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.

The University of Arizona program in solid propellants

NASA Technical Reports Server (NTRS)

Ramohalli, Kumar

1989-01-01

The University of Arizona program is aimed at introducing scientific rigor to the predictability and quality assurance of composite solid propellants. Two separate approaches are followed: to use the modern analytical techniques to experimentally study carefully controlled propellant batches to discern trends in mixing, casting, and cure; and to examine a vast bank of data, that has fairly detailed information on the ingredients, processing, and rocket firing results. The experimental and analytical work is described briefly. The principle findings were that: (1) pre- (dry) blending of the coarse and fine ammonium perchlorate can significantly improve the uniformity of mixing; (2) the Fourier transformed IR spectra of the uncured and cured polymer have valuable data on the state of the fuel; (3) there are considerable non-uniformities in the propellant slurry composition near the solid surfaces (blades, walls) compared to the bulk slurry; and (4) in situ measurements of slurry viscosity continuously during mixing can give a good indication of the state of the slurry. Several important observations in the study of the data bank are discussed.

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.

New high energetic composite propellants for space applications: refrigerated solid propellant

NASA Astrophysics Data System (ADS)

Franson, C.; Orlandi, O.; Perut, C.; Fouin, G.; Chauveau, C.; Gökalp, I.; Calabro, M.

2009-09-01

Cryogenic solid propellants (CSP) are a new kind of chemical propellants that use frozen products to ensure the mechanical resistance of the grain. The objective is to combine the high performances of liquid propulsion and the simplicity of solid propulsion. The CSP concept has few disadvantages. Storability is limited by the need of permanent cooling between motor loading and firing. It needs insulations that increase the dry mass. It is possible to limit significantly these drawbacks by using a cooling temperature near the ambient one. It will permit not to change the motor materials and to minimize the supplementary dry mass due to insulator. The designation "Refrigerated Solid Propellant" (RPS) is in that case more appropriate as "Cryogenic Solid Propellant." SNPE Matériaux Energétiques is developing new concept of composition e e with cooling temperature as near the ambient temperature as possible. They are homogeneous and the main ingredients are hydrogen peroxide, polymer and metal or metal hydride, they are called "HydroxalaneTM." This concept allows reaching a high energy level. The expected specific impulse is between 355 and 375 s against 315 s for hydroxyl-terminated polybutadiene (HTPB) / ammonium perchlorate (AP) / Al composition. However, the density is lower than for current propellants, between 1377 and 1462 kg/m3 compared to around 1800 kg/m3 . This is an handicap only for volume-limited application. Works have been carried out at laboratory scale to define the quality of the raw materials and the manufacturing process to realize sample and small grain in a safer manner. To assess the process, a small grain with an internal bore had been realized with a composition based on aluminum and water. This grain had shown very good quality, without any defect, and good bonding properties on the insulator.

Space Shuttle Projects

NASA Image and Video Library

1987-05-27

This photograph is a long shot view of a full scale solid rocket motor (SRM) for the solid rocket booster (SRB) being test fired at Morton Thiokol's Wasatch Operations in Utah. The twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the SRM's were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. Under the direction of the Marshall Space Flight Center, the SRM's are provided by the Morton Thiokol Corporation.

The Initial Atmospheric Transport (IAT) Code: Description and Validation

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

Morrow, Charles W.; Bartel, Timothy James

The Initial Atmospheric Transport (IAT) computer code was developed at Sandia National Laboratories as part of their nuclear launch accident consequences analysis suite of computer codes. The purpose of IAT is to predict the initial puff/plume rise resulting from either a solid rocket propellant or liquid rocket fuel fire. The code generates initial conditions for subsequent atmospheric transport calculations. The Initial Atmospheric Transfer (IAT) code has been compared to two data sets which are appropriate to the design space of space launch accident analyses. The primary model uncertainties are the entrainment coefficients for the extended Taylor model. The Titan 34Dmore » accident (1986) was used to calibrate these entrainment settings for a prototypic liquid propellant accident while the recent Johns Hopkins University Applied Physics Laboratory (JHU/APL, or simply APL) large propellant block tests (2012) were used to calibrate the entrainment settings for prototypic solid propellant accidents. North American Meteorology (NAM )formatted weather data profiles are used by IAT to determine the local buoyancy force balance. The IAT comparisons for the APL solid propellant tests illustrate the sensitivity of the plume elevation to the weather profiles; that is, the weather profile is a dominant factor in determining the plume elevation. The IAT code performed remarkably well and is considered validated for neutral weather conditions.« less

Space Shuttle Projects

NASA Image and Video Library

1979-07-13

This is a photograph of the solid rocket booster's (SRB's) Qualification Motor-1 (QM-1) being prepared for a static firing in a test stand at the Morton Thiokol Test Site in Wasatch, Utah, showing the aft end of the booster. The twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. Under the direction of the Marshall Space Flight Center, the SRM's are provided by the Morton Thiokol Corporation.

Real-Time Inhibitor Recession Measurements in the Space Shuttle Reusable Solid Rocket Motors

NASA Technical Reports Server (NTRS)

McWhorter, Bruce B.; Ewing, Mark E.; McCool, Alex (Technical Monitor)

2001-01-01

Real-time char line recession measurements were made on propellant inhibitors of the Space Shuttle Reusable Solid Rocket Motor (RSRM). The RSRM FSM-8 static test motor propellant inhibitors (composed of a rubber insulation material) were successfully instrumented with eroding potentiometers and thermocouples. The data was used to establish inhibitor recession versus time relationships. Normally, pre-fire and post-fire insulation thickness measurements establish the thermal performance of an ablating insulation material. However, post-fire inhibitor decomposition and recession measurements are complicated by the fact that most of the inhibitor is back during motor operation. It is therefore a difficult task to evaluate the thermal protection offered by the inhibitor material. Real-time measurements would help this task. The instrumentation program for this static test motor marks the first time that real-time inhibitors. This report presents that data for the center and aft field joint forward facing inhibitors. The data was primarily used to measure char line recession of the forward face of the inhibitors which provides inhibitor thickness reduction versus time data. The data was also used to estimate the inhibitor height versus time relationship during motor operation.

Modeling of vortex generated sound in solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Flandro, G. A.

1980-01-01

There is considerable evidence based on both full scale firings and cold flow simulations that hydrodynamically unstable shear flows in solid propellant rocket motors can lead to acoustic pressure fluctuations of significant amplitude. Although a comprehensive theoretical understanding of this problem does not yet exist, procedures were explored for generating useful analytical models describing the vortex shedding phenomenon and the mechanisms of coupling to the acoustic field in a rocket combustion chamber. Since combustion stability prediction procedures cannot be successful without incorporation of all acoustic gains and losses, it is clear that a vortex driving model comparable in quality to the analytical models currently employed to represent linear combustion instability must be formulated.

Solid propellant rocket motor internal ballistics performance variation analysis, phase 3

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.; Murph, J. E.; Adams, G. W., Jr.

1977-01-01

Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

Low-Cost Propellant Launch From a Tethered Balloon

NASA Technical Reports Server (NTRS)

Wilcox, Brian

2006-01-01

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

Results of the NASA/MSFC FA-23 plume technology test program performed in the NASA/Ames unitary wind tunnels

NASA Technical Reports Server (NTRS)

Hendershot, K. C.

1977-01-01

A 2.25% scale model of the space shuttle external tank and solid rocket boosters was tested in the NASA/Ames Unitary 11 x 11 foot transonic and 9 x 7 foot supersonic tunnels to obtain base pressure data with firing solid propellant exhaust plumes. Data system difficulties prevented the acquisition of any useful data in the 9 x 7 tunnel. However, 28 successful rocket test firings were made in the 11 x 11 tunnel, providing base pressure data at Mach numbers of 0.5, 0.9, 1.05, 1.2, and 1.3 and at plume pressure ratios ranging from 11 to 89.

