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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Program For Simulation Of Trajectories And Events

Universal Simulation Executive (USE) program accelerates and eases generation of application programs for numerical simulation of continuous trajectories interrupted by or containing discrete events. Developed for simulation of multiple spacecraft trajectories with events as one spacecraft crossing the equator, two spacecraft meeting or parting, or firing rocket engine. USE also simulates operation of chemical batch processing factory. Written in Ada.

Gottlieb, Robert G.↗
NASA NTRS2004

Flame Retardant Epoxy Resins

As part of a program to develop fire resistant exterior composite structures for future subsonic commercial aircraft, flame retardant epoxy resins are under investigation. Epoxies and their curing agents (aromatic diamines) containing phosphorus were synthesized and used to prepare epoxy formulations. Phosphorus was incorporated within the backbone of the epoxy resin and not used as an additive. The resulting cured epoxies were characterized by thermogravimetric analysis, propane torch test, elemental analysis and microscale combustion calorimetry. Several formulations showed excellent flame retardation with phosphorous contents as low as 1.5% by weight. The fracture toughness of plaques of several cured formulations was determined on single-edge notched bend specimens. The chemistry and properties of these new epoxy formulations are discussed.

Thompson, C. M.↗
NASA NTRSDate not supplied

Development and Testing of Pulse Guns for Combustion Instability Testing

Combustion stability requirements of a liquid propellant rocket engine are most commonly verified by evaluating the oscillatory pressure or vibration characteristics following an artificial disturbance of the combustion in the combustion chamber. Two types of devices are most commonly used for dynamic stability rating – nondirectional bombs and pulse guns. Bombs have historically been used most often, especially for the large engines developed around the Apollo era. However, bombs in recent years have become much more expensive and difficult to procure, and their transportation and handling requirements have become more demanding. Pulse guns, with inherently less expensive materials, and far less restrictive handling and shipping requirements, have become an attractive alternative. The NASA Marshall Space Flight Center has developed a pulse gun for use during future development programs at the center, and as a design reference for interested commercial users. The objective of this paper is to present the different variants of the pulse gun and characterize performance over a range of parameters. The parameters include the breech diameter and corresponding volume, barrel length, amount of gun powder used, the method of containing the gunpowder within the breech, and the pressure rating of the burst disk. In total, 44 tests were conducted with the pulse gun firing into a test chamber pressurized to 2300 psig with gaseous nitrogen. The best experimental results – including the highest and most consistent peak pressure amplitudes recorded by dynamic pressure sensors in the test chamber – were from a configuration consisting of a 0.40” inner-diameter breech, 15-16 grains of gunpowder wrapped in cigarette paper, and a 24,000 psid burst disk. With this configuration, the zero-to-peak overpressures recorded in the test chamber ranged from about 37% to 58% of the 2300 psig mean pressure, which is adequate for typical combustion stability rating. To complement the data analysis, the acoustics of the pressure waves as they propagated through the pulse gun and into the test chamber were modeled. A more comprehensive final report of this pulse gun development program is planned to be available in the near future.

Pulse Gun↗
NASA NTRS2011

Comparison of DSMC and CFD Solutions of Fire II Including Radiative Heating

The ability to compute rarefied, ionized hypersonic flows is becoming more important as missions such as Earth reentry, landing high mass payloads on Mars, and the exploration of the outer planets and their satellites are being considered. These flows may also contain significant radiative heating. To prepare for these missions, NASA is developing the capability to simulate rarefied, ionized flows and to then calculate the resulting radiative heating to the vehicle's surface. In this study, the DSMC codes DAC and DS2V are used to obtain charge-neutral ionization solutions. NASA s direct simulation Monte Carlo code DAC is currently being updated to include the ability to simulate charge-neutral ionized flows, take advantage of the recently introduced Quantum-Kinetic chemistry model, and to include electronic energy levels as an additional internal energy mode. The Fire II flight test is used in this study to assess these new capabilities. The 1634 second data point was chosen for comparisons to be made in order to include comparisons to computational fluid dynamics solutions. The Knudsen number at this point in time is such that the DSMC simulations are still tractable and the CFD computations are at the edge of what is considered valid. It is shown that there can be quite a bit of variability in the vibrational temperature inferred from DSMC solutions and that, from how radiative heating is computed, the electronic temperature is much better suited for radiative calculations. To include the radiative portion of heating, the flow-field solutions are post-processed by the non-equilibrium radiation code HARA. Acceptable agreement between CFD and DSMC flow field solutions is demonstrated and the progress of the updates to DAC, along with an appropriate radiative heating solution, are discussed. In addition, future plans to generate more high fidelity radiative heat transfer solutions are discussed.

