs) triggered by the launch of a SpaceX Falcon 9 rocket on 17 January 2016. The rocket-triggered ionospheric disturbances show shock acoustic wave signature in the time rate change (time derivative) of total electron content (TEC), followed by CTIDs in the 8-15 min band-pass filtering of TEC. The CTIDs propagated northward with phase velocity of 241-617 m/s and reached distances more than 1000 km away from the source on the rocket trajectory. The wave characteristics of CTIDs with periods of 10.5-12.7 min and wavelength 200-400 km agree well with the gravity wave dispersion relation. The optimal wave source searching and gravity wave ray tracing technique suggested that the CTIDs have multiple sources which are originated from 38-120 km altitude before and after the ignition of the second-stage rocket, 200 s after the rocket was launched.

40 CFR 1039.107 - What evaporative emission standards and requirements apply?

Code of Federal Regulations, 2010 CFR

2010-07-01

... 1048 that apply to spark-ignition engines, as follows: (a) Follow the steps in 40 CFR 1048.245 to show...-IGNITION ENGINES Emission Standards and Related Requirements § 1039.107 What evaporative emission standards and requirements apply? There are no evaporative emission standards for diesel-fueled engines, or...

Final Rule for Phase 2 Emission Standards for New Nonroad Spark-Ignition Nonhandheld Engines At or Below 19 Kilowatts

EPA Pesticide Factsheets

Emission regulations to control emissions from new nonroad spark-ignition nonhandheld engines at or below 19 kilowatts (25 horsepower). These engines are used principally in lawn and garden equipment in applications such as lawnmowers and garden tractors.

2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND GARAGE (BUILDING 1930) AT LEFT REAR. Looking to west. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

7. Historic aerial photo of rocket engine test facility complex, ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

7. Historic aerial photo of rocket engine test facility complex, June 1962. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-60674. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

STS-90 Day 01 Highlights

NASA Technical Reports Server (NTRS)

1998-01-01

On this first day of the STS-90 mission, the flight crew, Cmdr. Richard A. Searfoss, Pilot Scott D. Altman, and Mission Specialists Richard M. Linnehan, Dafydd Rhys Williams and Kathryn P. Hire, and Payload Specialists Jay C. Buckey and James A. Pawelczyk, can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters. The shuttle's payload bay doors are then opened in anticipation of the 16-day scientific mission. The astronauts then are seen readying the Spacelab module for various experiments.

STS-106 Crew Activity Report/Flight Day 1 Highlights

NASA Technical Reports Server (NTRS)

2000-01-01

On this first day of the STS-106 Atlantis mission, the flight crew, Commander Terrence W. Wilcutt, Pilot Scott D. Altman, and Mission Specialists Daniel C. Burbank, Edward T. Lu, Richard A. Mastracchio, Yuri Ivanovich Malenchenko, and Boris V. Morukov are seen performing pre-launch activities. They are shown sitting around the breakfast table with the traditional cake, suiting-up, and riding out to the launch pad. The final inspection team is seen as they conduct their final check of the space shuttle on the launch complex. Also, included are various panoramic views of the shuttle on the pad. The crew is readied in the 'white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters.

STS-84 Mission Highlights Resource Tape

NASA Technical Reports Server (NTRS)

1997-01-01

The STS-84 mission flight crew, Cmdr. Charles J. Precourt, Pilot Eileen M. Collions, Payload Cmdr. Jean-Francois Clervoy (ESA), Mission Specialists Edward T. Lu, Carlos I. Noriega, Elena V. Kondakova, and Jerry M. Linenger can be seen performing pre-launch activities such as eating the traditional breakfast, crew suit-up, and the ride out to the launch pad. Also, included are various panoramic views of the shuttle on the pad. The crew can be seen being readied in the white room' for their mission. After the closing of the hatch and arm retraction, launch activities are shown including countdown, engine ignition, launch, and the separation of the Solid Rocket Boosters. The rendezvous with the Mir Space Station, along with onboard activities, and landing are included. Also included are shuttle-to-ground transmission between the crew and Mission Control and various earthviews.