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

NASA Technical Reports Server (NTRS)

Guman, W. J. (Editor)

1972-01-01

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.

The 260: The Largest Solid Rocket Motor Ever Tested

NASA Technical Reports Server (NTRS)

Crimmins, P.; Cousineau, M.; Rogers, C.; Shell, V.

1999-01-01

Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

Solid-propellant rocket motor ballistic performance variation analyses

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.

1975-01-01

Results are presented of research aimed at improving the assessment of off-nominal internal ballistic performance including tailoff and thrust imbalance of two large solid-rocket motors (SRMs) firing in parallel. Previous analyses using the Monte Carlo technique were refined to permit evaluation of the effects of radial and circumferential propellant temperature gradients. Sample evaluations of the effect of the temperature gradients are presented. A separate theoretical investigation of the effect of strain rate on the burning rate of propellant indicates that the thermoelastic coupling may cause substantial variations in burning rate during highly transient operating conditions. The Monte Carlo approach was also modified to permit the effects on performance of variation in the characteristics between lots of propellants and other materials to be evaluated. This permits the variabilities for the total SRM population to be determined. A sample case shows, however, that the effect of these between-lot variations on thrust imbalances within pairs of SRMs is minor in compariosn to the effect of the within-lot variations. The revised Monte Carlo and design analysis computer programs along with instructions including format requirements for preparation of input data and illustrative examples are presented.

Determining fate and transport parameters for nitroglycerine, 2,4-dinitrotoluine, and nitroguanidine in soils

NASA Astrophysics Data System (ADS)

Gosch, D. L.; Dontsova, K.; Chorover, J.; Ferré, T.; Taylor, S.

2010-12-01

During military operations, a small fraction of propellant mass is not consumed during firing and is deposited onto the ground surface (Jenkins et al., 2006). Soluble propellant constituents can be released from particulate residues into the environment. Propellant constituents of interest for this study are nitroglycerine (NG), 2,4-dinitrotoluine (2,4-DNT), 2,6-dinitrotoluine (2,6-DNT), and nitroguanidine (NQ). The goal of this work is to determine fate and transport parameters for these constituents in three soils that represent a range of geographic locations and soil properties. This supports a companion study that looks at dissolution of NG, 2,4-DNT, 2,6-DNT, and NQ from fired and unfired solid propellant formulations and their transport in soils. The three soils selected for the study are Catlin silt loam (fine-silty, mixed, mesic, superactive Oxyaquic Argiudoll), Plymouth sandy loam (mesic, coated Typic Quartzipsamment), and Sassafras loam (fine loamy, siliceous, mesic Typic Hapudult). Two of these soils, Plymouth sandy loam and Sassafras loam, were collected on military installations. Linear adsorption coefficients and transformation rates of propellant constituents were determined in batch kinetic experiments. Soils were mixed with propellant constituent solutions (2 mg L-1) at 4:1 solution/soil mass ratio and equilibrated for 0, 1, 2, 6, 12, 24, 48, and 120 hr at which time samples were centrifuged and supernatant solutions were analyzed for target compounds by high performance liquid chromatography (HPLC) using U.S. EPA Method 8330b for NG, 2,4-DNT, and 2,6-DNT, and Walsh (1989) method for NQ. Adsorption and transformation of propellant constituents were determined from the decrease in solution concentration of these compounds. It was determined that all studied compounds were subjected to sorption by the solid phase and degradation. Catlin soil, with finer texture and high organic matter content, influenced solution concentration of NG, 2,4-DNT, 2,6-DNT, and NQ to the greatest extent. Estimated fate and transport parameters will support ongoing release and column transport studies and will allow environmental managers on military installations to better estimate potential for propellant constituent transport off-site. Jenkins, T.F., A.D. Hewitt, C.L. Grant, S. Thiboutot, G. Ampleman, M.E. Walsh, T.A. Ranney, C.A. Ramsey, A.J. Palazzo, and J.C. Pennington. 2006. Identity and distribution of residues of energetic compounds at army live-fire training ranges. Chemosphere 63:1280-1290. Walsh, M.E. 1989. Analytical Methods for Determining Nitroguanidine in Soil and Water. Special Report 89-35. U.S. Army Cold Regions Research and Engineering Laboratory, Hanover, NH.

Solid propellant exhausted aluminum oxide and hydrogen chloride - Environmental considerations

NASA Technical Reports Server (NTRS)

Cofer, W. R., III; Winstead, E. L.; Purgold, G. C.; Edahl, R. A.

1993-01-01

Measurements of gaseous hydrogen chloride (HCl) and particulate aluminum oxide (Al2O3) were made during penetrations of five Space Shuttle exhaust clouds and one static ground test firing of a shuttle booster. Instrumented aircraft were used to penetrate exhaust clouds and to measure and/or collect samples of exhaust for subsequent analyses. The focus was on the primary solid rocket motor exhaust products, HCl and Al2O3, from the Space Shuttle's solid boosters. Time-dependent behavior of HCl was determined for the exhaust clouds. Composition, morphology, surface chemistry, and particle size distributions were determined for the exhausted Al2O3. Results determined for the exhaust cloud from the static test firing were complicated by having large amounts of entrained alkaline ground debris (soil) in the lofted cloud. The entrained debris may have contributed to neutralization of in-cloud HCl.

30 CFR 57.4230 - Surface self-propelled equipment.

Code of Federal Regulations, 2011 CFR

2011-07-01

... Prevention and Control Firefighting Equipment § 57.4230 Surface self-propelled equipment. (a)(1) Whenever a fire or its effects could impede escape from self-propelled equipment, a fire extinguisher shall be on... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Surface self-propelled equipment. 57.4230...

Test data from solid propellant plume aerodynamics test program in Ames 6 x 6 foot supersonic wind tunnel (shuttle test FA7) (Ames test 033-66)

NASA Technical Reports Server (NTRS)

Hair, L. M.

1975-01-01

The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solid propellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.

Coated oxidizers for combustion stability in solid-propellant rockets

NASA Technical Reports Server (NTRS)

Helmy, A. M.; Ramohalli, K. N. R.

1985-01-01

Experiments are conducted in a laboratory-scale (6.25-cm diameter) end-burning rocket motor with state-of-the-art, ammonium perchlorate hydroxy-terminated polybutadiene (HTPB), nonmetallized propellants. The concept of tailoring the stability characteristics with a small amount (less than 1 percent by weight) of COATING on the oxidizer is explored. The thermal degradation characteristics of the coat chemical are deduced through theoretical arguments on thermal diffusivity of the composite material (propellant). Several candidate coats are selected and propellants are cast. These propellants (with coated oxidizers) are fired in a laboratory-scale end-burning rocket motor, and real-time pressure histories are recorded. The control propellant (with no coating) is also tested for comparison. The uniformity of the coating, confirmed by SEM pictures and BET adsorption measurements, is thought to be an advance in technology. The frequency of bulk mode instability (BMI), the pressure fluctuation amplitudes, and stability boundaries are correlated with parameters related to the characteristic length (L-asterisk) of the rocket motor. The coated oxidizer propellants, in general, display greater combustion stability than the control (state-of-the-art). The correlations of the various parameters are thought to be new to a field filled with much uncertainty.

40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

Code of Federal Regulations, 2014 CFR

2014-07-01

... 40 Protection of Environment 9 2014-07-01 2014-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

Code of Federal Regulations, 2013 CFR

2013-07-01

... 40 Protection of Environment 9 2013-07-01 2013-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

Code of Federal Regulations, 2011 CFR

2011-07-01

... 40 Protection of Environment 8 2011-07-01 2011-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

Code of Federal Regulations, 2012 CFR

2012-07-01

... 40 Protection of Environment 9 2012-07-01 2012-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

Code of Federal Regulations, 2010 CFR

2010-07-01

... 40 Protection of Environment 8 2010-07-01 2010-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

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.