Liechty, Derek S.↗
NASA NTRS2009

Lessons Learned Entry: Hypergolic Propellant Related Spills and Fires

The attached report is a compilation of all credible, unintentional hypergolic fluid related spills, fires, and explosions from the Apollo Program, the Space Shuttle Program, Titan Program, and a few other programs. Spill sites include the following government facilities: KSC, JSC, WSTF, VAFB, CCAFS, EAFB, Little Rock AFB, and McConnell AFB. The root causes and consequences of the incidents contained in this document vary drastically; however, certain "themes" can be deduced and utilized for future hypergolic propellant handling. Some of those common "themes" are summarized below: (1) Improper configuration control and complacency can lead to being falsely comfortable with a system (2) Communication breakdown can escalate an incident to a level where injuries occur and/or hardware is damaged (3) Improper propulsion system and ground support system designs can destine a system for failure (4) Improper training of technicians, engineers, and safety personnel can put lives in danger (5) Improper PPE, spill protection, and staging of fire extinguishing equipment can result in unnecessary injuries or hardware damage if an incident occurs (6) Improper procedural oversight, development, and adherence to the procedure can be detrimental and quickly lead to an undesirable incident (7) Improper local cleanliness or compatibility can result in fires or explosions The items listed above are only a short list of the issues that should be recognized prior to handling of hypergolic fluids or processing of vehicles containing hypergolic propellants. The summary of incidents in this report is intended to cover many more issues than those listed above that have been found during nearly the entire spectrum. of hypergolic propellant and/or vehicle processing.

Nufer, Brian↗
NASA NTRS2000

Lead-Free Double-Base Propellant for the 2.75 Inch Rocket Motor

The current MK 66 2.75 inch Rocket Motor double-base propellant contains the lead-based ballistic modifier LC-12-15 to achieve the desired plateau and mesa burning rate characteristics. The use of lead compounds poses a concern for the environment and for personal safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacture, rocket motor firing, and disposal. Copper beta-resorcylate (copper 2,4-di-hydroxy-benzoate) was successfully used in propellant as a simple modifier in the mid 1970's. This and other compounds have also been mixed with lead salts to obtain more beneficial ballistic results. Synthesized complexes of lead and copper compounds soon replaced the mixtures. The complexes incorporate the lead, copper lack of organic liquids, which allows for easier propellant processing. About ten years ago, the Indian Head Division, Naval Surface Warfare Center (NSWC), initiated an effort to develop a lead-free propellant for use in missile systems. Several lead-free propellant candidate formulations were developed. About five years ago, NSWC, in conjunction with Alliant Techsystems, Radford Army Ammunition Plant, continued ballistic modifier investigations. A four component ballistic modifier system without lead for double-base propellants that provide adequate plateau and mesa burn rate characteristics was developed and patented. The ballistic modifier's system contains bismuth subsalicylate, 1.5 percent; copper salicylate, 1.0 percent, copper stannate, 0.77 percent; and carbon black, 0.1 percent. Action time and impulse data obtained through multiple static firings indicate that the new lead-free double-base propellant, while not a match for NOSIH-AA-2, will be a very suitable replacement in the 2.75 inch Rocket Motor. Accelerated aging of the double-base propellant containing the lead-free ballistic modifier showed that it had a much higher rate of stabilizer depletion than the AA-2. A comprehensive study showed that an increased rate of stabilizer depletion occurred in propellants containing monobasic copper salicylate. The study also showed that propellants containing a mixture of bismuth subsalicylate and copper salicylate, had only about one-half the stabilizer depletion rate than those with copper salicylate alone. The copper salicylate catalyzes the decomposition of nitroglycerin, which triggers a chain of events leading to the increased rate of stabilizer depletion. A program has been initiated to coat the ballistic modifier, thus isolating it from the nitroglycerin.