KSC-2014-2206

NASA Image and Video Library

2014-04-18

CAPE CANAVERAL, Fla. - A blinding flash of light under the Falcon 9 rocket signals engine ignition and liftoff of the SpaceX-3 mission from Space Launch Complex 40 on Cape Canaveral Air Force Station, sending the Dragon resupply spacecraft on its way to the International Space Station. Launch was during an instantaneous window at 3:25 p.m. EDT. Dragon is making its fourth trip to the space station. The SpaceX-3 mission, carrying almost 2.5 tons of supplies, technology and science experiments, is the third of 12 flights through a $1.6 billion NASA Commercial Resupply Services contract. Dragon's cargo will support more than 150 experiments that will be conducted during the station's Expeditions 39 and 40. For more information, visit http://www.nasa.gov/mission_pages/station/structure/launch/index.html. Photo credit: NASA/Tony Gray

Influence of several factors on ignition lag in a compression-ignition engine

NASA Technical Reports Server (NTRS)

Gerrish, Harold C; Voss, Fred

1932-01-01

This investigation was made to determine the influence of fuel quality, injection advance angle, injection valve-opening pressure, inlet-air pressure, compression ratio, and engine speed on the time lag of auto-ignition of a Diesel fuel oil in a single-cylinder compression-ignition engine as obtained from an analysis of indicator diagrams. Three cam-operated fuel-injection pumps, two pumps cams, and an automatic injection valve with two different nozzles were used. Ignition lag was considered to be the interval between the start of injection of the fuel as determined with a Stroborama and the start of effective combustion as determined from the indicator diagram, the latter being the point where 4.0 x 10(exp-6) pound of fuel had been effectively burned. For this particular engine and fuel it was found that: (1) for a constant start and the same rate of fuel injection up the point of cut-off, a variation in fuel quantity from 1.2 x 10(exp-4) to 4.1 x 10(exp-4) pound per cycle has no appreciable effect on the ignition lag; (2) injection advance angle increases or decreases the lag according to whether density, temperature, or turbulence has the controlling influence; (3) increase in valve-opening pressure slightly increases the lag; and (4) increase of inlet-air pressure, compression ratio, and engine speed reduces the lag.

Free-piston Stirling hydraulic engine and drive system for automobiles

NASA Technical Reports Server (NTRS)

Beremand, D. G.; Slaby, J. G.; Nussle, R. C.; Miao, D.

1982-01-01

The calculated fuel economy for an automotive free piston Stirling hydraulic engine and drive system using a pneumatic accumulator with the fuel economy of both a conventional 1980 spark ignition engine in an X body class vehicle and the estimated fuel economy of a 1984 spark ignition vehicle system are compared. The results show that the free piston Stirling hydraulic system with a two speed transmission has a combined fuel economy nearly twice that of the 1980 spark ignition engine - 21.5 versus 10.9 km/liter (50.7 versus 25.6 mpg) under comparable conditions. The fuel economy improvement over the 1984 spark ignition engine was 81 percent. The fuel economy sensitivity of the Stirling hydraulic system to system weight, number of transmission shifts, accumulator pressure ratio and maximum pressure, auxiliary power requirements, braking energy recovery, and varying vehicle performance requirements are considered. An important finding is that a multispeed transmission is not required. The penalty for a single speed versus a two speed transmission is about a 12 percent drop in combined fuel economy to 19.0 km/liter (44.7 mpg). This is still a 60 percent improvement in combined fuel economy over the projected 1984 spark ignition vehicle.