Plume Particle Collection and Sizing from Static Firing of Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Sambamurthi, Jay K.

1995-01-01

Thermal radiation from the plume of any solid rocket motor, containing aluminum as one of the propellant ingredients, is mainly from the microscopic, hot aluminum oxide particles in the plume. The plume radiation to the base components of the flight vehicle is primarily determined by the plume flowfield properties, the size distribution of the plume particles, and their optical properties. The optimum design of a vehicle base thermal protection system is dependent on the ability to accurately predict this intense thermal radiation using validated theoretical models. This article describes a successful effort to collect reasonably clean plume particle samples from the static firing of the flight simulation motor (FSM-4) on March 10, 1994 at the T-24 test bed at the Thiokol space operations facility as well as three 18.3% scaled MNASA motors tested at NASA/MSFC. Prior attempts to collect plume particles from the full-scale motor firings have been unsuccessful due to the extremely hostile thermal and acoustic environment in the vicinity of the motor nozzle.

Influence of different propellant systems on ablation of EPDM insulators in overload state

NASA Astrophysics Data System (ADS)

Guan, Yiwen; Li, Jiang; Liu, Yang; Xu, Tuanwei

2018-04-01

This study examines the propellants used in full-scale solid rocket motors (SRM) and investigates how insulator ablation is affected by two propellant formulations (A and B) during flight overload conditions. An experimental study, theoretical analysis, and numerical simulations were performed to discover the intrinsic causes of insulator ablation rates from the perspective of lab-scaled ground-firing tests, the decoupling of thermochemical ablation, and particle erosion. In addition, the difference in propellant composition, and the insulator charring layer microstructure were analyzed. Results reveal that the degree of insulator ablation is positively correlated with the propellant burn rate, particle velocity, and aggregate concentrations during the condensed phase. A lower ratio of energetic additive material in the AP oxidizer of the propellant is promising for the reduction in particle size and increase in the burn rate and pressure index. However, the overall higher velocity of a two-phase flow causes severe erosion of the insulation material. While the higher ratio of energetic additive to the AP oxidizer imparts a smaller ablation rate to the insulator (under lab-scale test conditions), the slag deposition problem in the combustion chamber may cause catastrophic consequences for future large full-scale SRM flight experiments.

Design, analysis, fabrication and test of the Space Shuttle solid rocket booster motor case

NASA Technical Reports Server (NTRS)

Kapp, J. R.

1978-01-01

The motor case used in the solid propellant booster for the Space Shuttle is unique in many respects, most of which are indigenous to size and special design requirements. The evolution of the case design from initial requirements to finished product is discussed, with increased emphasis of reuse capability, special design features, fracture mechanics and corrosion control. Case fabrication history and the resulting procedure are briefly reviewed with respect to material development, processing techniques and special problem areas. Case assembly, behavior and performance during the DM-1 static firing are reviewed, with appropriate comments and conclusions.

Solid rocket motor fire tests: Phases 1 and 2

NASA Astrophysics Data System (ADS)

Chang, Yale; Hunter, Lawrence W.; Han, David K.; Thomas, Michael E.; Cain, Russell P.; Lennon, Andrew M.

2002-01-01

JHU/APL conducted a series of open-air burns of small blocks (3 to 10 kg) of solid rocket motor (SRM) propellant at the Thiokol Elkton MD facility to elucidate the thermal environment under burning propellant. The propellant was TP-H-3340A for the STAR 48 motor, with a weight ratio of 71/18/11 for the ammonium perchlorate, aluminum, and HTPB binder. Combustion inhibitor applied on the blocks allowed burning on the bottom and/or sides only. Burns were conducted on sand and concrete to simulate near-launch pad surfaces, and on graphite to simulate a low-recession surface. Unique test fixturing allowed propellant self-levitation while constraining lateral motion. Optics instrumentation consisted of a longwave infrared imaging pyrometer, a midwave spectroradiometer, and a UV/visible spectroradiometer. In-situ instrumentation consisted of rod calorimeters, Gardon gauges, elevated thermocouples, flush thermocouples, a two-color pyrometer, and Knudsen cells. Witness materials consisted of yttria, ceria, alumina, tungsten, iridium, and platinum/rhodium. Objectives of the tests were to determine propellant burn characteristics such as burn rate and self-levitation, to determine heat fluxes and temperatures, and to carry out materials analyses. A summary of qualitative results: alumina coated almost all surfaces, the concrete spalled, sand moisture content matters, the propellant self-levitated, the test fixtures worked as designed, and bottom-burning propellant does not self-extinguish. A summary of quantitative results: burn rate averaged 1.15 mm/s, thermocouples peaked at 2070 C, pyrometer readings matched MWIR data at about 2400 C, the volume-averaged plume temperatures were 2300-2400 C with peaks of 2400-2600 C, and the heat fluxes peaked at 125 W/cm2. These results are higher than other researchers' measurements of top-burning propellant in chimneys, and will be used, along with Phase 3 test results, to analyze hardware response to these environments, including General Purpose Heat Sources (GPHS) and Radioisotope Heater Units (RHU). Follow-on Phase 3 tests burning propellant blocks up to 90 kg will be briefly described. .

Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

NASA Astrophysics Data System (ADS)

Sathiyanathan, Kartheephan

This thesis describes the design, fabrication, and testing of a solid propellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

30 CFR 57.4230 - Surface self-propelled equipment.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Section 57.4230 Mineral Resources MINE SAFETY AND HEALTH ADMINISTRATION, DEPARTMENT OF LABOR METAL AND NONMETAL MINE SAFETY AND HEALTH SAFETY AND HEALTH STANDARDS-UNDERGROUND METAL AND NONMETAL MINES Fire... fire or its effects could impede escape from self-propelled equipment, a fire extinguisher shall be on...

Space Shuttle Flight Support Motor no. 1 (FSM-1)

NASA Technical Reports Server (NTRS)

Hughes, Phil D.

1990-01-01

Space Shuttle Flight Support Motor No. 1 (FSM-1) was static test fired on 15 Aug. 1990 at the Thiokol Corporation Static Test Bay T-24. FSM-1 was a full-scale, full-duration static test fire of a redesigned solid rocket motor. FSM-1 was the first of seven flight support motors which will be static test fired. The Flight Support Motor program validates components, materials, and manufacturing processes. In addition, FSM-1 was the full-scale motor for qualification of Western Electrochemical Corporation ammonium perchlorate. This motor was subjected to all controls and documentation requirements CTP-0171, Revision A. Inspection and instrumentation data indicate that the FSM-1 static test firing was successful. The ambient temperature during the test was 87 F and the propellant mean bulk temperature was 82 F. Ballistics performance values were within the specified requirements. The overall performance of the FSM-1 components and test equipment was nominal.

The space shuttle advanced solid rocket motor: Quality control and testing

NASA Technical Reports Server (NTRS)

1991-01-01

The Congressional committees that authorize the activities of NASA requested that the National Research Council (NRC) review the testing and quality assurance programs for the Advanced Solid Rocket Motor (ASRM) program. The proposed ASRM design incorporates numerous features that are significant departures from the Redesigned Solid Rocket Motor (RSRM). The NRC review concentrated mainly on these features. Primary among these are the steel case material, welding rather than pinning of case factory joints, a bolted field joint designed to close upon firing the rocket, continuous mixing and casting of the solid propellant in place of the current batch processes, use of asbestos-free insulation, and a lightweight nozzle. The committee's assessment of these and other features of the ASRM are presented in terms of their potential impact on flight safety.

RSRM TP-H1148 Main Grain Propellant Crack Initiation Evaluation

NASA Technical Reports Server (NTRS)

Earnest, Todd E.