Magill, B. T.↗
NASA NTRS2009

Polyimide/Glass Composite High-Temperature Insulation

Lightweight composites of RP46 polyimide and glass fibers have been found to be useful as extraordinarily fire-resistant electrical-insulation materials. RP46 is a polyimide of the polymerization of monomeric reactants (PMR) type, developed by NASA Langley Research Center. RP46 has properties that make it attractive for use in electrical insulation at high temperatures. These properties include high-temperature resistance, low relative permittivity, low dissipation factor, outstanding mechanical properties, and excellent resistance to moisture and chemicals. Moreover, RP46 contains no halogen or other toxic materials and when burned it does not produce toxic fume or gaseous materials. The U. S. Navy has been seeking lightweight, high-temperature-resistant electrical-insulation materials in a program directed toward reducing fire hazards and weights in ship electrical systems. To satisfy the requirements of this program, an electrical-insulation material must withstand a 3-hour gas-flame test at 1,600 F (about 871 C). Prior to the development reported here, RP46 was rated for use at temperatures from -150 to +700 F (about -101 to 371 C), and no polymeric product - not even RP46 - was expected to withstand the Navy 3-hour gas-flame test.

Pater, Ruth H.↗
NASA NTRS1972

Pulsed plasma solid propellant microthruster for the synchronous meteorological satellite. Task 4: Engineering model fabrication and test report

Two flight prototype solid propellant pulsed plasma microthruster propulsion systems for the SMS satellite were fabricated, assembled and tested. The propulsion system is a completely self contained system requiring only three electrical inputs to operate: a 29.4 volt power source, a 28 volt enable signal and a 50 millsec long command fire signal that can be applied at any rate from 50 ppm to 110 ppm. The thrust level can be varied over a range 2.2 to 1 at constant impulse bit amplitude. By controlling the duration of the 28 volt enable either steady state thrust or a series of discrete impulse bits can be generated. A new technique of capacitor charging was implemented to reduce high voltage stress on energy storage capacitors.

Guman, W. J.↗
NASA NTRS2000

BOREAS TGB-5 CO2, CH4 and CO Chamber Flux Data Over the NSA

The BOReal Ecosystem-Atmosphere Study Trace Gas Biogeochemistry (BOREAS TGB-5) team collected a variety of trace gas concentration and flux measurements at several NSA sites. This data set contains carbon dioxide (CO2), methane (CH4), and carbon monoxide (CO) chamber flux measurements conducted in 1994 at upland forest sites that experienced stand-replacement fires. These measurements were acquired to understand the impact of fires on soil biogeochemistry and related changes in trace gas exchange in boreal forest soils. Relevant ancillary data, including data concerning the soil temperature, solar irradiance, and information from nearby un-burned control sites, are included to provide a basis for modeling the regional impacts of fire and climate changes on trace gas biogeochemistry. The data are provided in tabular ASCII files.

Burke, Roger↗
NASA NTRS1993

Automated cloud classification with a fuzzy logic expert system

An unresolved problem in current cloud retrieval algorithms concerns the analysis of scenes containing overlapping cloud layers. Cloud parameterizations are very important both in global climate models and in studies of the Earth's radiation budget. Most cloud retrieval schemes, such as the bispectral method used by the International Satellite Cloud Climatology Project (ISCCP), have no way of determining whether overlapping cloud layers exist in any group of satellite pixels. One promising method uses fuzzy logic to determine whether mixed cloud and/or surface types exist within a group of pixels, such as cirrus, land, and water, or cirrus and stratus. When two or more class types are present, fuzzy logic uses membership values to assign the group of pixels partially to the different class types. The strength of fuzzy logic lies in its ability to work with patterns that may include more than one class, facilitating greater information extraction from satellite radiometric data. The development of the fuzzy logic rule-based expert system involves training the fuzzy classifier with spectral and textural features calculated from accurately labeled 32x32 regions of Advanced Very High Resolution Radiometer (AVHRR) 1.1-km data. The spectral data consists of AVHRR channels 1 (0.55-0.68 mu m), 2 (0.725-1.1 mu m), 3 (3.55-3.93 mu m), 4 (10.5-11.5 mu m), and 5 (11.5-12.5 mu m), which include visible, near-infrared, and infrared window regions. The textural features are based on the gray level difference vector (GLDV) method. A sophisticated new interactive visual image Classification System (IVICS) is used to label samples chosen from scenes collected during the FIRE IFO II. The training samples are chosen from predefined classes, chosen to be ocean, land, unbroken stratiform, broken stratiform, and cirrus. The November 28, 1991 NOAA overpasses contain complex multilevel cloud situations ideal for training and validating the fuzzy logic expert system.