Mechanism of plasma-assisted ignition for H2 and C1-C5 hydrocarbons

NASA Astrophysics Data System (ADS)

Starikovskiy, Andrey; Aleksandrov, Nikolay

2016-09-01

Nonequilibrium plasma demonstrates ability to control ultra-lean, ultra-fast, low-temperature flames and appears to be an extremely promising technology for a wide range of applications, including aviation GTEs, piston engines, ramjets, scramjets and detonation initiation for pulsed detonation engines. To use nonequilibrium plasma for ignition and combustion in real energetic systems, one must understand the mechanisms of plasma-assisted ignition and combustion and be able to numerically simulate the discharge and combustion processes under various conditions. A new, validated mechanism for high-temperature hydrocarbon plasma assisted combustion was built and allows to qualitatively describe plasma-assisted combustion close and above the self-ignition threshold. The principal mechanisms of plasma-assisted ignition and combustion have been established and validated for a wide range of plasma and gas parameters. These results provide a basis for improving various energy-conversion combustion systems, from automobile to aircraft engines, using nonequilibrium plasma methods.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Mehl, M; Kukkadapu, G; Kumar, K

The use of gasoline in homogeneous charge compression ignition engines (HCCI) and in duel fuel diesel - gasoline engines, has increased the need to understand its compression ignition processes under engine-like conditions. These processes need to be studied under well-controlled conditions in order to quantify low temperature heat release and to provide fundamental validation data for chemical kinetic models. With this in mind, an experimental campaign has been undertaken in a rapid compression machine (RCM) to measure the ignition of gasoline mixtures over a wide range of compression temperatures and for different compression pressures. By measuring the pressure history duringmore » ignition, information on the first stage ignition (when observed) and second stage ignition are captured along with information on the phasing of the heat release. Heat release processes during ignition are important because gasoline is known to exhibit low temperature heat release, intermediate temperature heat release and high temperature heat release. In an HCCI engine, the occurrence of low-temperature and intermediate-temperature heat release can be exploited to obtain higher load operation and has become a topic of much interest for engine researchers. Consequently, it is important to understand these processes under well-controlled conditions. A four-component gasoline surrogate model (including n-heptane, iso-octane, toluene, and 2-pentene) has been developed to simulate real gasolines. An appropriate surrogate mixture of the four components has been developed to simulate the specific gasoline used in the RCM experiments. This chemical kinetic surrogate model was then used to simulate the RCM experimental results for real gasoline. The experimental and modeling results covered ultra-lean to stoichiometric mixtures, compressed temperatures of 640-950 K, and compression pressures of 20 and 40 bar. The agreement between the experiments and model is encouraging in terms of first-stage (when observed) and second-stage ignition delay times and of heat release rate. The experimental and computational results are used to gain insight into low and intermediate temperature processes during gasoline ignition.« less

Qualification Test of the Thiokol TE-M-364-19 Solid-Propellant Rocket Motor (S/N 19006)

DTIC Science & Technology

1977-05-01

cell by a steam ejector operating in series with the ETF exhaust gas compressors. During the motor firing, the motor exhaust gases were used as a...driving gas for the 42-in.-diam, water-cooled, ejector-diffuser system incorporating a 24-deg (half-angle) conical inlet to maintain test cell pressure...after Ignition, sec 0.5 0.6 0.7 Figure 4. Variation of thrust and chamber pressure during motor ignition. - CO Q_ OH LU CO TL cr x CJ 1400

A hybrid rocket engine design for simple low cost sounding rocket use

NASA Astrophysics Data System (ADS)

Grubelich, Mark; Rowland, John; Reese, Larry

1993-06-01

Preliminary test results on a nitrous oxide/HTPB hybrid rocket engine suitable for powering a small sounding rocket to altitudes of 50-100 K/ft are presented. It is concluded that the advantage of the N2O hybrid engine over conventional solid propellant rocket motors is the ability to obtain long burn times with core burning geometries due to the low regression rate of the fuel. Long burn times make it possible to reduce terminal velocity to minimize air drag losses.

Rocket engine exhaust plume diagnostics and health monitoring/management during ground testing

NASA Technical Reports Server (NTRS)

Chenevert, D. J.; Meeks, G. R.; Woods, E. G.; Huseonica, H. F.

1992-01-01

The current status of a rocket exhaust plume diagnostics program sponsored by NASA is reviewed. The near-term objective of the program is to enhance test operation efficiency and to provide for safe cutoff of rocket engines prior to incipient failure, thereby avoiding the destruction of the engine and the test complex and preventing delays in the national space program. NASA programs that will benefit from the nonintrusive remote sensed rocket plume diagnostics and related vehicle health management and nonintrusive measurement program are Space Shuttle Main Engine, National Launch System, National Aero-Space Plane, Space Exploration Initiative, Advanced Solid Rocket Motor, and Space Station Freedom. The role of emission spectrometry and other types of remote sensing in rocket plume diagnostics is discussed.