2005-01-01

Pressurized TP-HI 148 propellant fracture toughness testing was performed to assess the potential for initiation of visually undetectable cracks in the RSRM forward segment transition region during motor ignition. Two separate test specimens were used in this evaluation. Testing was performed in cold-gas and hot-fire environments, and under both static and dynamic pressurization conditions. Analysis of test results demonstrates safety factors against initiation of visually undetectable cracks in excess of 8.0. The Reusable Solid Rocket Motor (RSRM) forward segment is cast with PBAN propellant (TP-HI 148) to form T an 1 1-point star configuration that transitions to a tapered center perforated bore (see Figure 1). The geometry of the transition region between the fin valleys and the bore causes a localized area of high strain during horizontal storage. Updated analyses using worst-case mechanical properties at 40 F and improved modeling techniques indicated a slight reduction in safety margins over previous predictions. Although there is no history of strain induced cracks or flaws in the transition region propellant, a proactive test effort was initiated to better understand the implications of the new analysis, primarily the resistance of TP-H1148 propellant to crack initiation' during RSRM ignition.

The Fire Environment of a Solid Rocket Propellant Burning in Air

DTIC Science & Technology

1979-03-01

Weapons Labo.*atory. K4rtland Air Force Base, New Mexico. Lieutenant Michael L. Crawford (DY\\) was the .aboratory Project Officer.in-Charge. When US ...Government drawirgs. specifications, or other data are used fcr any purp:se other than a definitely related Government procurement operaticn. ".’a 4overnment...other, person or corporation, cr conveying any rights or permission to manufacture, use , or sell any patented Invention that may in any tray be related

Bruel and Kjaer 4944 Microphone Grid Frequency Response Function System Identification

NASA Technical Reports Server (NTRS)

Bennett, Reginald; Lee, Erik

2010-01-01

Br el & Kjaer (B&K) 4944B pressure field microphone was judiciously selected to measure acoustic environments, 400Hz 50kHz, in close proximity of the nozzle during multiple firings of solid propellant rocket motors. It is well known that protective grids can affect the frequency response of microphones. B&K recommends operation of the B&K 4944B without a protective grid when recording measurements above 10 to 15 kHz.

Effect of ambient vibration on solid rocket motor grain and propellant/liner bonding interface

NASA Astrophysics Data System (ADS)

Cao, Yijun; Huang, Weidong; Li, Jinfei

2017-05-01

In order to study the condition of structural integrity in the process of the solid propellant motor launching and transporting, the stress and strain field analysis were studied on a certain type of solid propellant motor. the vibration acceleration on the solid propellant motors' transport process were monitored, then the original vibration data was eliminated the noise and the trend term efficiently, finally the characteristic frequency of vibration was got to the finite element analysis. Experiment and simulation results show that the monitored solid propellant motor mainly bear 0.2 HZ and 15 HZ low frequency vibration in the process of transportation; Under the low frequency vibration loading, solid propellant motor grain stress concentration position is respectively below the head and tail of the propellant/liner bonding surface and the grain roots.

Solid-propellant rocket motor internal ballistic performance variation analysis, phase 2

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.

1976-01-01

The Monte Carlo method was used to investigate thrust imbalance and its first time derivative throughtout the burning time of pairs of solid rocket motors firing in parallel. Results obtained compare favorably with Titan 3 C flight performance data. Statistical correlations of the thrust imbalance at various times with corresponding nominal trace slopes suggest several alternative methods of predicting thrust imbalance. The effect of circular-perforated grain deformation on internal ballistics is discussed, and a modified design analysis computer program which permits such an evaluation is presented. Comparisons with SRM firings indicate that grain deformation may account for a portion of the so-called scale factor on burning rate between large motors and strand burners or small ballistic test motors. Thermoelastic effects on burning rate are also investigated. Burning surface temperature is calculated by coupling the solid phase energy equation containing a strain rate term with a model of gas phase combustion zone using the Zeldovich-Novozhilov technique. Comparisons of solutions with and without the strain rate term indicate a small but possibly significant effect of the thermoelastic coupling.

Hybrid rocket motor testing at Nammo Raufoss A/S

NASA Astrophysics Data System (ADS)

Rønningen, Jan-Erik; Kubberud, Nils

2005-08-01

Hybrid rocket motor technology and the use of hybrid rockets have gained increased interest in recent years in many countries. A typical hybrid rocket consists of a tank containing the oxidizer in either liquid or gaseous state connected to the combustion chamber containing an injector, inert solid fuel grain and nozzle. Nammo Raufoss A/S has for almost 40 years designed and produced high-performance solid propellant rocket motors for many military missile systems as well as solid propellant rocket motors for civil space use. In 2003 an in-house technology program was initiated to investigate and study hybrid rocket technology. On 23 September 2004 the first in-house designed hybrid test rocket motor was static test fired at Nammo Raufoss Test Center. The oxidizer was gaseous oxygen contained in a tank pressurized to 10MPa, flow controlled through a sonic orifice into the combustion chamber containing a multi port radial injector and six bore cartridge-loaded fuel grain containing a modified HTPB fuel composition. The motor was ignited using a non-explosive heated wire. This paper will present what has been achieved at Nammo Raufoss since the start of the program.

Hot-Fire Testing of a 1N AF-M315E Thruster

NASA Technical Reports Server (NTRS)

Burnside, Christopher G.; Pedersen, Kevin; Pierce, Charles W.

2015-01-01

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends. NASA completed a hot-fire test of a 1N AF-M315E monopropellant thruster at the Marshall Space Flight Center in the small altitude test stand located in building 4205. The thruster is a ground test article used for basic performance determination and catalyst studies. The purpose of the hot-fire testing was for performance determination of a 1N size thruster and form a baseline from which to study catalyst performance and life with follow-on testing to be conducted at a later date. The thruster performed as expected. The result of the hot-fire testing are presented in this paper and presentation.

Distributed Combustion in Solid Propellants

DTIC Science & Technology

1993-03-01

SENTRY. During that year three full scale development motors were test fired. All three motors experienced an unacceptabiy high level of combustion...CO. Thermochemical Implications," Journal of Physical Chemistry , 1986, Vol. 90, pp. 1688-1691. Rundinger, G., "Effect of Velocity Slip on the...resulting equation is found to be M (r, l = Lelnf 1 F (T-f- T’) I F(Tf- Ts) -J (B.20) where (p is given by P = (MvQ1 + McQ + H) Mil and F is the ratio of

Nozzle erosion characterization and minimization for high-pressure rocket motor applications

NASA Astrophysics Data System (ADS)

Evans, Brian

Understanding of the processes that cause nozzle throat erosion and developing methods for mitigation of erosion rate can allow higher operating pressures for advanced rocket motors. However, erosion of the nozzle throat region, which is a strong function of operating pressure, must be controlled to realize the performance gains of higher operating pressures. The objective of this work was the study the nozzle erosion rates at a broad range of pressures from 7 to 34.5 MPa (1,000 to 5,000 psia) using two different rocket motors. The first is an instrumented solidpropellant motor (ISPM), which uses two baseline solid propellants; one is a non-metallized propellant called Propellant S and the other is a metallized propellant called Propellant M. The second test rig is a non-metallized solid-propellant rocket motor simulator (RMS). The RMS is a gas rocket with the ability to vary the combustion-product species composition by systematically varying the flow rates of gaseous reactants. Several reactant mixtures were utilized in the study to determine the relative importance of different oxidizing species (such as H2O, OH, and CO2). Both test rigs are equipped with a windowed nozzle section for real-time X-ray radiography diagnostics of the instantaneous throat variations for deducing the instantaneous erosion rates. The nozzle test section for both motors can also incorporate a nozzle boundary-layer control system (NBLCS) as a means of nozzle erosion mitigation. The effectiveness of the NBLCS at preventing nozzle throat erosion was demonstrated for both the RMS and the ISPM motors at chamber pressures up to 34 MPa (4930 psia). All tests conducted with the NBLCS showed signs of coning of the propellant surface, leading to increased mass burning rate and resultant chamber pressure. Two correlations were developed for the nozzle erosion rates from solid propellant testing, one for metallized propellant and one for non-metallized propellants. The non-metallized propellant correlation also incorporates the RMS data, accounting for swirling flow of the products in the RMS combustor. These correlations are useful for rocket nozzle designs. The correlation for non-metallized propellant and RMS firings was developed in terms of the effective oxidizer mass fraction and effective Reynolds number. The results calculated from this correlation were compared with measured erosion rate data within +/-15% or 0.05 mm/s (2 mils/s). For metallized propellant, the nozzle erosion rate was found to be relatively independent of the concentration of oxidizing species due to the diffusion-controlled process and the partial surface coverage by the liquid Al/Al2O3 layer. The nozzle erosion rate was also found to be lower than those of non-metallized propellant cases. Agreement between predicted and measured erosion rates was found to be within +/-20% or 0.04 mm/s (2 mils/s).