Tovinkere, Vasanth↗
NASA NTRS1995

Plume Diagnostics of the RSRM Static Firings for the Pressure Perturbation Studies

During the STS-54 launch (RSRM-29), the right hand solid rocket motor experienced a 13.9 psi chamber pressure perturbation at 67 seconds into the motor operation. This pressure augmentation equated to a thrust change of 51 klb. Concerns were raised regarding the adverse effects of this thrust imbalance on the shuttle system and the overall thrust into the external tank structural elements. Pressure perturbations have been observed in solid rocket motors due to expulsion of igniter or insulation materials; the motor thrust during such events drop abruptly before rising. However, the RSRM motors do not exhibit such behavior during the large chamber pressure perturbation events. Several scenarios were investigated to explain these pressure perturbations in the RSRM motors based on a fault tree developed after STS-54. Of these, the expulsion of the slag accumulated in the submerged nozzle region appeared to be the most plausible scenario to explain the observations. Slag is a natural combustion product of aluminized solid rocket motors. The RSRM propellant contains 16% by weight of aluminum. Any ejection of this slag mass during nozzle vectoring or other side loads on the motor will result in the chamber pressure perturbation. Two RSRM static firings were instrumented extensively to further understand the slag expulsion phenomenon in the RSRM and the associated pressure perturbations.

Mathias, Edward C.↗
NASA NTRS1985

To catch a comet: Technical overview of CAN DO G-324

The primary photographing of aboard STS 61-E. INTRODUCTION objective of the C.E. Williams Middle School Get Away Special "CAN DO" is the Comet Halley. As it is currently scheduled, G-324 will fly in March of 1986 The project will involve middle school students, grades 6 through 8, in the study and interpretation of astronomical photographs and techniques. -The-objectives of these studies are detailed in another paper to be presented at this symposium. 6-324 is contained in a 5 cubic foot G.A.S. Canister with an opening door and pyrex window for photography. It will be pressurized with one atmosphere of dry nitrogen. Three 35mm still cameras with 250 exposure film backs and different focal length lenses will be fired by a combination of automatic timer and an active comet detector. A lightweight 35mm movie camera will shoot single exposures at about 4 minute intervals to give an overlapping "skymap" of the mission. The fifth camera is a solid state television camera specially constructed for detection of the comet by the microprocessor. In addition to the photographic goals of CAN DO, the canister also contains a separate package of middle school student designed passive experiments. These-projects will be described in detail in the accompanying paper on educational goals.

Thomas J. O'Brien↗
NASA NTRS1993

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Niiya, Karen E.↗
NASA NTRS2007

Ballistic Impact Response of Kevlar 49 and Zylon under Conditions Representing Jet Engine Fan Containment

A ballistic impact test program was conducted to provide validation data for the development of numerical models of blade out events in fabric containment systems. The impact response of two different fiber materials - Kevlar 49 (E.I. DuPont Nemours and Company) and Zylon AS (Toyobo Co., Ltd.) was studied by firing metal projectiles into dry woven fabric specimens using a gas gun. The shape, mass, orientation and velocity of the projectile were varied and recorded. In most cases the tests were designed such that the projectile would perforate the specimen, allowing measurement of the energy absorbed by the fabric. The results for both Zylon and Kevlar presented here represent a useful set of data for the purposes of establishing and validating numerical models for predicting the response of fabrics under conditions simulating those of a jet engine blade release situation. In addition some useful empirical observations were made regarding the effects of projectile orientation and the relative performance of the different materials.

Pereira, J. Michael↗
NASA NTRS2002

Lightweight, Actively Cooled Ceramic Matrix Composite Thrustcells Successfully Tested in Rocket Combustion Lab