12. Historic plot plan and drawings index for rocket engine ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

12. Historic plot plan and drawings index for rocket engine test facility, June 28, 1956. NASA GRC drawing number CE-101810. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

9. Historic aerial photo of rocket engine test facility complex, ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

9. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1270. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

10. Historic photo of rendering of rocket engine test facility ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

10. Historic photo of rendering of rocket engine test facility complex, April 28, 1964. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-69472. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

5. Historic photo of scale model of rocket engine test ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

5. Historic photo of scale model of rocket engine test facility, June 18, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45264. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

8. Historic aerial photo of rocket engine test facility complex, ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

8. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1271. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Hydrocarbon-Fueled Rocket Engine Plume Diagnostics: Analytical Developments and Experimental Results

NASA Technical Reports Server (NTRS)

Tejwani, Gopal D.; McVay, Gregory P.; Langford, Lester A.; St. Cyr, William W.

2006-01-01

A viewgraph presentation describing experimental results and analytical developments about plume diagnostics for hydrocarbon-fueled rocket engines is shown. The topics include: 1) SSC Plume Diagnostics Background; 2) Engine Health Monitoring Approach; 3) Rocket Plume Spectroscopy Simulation Code; 4) Spectral Simulation for 10 Atomic Species and for 11 Diatomic Molecular Electronic Bands; 5) "Best" Lines for Plume Diagnostics for Hydrocarbon-Fueled Rocket Engines; 6) Experimental Set Up for the Methane Thruster Test Program and Experimental Results; and 7) Summary and Recommendations.

NASA Tests RS-25 Flight Engine for Space Launch System

NASA Image and Video Library

2017-10-19

Engineers at NASA’s Stennis Space Center in Mississippi on Oct. 19 completed a hot-fire test of RS-25 rocket engine E2063, a flight engine for NASA’s new Space Launch System (SLS) rocket. Engine E2063 is scheduled to help power SLS on its Exploration Mission-2 (EM-2), the first flight of the new rocket to carry humans.

Method for operating a spark-ignition, direct-injection internal combustion engine

DOEpatents

Narayanaswamy, Kushal; Koch, Calvin K.; Najt, Paul M.; Szekely, Jr., Gerald A.; Toner, Joel G.

2015-06-02

A spark-ignition, direct-injection internal combustion engine is coupled to an exhaust aftertreatment system including a three-way catalytic converter upstream of an NH3-SCR catalyst. A method for operating the engine includes operating the engine in a fuel cutoff mode and coincidentally executing a second fuel injection control scheme upon detecting an engine load that permits operation in the fuel cutoff mode.

Fuel Vaporization and Its Effect on Combustion in a High-Speed Compression-Ignition Engine

NASA Technical Reports Server (NTRS)

Rothrock, A M; Waldron, C D

1933-01-01

The tests discussed in this report were conducted to determine whether or not there is appreciable vaporization of the fuel injected into a high-speed compression-ignition engine during the time available for injection and combustion. The effects of injection advance angle and fuel boiling temperature were investigated. The results show that an appreciable amount of the fuel is vaporized during injection even though the temperature and pressure conditions in the engine are not sufficient to cause ignition either during or after injection, and that when the conditions are such as to cause ignition the vaporization process affects the combustion. The results are compared with those of several other investigators in the same field.

Hybrid boosters for future launch vehicles

NASA Astrophysics Data System (ADS)

Dargies, E.; Lo, R. E.