JANNAF 28th Propellant Development and Characterization Subcommittee and 17th Safety and Environmental Protection Subcommittee Joint Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Cocchiaro, James E. (Editor); Mulder, Edwin J. (Editor); Gomez-Knight, Sylvia J. (Editor)

1999-01-01

This volume contains 37 unclassified/unlimited-distribution technical papers that were presented at the JANNAF 28th Propellant Development & Characterization Subcommittee (PDCS) and 17th Safety & Environmental Protection Subcommittee (S&EPS) Joint Meeting, held 26-30 April 1999 at the Town & Country Hotel and the Naval Submarine Base, San Diego, California. Volume II contains 29 unclassified/limited-distribution papers that were presented at the 28th PDCS and 17th S&EPS Joint Meeting. Volume III contains a classified paper that was presented at the 28th PDCS Meeting on 27 April 1999. Topics covered in PDCS sessions include: solid propellant rheology; solid propellant surveillance and aging; propellant process engineering; new solid propellant ingredients and formulation development; reduced toxicity liquid propellants; characterization of hypergolic propellants; and solid propellant chemical analysis methods. Topics covered in S&EPS sessions include: space launch range safety; liquid propellant hazards; vapor detection methods for toxic propellant vapors and other hazardous gases; toxicity of propellants, ingredients, and propellant combustion products; personal protective equipment for toxic liquid propellants; and demilitarization/treatment of energetic material wastes.

Rheology of composite solid propellants during motor casting

NASA Technical Reports Server (NTRS)

Rogers, C. J.; Smith, P. L.; Klager, K.

1978-01-01

In a study conducted to evaluate flow parameters of uncured solid composite propellants during motor casting, two motors (1.8M-lb grain wt) were cast with a PBAN propellant exhibiting good flow characteristics in a 260-in. dia solid rocket motor. Attention is given to the effects of propellant compositional and processing variables on apparent viscosity as they pertain to rheological behavior and grain defect formation during casting. It is noted that optimized flow behavior is impaired with solid propellant loading. Non-Newtonian pseudoplastic flow is observed, which is dependent upon applied shear stress and the age of the uncured propellant.

Performance analysis of SA-3 missile second stage

NASA Technical Reports Server (NTRS)

Helmy, A. M.

1981-01-01

One SA-3 missile was disassembled. The constituents of the second stage were thoroughly investigated for geometrical details. The second stage slotted composite propellant grain was subjected to mechanical properties testing, physiochemical analyses, and burning rate measurements at different conditions. To determine the propellant performance parameters, the slotted composite propellant grain was machined into a set of small-size tubular grains. These grains were fired in a small size rocket motor with a set of interchangeable nozzles with different throat diameters. The firings were carried out at three different conditions. The data from test motor firings, physiochemical properties of the propellant, burning rate measurement results and geometrical details of the second stage motor, were used as input data in a computer program to compute the internal ballistic characteristics of the second stage.

Laminated chemical and physical micro-jet actuators based on conductive media

NASA Astrophysics Data System (ADS)

Gadiraju, Priya D.

2008-04-01

This dissertation presents the development of electrically-powered, lamination-based microactuators for the realization of large arrays of high impulse and short duration micro-jets with potential applications in the field of micro-electro-mechanical systems (MEMS). Microactuators offer unique control opportunities by converting the input electrical or chemical energy stored in a propellant into useful mechanical energy. This small and precise control obtained can potentially be applied towards aerodynamic control and transdermal drug delivery applications. This thesis work discusses the feasibility of using microactuators for two such applications: Control of the motion of a spinning projectile by utilizing the chemically-driven microjets ejected from the actuators, and enhancement of the permeability properties of skin by selectively ablating the stratum corneum layer of skin using the physical microjets ejected from the actuators. This enhanced permeability of skin can later be used for the delivery of high molecular weight drugs for transdermal drug delivery. The development of electrically powered microactuators starts by fabricating an array of radially firing microactuators using lamination-based microfabrication techniques that potentially enable batch fabrication at low cost. The microactuators of this thesis consist of three main parts: a micro chamber in which the propellant is stored; two electrode structures through which electrical energy is supplied to the propellant; and a micro nozzle through which the propellant or released gases from the propellant are expanded as a jet. Once the actuators are fabricated, they are integrated with MEMS-process-compatible propellants and optimized so as to produce instantaneous ignition of the propellant. This instantaneous ignition is achieved either by making the propellant itself conductive, thus, passing an electric current directly through the propellant; or by discharging an arc across the propellant by placing it between two closely spaced electrodes. The first concept is demonstrated for the application of projectile maneuvering where energetic solid propellant is used in generating a high velocity gaseous jet and the second concept is demonstrated for transdermal drug delivery application where a rapid physical jet of a non-energetic propellant is generated. In the case of chemical-based microactuators, the feasibility of using conductive solid propellant based actuators for maneuvering a 25 mm bluff body projectile spinning at 600 Hz is presented. Several conductive solid propellants are developed and characterized for their electrical conductivity and required ignition energy. Finally, the propellant integrated microactuators are characterized for performance in terms of impulse delivered, thrust generated and duration of the jet. These experimental results are then compared to predicted results from simulations. In the case of physical based microactuators, the feasibility of using released physical jets from the microactuator array for transdermal drug delivery application is presented. Several bio-compatible and FDA-approved liquids are used as propellants and are characterized in terms of thrusts delivered and duration of the released jets. These thermo-mechanical jets are then used to expose skin locally so as to create micro conduits in the stratum corneum layer of skin. Both thermal effects and thermo-mechanical effects of the jet on exposed skin are studied. For both cases, histology of exposed skin is presented and its permeability to drug analog molecules is studied.

Green Monopropellant Status at Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Burnside, Christopher G.; Pierce, Charles W.; Pedersen, Kevin W.