In a joint effort between the NASA Glenn Research Center and the NASA Marshall Space Flight Center, regeneratively cooled ceramic matrix composite (CMC) thrustcells were developed and successfully tested in Glenn's Rocket Combustion Lab. Cooled CMC's offer the potential for substantial weight savings over more traditional metallic parts. Two CMC concepts were investigated. In the first of these concepts, an innovative processing approach utilized by Hyper-Therm, Inc., allowed woven CMC coolant containment tubes to be incorporated into the complex thruster design. In this unique design, the coolant passages had varying cross-sectional shapes but maintained a constant cross-sectional area along the length of the thruster. These thrusters were silicon carbide matrix composites reinforced with silicon carbide fibers. The second concept, which was supplied by Ceramic Composites, Inc., utilized copper cooling coils surrounding a carbon-fiber-reinforced carbon matrix composite. In this design, a protective gradient coating was applied to the inner thruster wall. Ceramic Composites, Inc.'s, method of incorporating the coating into the fiber and matrix eliminated the spallation problem often observed with thermal barrier coatings during hotfire testing. The focus of the testing effort was on screening the CMC material's capabilities as well as evaluating the performance of the thermal barrier or fiber-matrix interfacial coatings. Both concepts were hot-fire tested in gaseous O2/H2 environments. The test matrix included oxygen-to-fuel ratios ranging from 1.5 to 7 with chamber pressures to 400 psi. Steady-state internal wall temperatures in excess of 4300 F were measured in situ for successful 30-sec test runs. Photograph of actively cooled composite thrustcell fabricated by Hyper-Therm is shown. The thrustcell is a silicon-carbide-fiber-reinforced silicon carbide matrix composite with woven cooling channels. The matrix is formed via chemical vapor infiltration. Photograph of hot-fire test of an actively cooled carbon-fiber-reinforced carbon matrix composite thrustcell is also shown. This composite thrustcell, which was fabricated by CCI, Inc., was wound with copper cooling coils to contain the water coolant. The tests were run with oxygen fuel ratios up to seven with chamber pressures of 200 psia.

Jaskowiak, Martha H.↗
NASA NTRS1971

Technology for the cities

The activities of an organization for making the technology derived from NASA programs available for use by urban communities are discussed. Examples of technology utilization accomplished by the organization are: (1) self-contained breathing apparatus for firemen, (2) short range communication systems, (3) pavement marking, (4) protective clothing for firemen, (5) underground pipe and conduit locator, and (6) automatic fire hose flow regulator. The manner in which urban problems are identified and the designation of applicable technical developments are described.

Source record↗
NASA NTRS1995

Fundamental phenomena on fuel decomposition and boundary-layer combustion processes with applications to hybrid rocket motors

The experimental study on the fundamental processes involved in fuel decomposition and boundary-layer combustion in hybrid rocket motors is continuously being conducted at the High Pressure Combustion Laboratory of The Pennsylvania State University. This research will provide a useful engineering technology base in the development of hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high-pressure, 2-D slab motor has been designed, manufactured, and utilized for conducting seven test firings using HTPB fuel processed at PSU. A total of 20 fuel slabs have been received from the Mcdonnell Douglas Aerospace Corporation. Ten of these fuel slabs contain an array of fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. Diagnostic instrumentation used in the test include high-frequency pressure transducers for measuring static and dynamic motor pressures and fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. The ultrasonic pulse-echo technique as well as a real-time x-ray radiography system have been used to obtain independent measurements of instantaneous solid fuel regression rates.

Kuo, Kenneth K.↗
NASA NTRS1999

Kinetics and Structure of Superagglomerates Produced by Silane and Acetylene

The evolution of smoke in a laminar diffusion flame involves several steps. The first step is particle inception/nucleation in the high-temperature fuel-rich region of the flame followed by surface growth and coagulation/coalescence of the small particles. As the primary spheres grow in size and lose hydrogen, the colliding particles no longer coalesce but retain their identity as a cluster of primary spheres, termed an agglomerate. Finally, in the upper portion of the flame, the particles enter an oxidizing environment which may lead to partial or complete burnout of the agglomerates. Currently there is no quantitative model for describing the growth of smoke agglomerates up to superagglomerates with an overall dimension of 10 microns and greater. Such particles are produced during the burning of acetylene and fuels containing benzene rings such as toluene and polystyrene. In the case of polystyrene, smoke agglomerates in excess of 1 mm have been observed "raining" out from large fires. Evidence of the formation of superagglomerates in a laminar acetylene/air diffusion flame has been recently reported. Acetylene was chosen as the fuel since the particulate loading in acetylene/air diffusion flames is very high. Photographs were obtained by Sorensen using a microsecond xenon lamp of the "stream" of soot just above the flame. For low flow rates of acetylene, only submicrometer soot clusters are produced and they give rise to the homogeneous appearance of the soot stream. When the flow rate is increased to 1.7 cu cm/s, soot clusters up to 10 microns are formed and they are responsible for the graininess and at a flow rate of 3.4 cu cm/s, a web of interconnected clusters as large as the width of the flame is seen. This interconnecting web of superagglomerates is described as a gel state by Sorensen et al (1998). This is the first observation of a gel for a gas phase system. It was observed that this gel state immediately breaks up into agglomerates due to buoyancy induced turbulence and gravitational sedimentation.

Mulholland, G. W.↗