There is a striking similarity in the design of the US Space Transportation System, the European ARI-ANE 5P and the Japanese II-II: they all use a high energy cryogenic core stage along with two large solid propellant rocket boosters (SRB's) in order to provide for a high lift-off thrust level. Prior to last years disasters with Challenger and Titan it was widely held that SRB's were cheap, uncomplicated and safe. Even before the revelation by these accidents of severe safety hazards, shuttle operations demonstrated that the SRB's were by no means as cheap as reusable systems ought to be. In addition, they became known as sources of considerable environmental pollution. In contrast, hybrid rocket propulsion systems offer the following potential advantages: • much higher savety level (TNT equivalent almost zero, shut-down capability in case of ignition failure of one unit, inert against unbonding) • choice of non-toxic propellant combinations • equal or higher specific performance For these reasons, system analysis were carried out to examine hybrids as potential alternative to SRB's. Various analytical tools (mass- and performance models, trajectory simulation etc.) were developed for parametrical studies of hybrid propulsion systems. Special attention was devoted to the well-known primary concern of hybrids: geometrical design of the solid fuel grain and regression rate of the ablating surface. Experimental data were used as input wherever possible. In 1985 first studies were carried out to find possible fields of application for hybrid rocket engines. A mass model and a performance model for hybrid rocket motors were developed, taking into account the peculiarities of hybrid combustion as there are i.e. low regression rate and shifting mixture ratio during operation. After some analytical work was done, hybrids proved to be a promising alternative to SRB's. Compared with solids, hybrids offer many advantages.

Air-Breathing Rocket Engine Test

NASA Technical Reports Server (NTRS)

2000-01-01

This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

Compression-ignition Engine Performance at Altitudes and at Various Air Pressures and Temperatures

NASA Technical Reports Server (NTRS)

Moore, Charles S; Collins, John H

1937-01-01

Engine test results are presented for simulated altitude conditions. A displaced-piston combustion chamber on a 5- by 7-inch single cylinder compression-ignition engine operating at 2,000 r.p.m. was used. Inlet air temperature equivalent to standard altitudes up to 14,000 feet were obtained. Comparison between performance at altitude of the unsupercharged compression-ignition engine compared favorably with the carburetor engine. Analysis of the results for which the inlet air temperature, inlet air pressure, and inlet and exhaust pressure were varied indicates that engine performance cannot be reliably corrected on the basis of inlet air density or weight of air charge. Engine power increases with inlet air pressure and decreases with inlet air temperatures very nearly as straight line relations over a wide range of air-fuel ratios. Correction factors are given.

11. Historic photo of cutaway rendering of rocket engine test ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

11. Historic photo of cutaway rendering of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-74433. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

A Preliminary Study of Flame Propagation in a Spark-ignition Engine

NASA Technical Reports Server (NTRS)

Rothrock, A M; Spencer, R C

1937-01-01

The N.A.C.A. combustion apparatus was altered to operate as a fuel-injection, spark-ignition engine, and a preliminary study was made of the combustion of gasoline-air mixtures at various air-fuel ratios. Air-fuel ratios ranging from 10 to 21.6 were investigated. Records from an optical indicator and films from a high-speed motion-picture camera were the chief sources of data. Schlieren photography was used for an additional study. The results show that the altered combustion apparatus has characteristics similar to those of a conventional spark-ignition engine and should be useful in studying phenomena in spark-ignition engines. The photographs show the flame front to be irregularly shaped rather than uniformly curved. With a theoretically correct mixture the reaction, as indicated by the photographs, is not completed in the flame front but continues for some time after the combustion front has traversed the mixture.

A Study on Homogeneous Charge Compression Ignition Gasoline Engines

NASA Astrophysics Data System (ADS)

Kaneko, Makoto; Morikawa, Koji; Itoh, Jin; Saishu, Youhei

A new engine concept consisting of HCCI combustion for low and midrange loads and spark ignition combustion for high loads was introduced. The timing of the intake valve closing was adjusted to alter the negative valve overlap and effective compression ratio to provide suitable HCCI conditions. The effect of mixture formation on auto-ignition was also investigated using a direct injection engine. As a result, HCCI combustion was achieved with a relatively low compression ratio when the intake air was heated by internal EGR. The resulting combustion was at a high thermal efficiency, comparable to that of modern diesel engines, and produced almost no NOx emissions or smoke. The mixture stratification increased the local A/F concentration, resulting in higher reactivity. A wide range of combustible A/F ratios was used to control the compression ignition timing. Photographs showed that the flame filled the entire chamber during combustion, reducing both emissions and fuel consumption.