2016-01-01

NASA Marshall Space Flight Center is continuing investigations into the use of green monopropellants as a replacement for hydrazine in spacecraft propulsion systems. Work to date has been to push technology development through multiple activities designed to understand the capabilities of these technologies. Future work will begin to transition to mission pull as these technologies are mature while still keeping a solid goal of pushing technology development as opportunities become available. The AF-M315E activities began with hot-fire demonstration testing of a 1N monopropellant thruster in FY 14 and FY15. Following successful completion of the preliminary campaign, changes to the test stand to accommodate propellant conditioning capability and better control of propellant operations was incorporated to make testing more streamlined. The goal is to conduct hot-fire testing with warm and cold propellants using the existing feed system and original thruster design. Following the 1N testing, a NASA owned 100 mN thruster will be hot-fire tested in the same facility to show feasibility of scaling to smaller thrusters for cubesat applications. The end goal is to conduct a hot-fire test of an integrated cubesat propulsion system using an SLM printed propellant tank, an MSFC designed propulsion system electronic controller and the 100 mN thruster. In addition to the AF-M315E testing, MSFC is pursuing hot-fire testing with LMP-103S. Following our successful hot-fire testing of the 22N thruster in April 2015, a test campaign was proposed for a 440N LMP-103S thruster with Orbital ATK and Plasma Processes. This activity was funded through the Space Technology Mission Directorate (STMD) ACO funding call in the last quarter of CY15. Under the same funding source a test activity with Busek and Glenn Research Center for testing of 5N AF-M315E thrusters was proposed and awarded. Both activities are in-work with expected completion of hot-fire testing by the end of FY17. MSFC is continuing to coordinate with the AF and academia on understanding the chemical reactions that occur in AF-M315E. An on-going investigation of the catalyst bed species using Raman Spectroscopy through the NASA Technology Research Fellowship Program (NSTRF) is looking for ways to minimize the amount of computation required by understanding the intermediate species created in the catalyst bed. The MSFC team is also working with commercial partners through Cooperative Agreement Notices (CAN's). Partnerships with commercial and academia include work in non-catalytic ignition of AF-M315, spark ignition of hybrid cubesat systems, printed SLM tanks, and dual-mode (electric and chemical) propulsion systems is continuing.

Numerical modelling of the work of a pulsed aerosol system for fire fighting at the ignitions of liquid hydrocarbon fuels

NASA Astrophysics Data System (ADS)

Rychkov, A. D.

2009-06-01

The work of a pulsed aerosol system for fire fighting is modelled, which is designed for fire fighting at oil storages and at the spills of oil products, whose vapors were modelled by gaseous methane. The system represents a device for separate installation, which consists of a charge of solid propellant (the gas generator) and a container with fine-dispersed powder of the flame-damper substance. The methane combustion was described by a one-stage gross-reaction, the influence of the concentration of vapors of the flame-damper substance on the combustion process was taken into account by reducing the pre-exponent factor in the Arrhenius law and was described by an empirical dependence. The computational experiment showed that the application of the pulsed aerosol system for fire fighting ensures an efficient transport of fine-dispersed aerosol particles of the flame-damping substance and its forming vapors to the combustion zone; the concentration of particles ensures the damping of the heat source.

Viscoelastic propellant effects on Space Shuttle Dynamics

NASA Technical Reports Server (NTRS)

Bugg, F.

1981-01-01

The program of solid propellant research performed in support of the space shuttle dynamics modeling effort is described. Stiffness, damping, and compressibility of the propellant and the effects of many variables on these properties are discussed. The relationship between the propellant and solid rocket booster dynamics during liftoff and boost flight conditions and the effects of booster vibration and propellant stiffness on free free solid rocket booster modes are described. Coupled modes of the shuttle system and the effect of propellant stiffness on the interfaces of the booster and the external tank are described. A finite shell model of the solid rocket booster was developed.

30 CFR 57.4260 - Underground self-propelled equipment.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Section 57.4260 Mineral Resources MINE SAFETY AND HEALTH ADMINISTRATION, DEPARTMENT OF LABOR METAL AND NONMETAL MINE SAFETY AND HEALTH SAFETY AND HEALTH STANDARDS-UNDERGROUND METAL AND NONMETAL MINES Fire... self-propelled equipment is used underground, a fire extinguisher shall be on the equipment. This...

Predicting Slag Generation in Sub-Scale Test Motors Using a Neural Network

NASA Technical Reports Server (NTRS)

Wiesenberg, Brent

1999-01-01

Generation of slag (aluminum oxide) is an important issue for the Reusable Solid Rocket Motor (RSRM). Thiokol performed testing to quantify the relationship between raw material variations and slag generation in solid propellants by testing sub-scale motors cast with propellant containing various combinations of aluminum fuel and ammonium perchlorate (AP) oxidizer particle sizes. The test data were analyzed using statistical methods and an artificial neural network. This paper primarily addresses the neural network results with some comparisons to the statistical results. The neural network showed that the particle sizes of both the aluminum and unground AP have a measurable effect on slag generation. The neural network analysis showed that aluminum particle size is the dominant driver in slag generation, about 40% more influential than AP. The network predictions of the amount of slag produced during firing of sub-scale motors were 16% better than the predictions of a statistically derived empirical equation. Another neural network successfully characterized the slag generated during full-scale motor tests. The success is attributable to the ability of neural networks to characterize multiple complex factors including interactions that affect slag generation.

30 CFR 56.4230 - Self-propelled equipment.

Code of Federal Regulations, 2010 CFR

2010-07-01

....4230 Mineral Resources MINE SAFETY AND HEALTH ADMINISTRATION, DEPARTMENT OF LABOR METAL AND NONMETAL MINE SAFETY AND HEALTH SAFETY AND HEALTH STANDARDS-SURFACE METAL AND NONMETAL MINES Fire Prevention and Control Firefighting Equipment § 56.4230 Self-propelled equipment. (a)(1) Whenever a fire or its effects...

Effects of a Near Field Pyroshock on the Performance of a Nitramine Nitrocellulose Propellant

NASA Technical Reports Server (NTRS)

Baca, Arcenio

2016-01-01

The purpose of this study is to investigate the effects of a pyroshock environment on the performance characteristics of a propellant used in pyrotechnic devices such as guillotine cutters. A heritage pressure cartridge assembly which uses a nitramine nitrocellulose propellant with a known performance baseline will be exposed to a near field pyroshock event. The pressure cartridge will then be fired in an ambient closed bomb firing to collect pressure time history. This data will be compared to the baseline data to evaluate the effects of the shock on the performance of the propellant.

Composite Solid Propellant Predictability and Quality Assurance

NASA Technical Reports Server (NTRS)

Ramohalli, Kumar

1989-01-01

Reports are presented at the meeting at the University of Arizona on the study of predictable and reliable solid rocket motors. The following subject areas were covered: present state and trends in the research of solid propellants; the University of Arizona program in solid propellants, particularly in mixing (experimental and analytical results are presented).

Solid-propellant rocket motor internal ballistics performance variation analysis, phase 5

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Murph, J. E.

1980-01-01

The results of research aimed at improving the predictability of internal ballistics performance of solid-propellant rocket motors (SRM's) including thrust imbalance between two SRM's firing in parallel are presented. Static test data from the first six Space Shuttle SRM's is analyzed using a computer program previously developed for this purpose. The program permits intentional minor design biases affecting the imbalance between any two SMR's to be removed. Results for the last four of the six SRM's, with only the propellant bulk temperature as a non-random variable, are generally within limits predicted by theory. Extended studies of internal ballistic performance of single SRM's are presented based on an earlier developed mathematical model which includes an assessment of grain deformation. The erosive burning rate law used in the model is upgraded and made more general. Excellent results are obtained in predictions of the performances of five different SRM's of quite different sizes and configurations. These SRM's all employ PBAN type propellants with ammonium perchlorate oxidizer and 16 to 20% aluminum except one which uses carboxyl terminated butadiene binder. The only non-calculated parameters in the burning rate equations that are changed for the different SRM's are the zero crossflow velocity burning rate coefficients and exponents. The results, in general, confirm the importance of grain deformation. The improved internal ballistic model makes practical development of an effective computer program for application of an optimization technique to SRM design which is also demonstrated. The program uses a pattern search technique to minimize the difference between a desired thrust-time trace and one calculated based on the internal ballistic model.

Computer-Controlled Image Anaysis of Solid Propellant Combustion Holograms Using a Quantimet 720 and a PDP-11.