Injector element characterization methodology

NASA Technical Reports Server (NTRS)

Cox, George B., Jr.

1988-01-01

Characterization of liquid rocket engine injector elements is an important part of the development process for rocket engine combustion devices. Modern nonintrusive instrumentation for flow velocity and spray droplet size measurement, and automated, computer-controlled test facilities allow rapid, low-cost evaluation of injector element performance and behavior. Application of these methods in rocket engine development, paralleling their use in gas turbine engine development, will reduce rocket engine development cost and risk. The Alternate Turbopump (ATP) Hot Gas Systems (HGS) preburner injector elements were characterized using such methods, and the methodology and some of the results obtained will be shown.

Analysis of an Increase in the Efficiency of a Spark Ignition Engine Through the Application of an Automotive Thermoelectric Generator

NASA Astrophysics Data System (ADS)

Merkisz, Jerzy; Fuc, Pawel; Lijewski, Piotr; Ziolkowski, Andrzej; Galant, Marta; Siedlecki, Maciej

2016-08-01

We have analyzed the increase of the overall efficiency of a spark ignition engine through energy recovery following the application of an automotive thermoelectric generator (ATEG) of our own design. The design of the generator was developed following emission investigations during vehicle driving under city traffic conditions. The measurement points were defined by actual operation conditions (engine speed and load), subsequently reproduced on an engine dynamometer. Both the vehicle used in the on-road tests and the engine dynamometer were fit with the same, downsized spark ignition engine (with high effective power-to-displacement ratio). The thermodynamic parameters of the exhaust gases (temperature and exhaust gas mass flow) were measured on the engine testbed, along with the fuel consumption and electric current generated by the thermoelectric modules. On this basis, the power of the ATEG and its impact on overall engine efficiency were determined.

78 FR 49237 - Airworthiness Directives; the Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-13

... could cause a fuel leak near an ignition source (e.g., hot brakes or engine exhaust nozzle..., which could cause a fuel leak near an ignition source (e.g., hot brakes or engine exhaust nozzle... brakes or engine exhaust nozzle), consequently leading to a fuel-fed fire. (f) Compliance Comply with...

77 FR 52323 - Agency Information Collection Activities; Submission to OMB for Review and Approval; Comment...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-29

...- Ignition Engines (Renewal). ICR Numbers: EPA ICR No. 1695.10, OMB Control No. 2060-0338. ICR Status: This... Engines and Equipment, OMB Control Number 2060-0603) were incorporated into ICR 1695.10. This action was... Requirements for Nonroad Spark-Ignition Engines (Renewal) AGENCY: Environmental Protection Agency (EPA). ACTION...

Fabrication and testing of an enhanced ignition system to reduce cold-start emissions in an ethanol (E85) light-duty truck engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gardiner, D; Mallory, R; Todesco, M

This report describes an experimental investigation of the potential for an enhanced ignition system to lower the cold-start emissions of a light-duty vehicle engine using fuel ethanol (commonly referred to as E85). Plasma jet ignition and conventional inductive ignition were compared for a General Motors 4-cylinder, alcohol-compatible engine. Emission and combustion stability measurements were made over a range of air/fuel ratios and spark timing settings using a steady-state, cold-idle experimental technique in which the engine coolant was maintained at 25 C to simulate cold-running conditions. These tests were aimed at identifying the degree to which calibration strategies such as mixturemore » enleanment and retarded spark timing could lower engine-out hydrocarbon emissions and raise exhaust temperatures, as well as determining how such calibration changes would affect the combustion stability of the engine (as quantified by the coefficient of variation, or COV, of indicated mean effective pressure calculated from successive cylinder pressure measurements). 44 refs., 39 figs.« less

A simplified life-cycle cost comparison of various engines for small helicopter use

NASA Technical Reports Server (NTRS)

Civinskas, K. C.; Fishbach, L. M.