DTIC Science & Technology

1985-09-01

TND 1 96 PIN11. L 4. c. j;. NAVAL POSTGRADUATE SCHOOL Monterey, California NOV 19 19853 THESIS COMPUTER-CONTROLLED IMAGE ANALYSIS OF SOLID PROPELLANT...Controlled Image Analysis of Master’s Thesis Solid Propellant Combustion Holograms September, 1985 Using a Quantimet 720 and a PDP-11 S. PERFORMING ORG...unlimited Computer-Controlled Image Analysis of Solid Propellant * - Combustion Holograms Using a Quantimet 720 and a PDP-11 by Marvin Philip Shook

Space Shuttle Projects

NASA Image and Video Library

1987-07-01

A forward segment is being lowered into the Transient Pressure Test Article (TPTA) test stand at the Marshall Space Flight Center (MSFC) east test area. The TPTA test stand, 14-feet wide, 27-feet long, and 33-feet high, was built in 1987 to provide data to verify the sealing capability of the redesign solid rocket motor (SRM) field and nozzle joints. The test facility applies pressure, temperature, and external loads to a short stack of solid rocket motor hardware. The simulated SRM ignition pressure and temperature transients are achieved by firing a small amount of specially configured solid propellant. The pressure transient is synchronized with external programmable dynamic loads that simulate lift off loads at the external tank attach points. Approximately one million pounds of dead weight on top of the test article simulates the weight of the other Shuttle elements.

Space Shuttle Projects

NASA Image and Video Library

1987-07-01

A forward segment is being lowered into the Transient Pressure Test Article (TPTA) test stand at thw Marshall Space Flight Center (MSFC) east test area. The TPTA test stand, 14-feet wide, 27-feet long, and 33-feet high, was built in 1987 to provide data to verify the sealing capability of the redesign solid rocket motor (SRM) field and nozzle joints. The test facility applies pressure, temperature, and external loads to a short stack of solid rocket motor hardware. The simulated SRM ignition pressure and temperature transients are achieved by firing a small amount of specially configured solid propellant. The pressure transient is synchronized with external programmable dynamic loads that simulate lift off loads at the external tank attach points. Approximately one million pounds of dead weight on top of the test article simulates the weight of the other Shuttle elements.

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

30 CFR 57.4260 - Underground self-propelled equipment.

Code of Federal Regulations, 2011 CFR

2011-07-01

... Prevention and Control Firefighting Equipment § 57.4260 Underground self-propelled equipment. (a) Whenever self-propelled equipment is used underground, a fire extinguisher shall be on the equipment. This... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Underground self-propelled equipment. 57.4260...

The effects of particulates from solid rocket motors fired in space

NASA Technical Reports Server (NTRS)

Mueller, A. C.; Kessler, D. J.

1985-01-01

The orbits attained by kick motor solid propellant particulates are modeled, and an estimate is made of the number of particulates which will remain in orbit. The fuel, Al2O3, is burned while inserting spacecraft into a transfer orbit and again while circularizing the GEO station. It is shown that 23 percent of 1 micron particles deorbit immediately, while most particles enter a retrograde orbit. The resulting flux is an order of magnitude larger than the micrometeoroid flux. The pressures exerted by solar radiation ensure that only 5 percent of the original flux is still in orbit after the first year. The estimates provided are valid for a large number of transfer orbit operations, but will vary widely over the short term.

SOLID PROPELLANT COMBUSTION MECHANISM STUDIES.

DTIC Science & Technology

SOLID ROCKET PROPELLANTS, BURNING RATE), LOW PRESSURE, COMBUSTION PRODUCTS, QUENCHING, THERMAL CONDUCTIVITY, KINETIC THEORY, SURFACE PROPERTIES, PHASE STUDIES, SOLIDS, GASES, PYROLYSIS, MATHEMATICAL ANALYSIS.

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

14 CFR 420.69 - Solid and liquid propellants located together.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Solid and liquid propellants located together. 420.69 Section 420.69 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an...

Hot-Fire Testing of 5N and 22N HPGP Thrusters

NASA Technical Reports Server (NTRS)

Burnside, Christopher G.; Pedersen, Kevin W.; Pierce, Charles W.

2015-01-01

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends.NASA completed hot-fire testing of 5N and 22N HPGP thrusters at the Marshall Space Flight Center’s Component Development Area altitude test stand in April 2015. Both thrusters are ground test articles and not flight ready units, but are representative of potential flight hardware with a known path towards flight application. The purpose of the 5N testing was to perform facility check-outs and generate a small set of data for comparison to ECAPS and Orbital ATK data sets. The 5N thruster performed as expected with thrust and propellant flow-rate data generated that are similar to previous testing at Orbital ATK. Immediately following the 5N testing, and using the same facility, the 22N testing was conducted on the same test stand with the purpose of demonstrating the 22N performance. The results of 22N testing indicate it performed as expected.The results of the hot-fire testing are presented in this paper and presentation.

ISRO's solid rocket motors

NASA Astrophysics Data System (ADS)

Nagappa, R.; Kurup, M. R.; Muthunayagam, A. E.

1989-08-01

Solid rocket motors have been the mainstay of ISRO's sounding rockets and the first generation satellite launch vehicles. For the new launch vehicle under development also, the solid rocket motors contribute significantly to the vehicle's total propulsive power. The rocket motors in use and under development have been developed for a variety of applications and range in size from 30 mm dia employing 450 g of solid propellant—employed for providing a spin to the apogee motors—to the giant 2.8 m dia motor employing nearly 130 tonnes of solid propellant. The initial development, undertaken in 1967 was of small calibre motor of 75 mm dia using a double base charge. The development was essentially to understand the technological elements. Extruded aluminium tubes were used as a rocket motor casing. The fore and aft closures were machined from aluminium rods. The grain was a seven-pointed star with an enlargement of the port at the aft end and was charged into the chamber using a polyester resin system. The nozzle was a metallic heat sink type with graphite throat insert. The motor was ignited with a black powder charge and fired for 2.0 s. Subsequent to this, further developmental activities were undertaken using PVC plastisol based propellants. A class of sounding rockets ranging from 125 to 560 mm calibre were realized. These rocket motors employed improved designs and had delivered lsp ranging from 2060 to 2256 Ns/kg. Case bonding could not be adopted due to the higher cure temperatures of the plastisol propellants but improvements were made in the grain charging techniques and in the design of the igniters and the nozzle. Ablative nozzles based on asbestos phenolic and silica phenolic with graphite inserts were used. For the larger calibre rocket motors, the lsp could be improved by metallic additives. In the early 1970s designs were evolved for larger and more efficient motors. A series of 4 motors for the country's first satellite launch vehicle SLV-3 were developed. The first and second stages of 1 and 0.8 m dia respectively used low carbon steel casing and PBAN propellant. The first stage used segmented construction with a total propellant weight of 8600 kg. The second stage employed about 3 tonnes of the same propellant. The third and fourth stages were of GFRP construction and employed respectively 1100 and 275 kg of CTPB type propellants. Nozzle expansion ratios upto 30 were employed and delivered vacuum lsp of 2766 Ns/kg realized. The fourth stage motor was subsequently used as the apogee motor for orbit injection of India's first geosynchronous satellite—APPLE. All these motors have been flight proven a number of times. Further design improvements have been incorporated and these motors continue to be in use. Starting in 1984 design for a large booster was undertaken. This booster employs a nominal propellant weight of 125 tonne in a 2.8 m dia casing. The motor is expected to be qualified for flight test in 1989. Side by side a high performance motor housing nearly 7 tonnes of propellant in composite casing of 2 m dia and having flex nozzle control system is also under development for upper stage application. Details of the development of the motors, their leading specifications and performance are described.

Solid Propellant Grain Structural Integrity Analysis

NASA Technical Reports Server (NTRS)

1973-01-01

The structural properties of solid propellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.