1974-01-01

A ten-year, life-cycle cost comparison is made of the following engines for small helicopter use: (1) simple turboshaft; (2) regenerative turboshaft; (3) compression-ignition reciprocator; (4) spark-ignited rotary; and (5) spark-ignited reciprocator. Based on a simplified analysis and somewhat approximate data, the simple turboshaft engine apparently has the lowest costs for mission times up to just under 2 hours. At 2 hours and above, the regenerative turboshaft appears promising. The reciprocating and rotary engines are less attractive, requiring from 10 percent to 80 percent more aircraft to have the same total payload capability as a given number of turbine powered craft. A nomogram was developed for estimating total costs of engines not covered in this study.

A Comparison of Several Methods of Measuring Ignition Lag in a Compression-ignition Engine

NASA Technical Reports Server (NTRS)

Spanogle, J A

1934-01-01

The ignition lag of a fuel oil in the combustion chamber of a high speed compression-ignition engine was measured by three different methods. The start of injection of the fuel as observed with a Stoborama was taken as the start of the period of ignition lag in all cases. The end of the period of ignition lag was determined by observation of the appearance of incandescence in the combustion chamber, by inspection of a pressure-time card for evidence of pressure rise, and by analysis of the indicator card for evidence of the combustion of a small but definite quantity of fuel. A comparison of the values for ignition lags obtained by these three methods indicates that the appearance of incandescence is later than other evidences of the start of combustion, that visual inspection of a pressure-time diagram gives consistent and usable values with a minimum requirement of time and/or apparatus, and that analysis of the indicator card is not worth while for ignition lag alone.

The hard start phenomena in hypergolic engines. Volume 1: Bibliography

NASA Technical Reports Server (NTRS)

Miron, Y.; Perlee, H. E.

1974-01-01

A bibliography of reports pertaining to the hard start phenomenon in attitude control rocket engines on Apollo spacecraft is presented. Some of the subjects discussed are; (1) combustion of hydrazine, (2) one dimensional theory of liquid fuel rocket combustion, (3) preignition phenomena in small pulsed rocket engines, (4) experimental and theoretical investigation of the fluid dynamics of rocket combustion, and (5) nonequilibrium combustion and nozzle flow in propellant performance.

6. Historic photo of rocket engine test facility Building 202 ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

6. Historic photo of rocket engine test facility Building 202 complex in operation at night, September 12, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45924. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

13. Historic drawing of rocket engine test facility layout, including ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

13. Historic drawing of rocket engine test facility layout, including Buildings 202, 205, 206, and 206A, February 3, 1984. NASA GRC drawing number CF-101539. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

X-43A departs NASA Dryden Flight Research Center for first free-flight attempt.

NASA Technical Reports Server (NTRS)

2001-01-01

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A 'stack' lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing 'scramjet' engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz. The X-43A flights are the first actual flight tests of an aircraft powered by a scramjet engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). Some 90 minutes after takeoff, the Pegasus will launch from a B-52, rocketing the X-43A to Mach 7 at 95,000 feet altitude, or Mach 10 at 105,000 feet altitude. The X-43A will be powered by its revolutionary air-breathing supersonic-combustion ramjet or 'scramjet' engine. The X-43A will then fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments as it descends until it splashes into the Pacific Ocean.

The X-43A/Pegasus combination dropped into the Pacific Ocean after losing control early in the first

NASA Technical Reports Server (NTRS)

2001-01-01

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A 'stack' lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing 'scramjet' engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz. The X-43A flights are the first actual flight tests of an aircraft powered by a scramjet engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). Some 90 minutes after takeoff, the Pegasus will launch from a B-52, rocketing the X-43A to Mach 7 at 95,000 feet altitude, or Mach 10 at 105,000 feet altitude. The X-43A will be powered by its revolutionary air-breathing supersonic-combustion ramjet or 'scramjet' engine. The X-43A will then fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments as it descends until it splashes into the Pacific Ocean.

RS-25 Rocket Engine Test

NASA Image and Video Library

2017-08-09

The 8.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

Fusion Ignition Rocket Engine with Ballistic Ablative Lithium Liner

NASA Technical Reports Server (NTRS)

Martin, Adam; Eskridge, Richard; Fimognari, Peter J., III.

2005-01-01

Thermo-nuclear fusion may be the key to a high Isp, high specific power (low alpha) propulsion system. In a fusion system energy is liberated within, and imparted directly to, the propellant. In principle, this can overcome the performance limitations inherent in systems that require thermal power transfer across a material boundary, and/or multiple power conversion stages (NTR, NEP). A thermo-nuclear propulsion system, which attempts to overcome some of the problems inherent in the ORION concept, is described. A passive tapered liner is launched behind a vehicle, through a hole in a pusher-plate, that is connected to the vehicle by a shock-absorbing mechanism. A dense FRC plasmoid is then accelerated to high velocity (in excess of 1,000 km/s) and shot through the hole into the liner, when it has reached a given point down-range. The kinetic energy of the FRC is converted into thermal and magnetic-field energy, igniting a fusion bum in the magnetically confined plasma. The fusion reaction serves as an ignition source for the liner, which is made out of detonable materials. The energy liberated in this process is converted to thrust by the pusher-plate, as in the classic ORION concept. However with this concept, the vehicle does not carry a magazine of pre-fabricated pulse-units. A magnetic nozzle may also be used, in place of the pusher-plate. Estimates of the conditions needed to achieve a sufficient gain will be presented, along with a description of the driver characteristics. The incorporation of this concept into the propulsion system of a spacecraft will also be discussed.

Over compression influence to the performances of the spark ignition engines

NASA Astrophysics Data System (ADS)

Rakosi, E.; Talif, S. G.; Manolache, G.

2016-08-01

This paper presents the theoretical and experimental results of some procedures used in improving the performances of the automobile spark ignition engines. The study uses direct injection and high over-compression applied to a standard engine. To this purpose, the paper contains both the constructive solutions and the results obtained from the test bed concerning the engine power indices, fuel consumption and exhaust emissions.

Combustion characteristics of the LO2/GCH4 fuel-rich preburners for staged combustion cycle rocket engines

NASA Astrophysics Data System (ADS)

Ono, Fumiei; Tamura, Hiroshi; Sakamoto, Hiroshi; Sasaki, Masaki

1991-09-01

The combustion characteristics of Liquid Oxygen (LO2)/Gaseous Methane (GCH4) fuel rich preburners were experimentally studied using subscale hardware. Three types of preburners with coaxial type propellant injection elements were designed and fabricated, and were used for hot fire testing. LO2 was used as oxidizer, and GCH4 at room temperature was used as fuel. The tests were conducted at chamber pressures ranging from 6.7 to 11.9 M Pa, and oxidizer to fuel ratios ranged from 0.16 to 0.42. The test results, which include combustion gas temperature T(sub c), characteristic velocity C(sup *) and soot adhesion data, are presented. The T(sub c) efficiency and the C(sup *) efficiency were found to be a function of oxidizer to fuel ratio and chamber pressure. These efficiencies are correlated by an empirical correlation parameter which accounts for the effects of oxidizer to fuel ratio and chamber pressure. The exhaust plumes were colorless and transparent under all tests conditions. There was some soot adhesion to the chamber wall, but no soot adhesion was observed on the main injector simulator orifices. Higher temperature igniter gas was required to ignite the main propellants of the preburner compared with that of the LO2/Gaseous Hydrogen (GH2) propellants combination.

LITTLE JOE 2 - LAUNCH VEHICLES - VA

NASA Image and Video Library

1961-04-13

G61-00030 (4 Nov. 1959) --- Launch of Little Joe-2 from Wallops Island carrying Mercury spacecraft test article. The suborbital test flight of the Mercury capsule was to test the escape system. Vehicle functioned perfectly, but escape rocket ignited several seconds too late. Photo credit: NASA

40 CFR 94.1 - Applicability.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF EMISSIONS FROM MARINE COMPRESSION-IGNITION ENGINES General Provisions for Emission Regulations for Compression-Ignition Marine Engines § 94.1 Applicability. (a) Except as noted in paragraphs (b) and (c) of...