Portable propellant cutting assembly, and method of cutting propellant with assembly

NASA Technical Reports Server (NTRS)

Sharp, Roger A. (Inventor); Hoskins, Shawn W. (Inventor); Payne, Brett D. (Inventor)

2002-01-01

A propellant cutting assembly and method of using the assembly to cut samples of solid propellant in a repeatable and consistent manner is disclosed. The cutting assembly utilizes two parallel extension beams which are shorter than the diameter of a central bore of an annular solid propellant grain and can be loaded into the central bore. The assembly is equipped with retaining heads at its respective ends and an adjustment mechanism to position and wedge the assembly within the central bore. One end of the assembly is equipped with a cutting blade apparatus which can be extended beyond the end of the extension beams to cut into the solid propellant.

Application of X-ray television image system to observation in solid rocket motor

NASA Astrophysics Data System (ADS)

Fujiwara, T.; Ito, K.; Tanemura, T.; Shimizu, M.; Godai, T.

The X-ray television image system is used to observe the solid propellant burning surface during rocket motor operation as well as to inspect defects in solid rocket motors in a real time manner. This system can test 200 mm diameter dummy propellant rocket motors with under 2 percent discriminative capacity. Viewing of a 50 mm diameter internal-burning rocket motor, propellant burning surface time transition and propellant burning process of the surroundings of artificial defects were satisfactorily observed. The system was demonstrated to be effective for nondestructive testing and combustion research of solid rocket motors.

Propellant Residues Deposition from Firing of 40-mm Grenades

DTIC Science & Technology

2010-09-01

the snow surface downrange of the firing positions in three sampling units on each pad. Samples were analyzed and results compo- sited to derive an...Processing and Analysis ..................................................................... 10 3.1 Snow samples...mm howitzers, propel- lant residues containing DNT were collected from the snow -covered area in front of one of the guns (Walsh, M.E. et al. 2004

Air-Powered Projectile Launcher

NASA Technical Reports Server (NTRS)

Andrews, T.; Bjorklund, R. A.; Elliott, D. G.; Jones, L. K.

1987-01-01

Air-powered launcher fires plastic projectiles without using explosive propellants. Does not generate high temperatures. Launcher developed for combat training for U.S. Army. With reservoir pressurized, air launcher ready to fire. When pilot valve opened, sleeve (main valve) moves to rear. Projectile rapidly propelled through barrel, pushed by air from reservoir. Potential applications in seismic measurements, avalanche control, and testing impact resistance of windshields on vehicles.

Propellant development for the Advanced Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Landers, L. C.; Stanley, C. B.; Ricks, D. W.

1991-01-01

The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

Multiphysics simulation of a novel concept of MEMS-based solid propellant thruster for space propulsion

NASA Astrophysics Data System (ADS)

Moríñigo, José A.; Hermida-Quesada, José

2011-12-01

This work analyzes a novel MEMS-based architecture of submillimeter size thruster for the propulsion of small spacecrafts, addressing its preliminary characterization of performance. The architecture of microthruster comprises a setup of miniaturized channels surrounding the solid-propellant reservoir filled up with a high-energetic polymer. These channels guide the hot gases from the combustion region towards the nozzle entrance located at the opposite side of the thruster. Numerical simulations of the transient response of the combustion gases and wafer heating in thruster firings have been conducted with FLUENT under a multiphysics modelling that fully couples the gas and solid parts involved. The approach includes the gas-wafer and gas-polymer thermal exchange, burnback of the polymer with a simplified non-reacting gas pyrolysis model at its front, and a slip-model inside the nozzle portion to incorporate the effect of gas-surface and rarefaction onto the gas expansion. Besides, accurate characterization of thruster operation requires the inclusion of the receding front of the polymer and heat transfer in the moving gas-solid interfaces. The study stresses the improvement attained in thermal management by the inclusion of lateral micro-channels in the device. In particular, the temperature maps reveal the significant dependence of the thermal loss on the instantaneous surface of the reservoir wall exposed to the heat flux of hot gases. Specifically, the simulations stress the benefit of implementing such a pattern of micro-channels connecting the exit of the combustion reservoir with the nozzle. The results prove that hot gases flowing along the micro-channels exert a sealing action upon the heat flux at the reservoir wall and partly mitigate the overall thermal loss at the inner-wall vicinity during the burnback. The analysis shows that propellant decomposition rate is accelerated due to surface preheating and it suggests that a delay of the flame extinction into the reservoir is possible. The simulated operation of the thruster concept shows encouraging performance.

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.

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.

46 CFR 160.036-2 - Type.

Code of Federal Regulations, 2012 CFR

2012-10-01

... APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-2 Type. (a) Handheld rocket-propelled parachute red flare distress signals specified by this subpart... fired from the hand to provide a rocket-propelled parachute red flare distress signal. (b) [Reserved] ...

46 CFR 160.036-2 - Type.

Code of Federal Regulations, 2014 CFR

2014-10-01

... APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-2 Type. (a) Handheld rocket-propelled parachute red flare distress signals specified by this subpart... fired from the hand to provide a rocket-propelled parachute red flare distress signal. (b) [Reserved] ...

46 CFR 160.036-2 - Type.

Code of Federal Regulations, 2010 CFR

2010-10-01

... APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-2 Type. (a) Handheld rocket-propelled parachute red flare distress signals specified by this subpart... fired from the hand to provide a rocket-propelled parachute red flare distress signal. (b) [Reserved] ...

46 CFR 160.036-2 - Type.

Code of Federal Regulations, 2011 CFR

2011-10-01

... APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-2 Type. (a) Handheld rocket-propelled parachute red flare distress signals specified by this subpart... fired from the hand to provide a rocket-propelled parachute red flare distress signal. (b) [Reserved] ...

46 CFR 160.036-2 - Type.

Code of Federal Regulations, 2013 CFR

2013-10-01

... APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-2 Type. (a) Handheld rocket-propelled parachute red flare distress signals specified by this subpart... fired from the hand to provide a rocket-propelled parachute red flare distress signal. (b) [Reserved] ...

Experimental Characteristics of Particle Dynamics within Solid Rocket Motors Environments

DTIC Science & Technology

2009-04-03

McCrorie, J. D., Vaughn, J. K., Netzer, D. W., “Motor and Plume Particle Size Measurements in Solid Propellant Micromotors ,” Journal of Propulsion...Solid Propellant Micromotors ,” Journal of Propulsion and Power 10(3), 410-418 (1994). 6. Kovalev, O. B., “Motor and Plume Particle Size Prediction in...McCrorie, J. D., Vaughn, J. K., Netzer, D. W., “Motor and Plume Particle Size Measurements in Solid Propellant Micromotors ,” Journal of Propulsion

Comparative Analyses of Creep Models of a Solid Propellant

NASA Astrophysics Data System (ADS)

Zhang, J. B.; Lu, B. J.; Gong, S. F.; Zhao, S. P.

2018-05-01

The creep experiments of a solid propellant samples under five different stresses are carried out at 293.15 K and 323.15 K. In order to express the creep properties of this solid propellant, the viscoelastic model i.e. three Parameters solid, three Parameters fluid, four Parameters solid, four Parameters fluid and exponential model are involved. On the basis of the principle of least squares fitting, and different stress of all the parameters for the models, the nonlinear fitting procedure can be used to analyze the creep properties. The study shows that the four Parameters solid model can best express the behavior of creep properties of the propellant samples. However, the three Parameters solid and exponential model cannot very well reflect the initial value of the creep process, while the modified four Parameters models are found to agree well with the acceleration characteristics of the creep process.

Three-dimensional finite element analysis of acoustic instability of solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Hackett, R. M.; Juruf, R. S.

1976-01-01

A three dimensional finite element solution of the acoustic vibration problem in a solid propellant rocket motor is presented. The solution yields the natural circular frequencies of vibration and the corresponding acoustic pressure mode shapes, considering the coupled response of the propellant grain to the acoustic oscillations occurring in the motor cavity. The near incompressibility of the solid propellant is taken into account in the formulation. A relatively simple example problem is solved in order to illustrate the applicability of the analysis and the developed computer code.

An Overview of Combustion Mechanisms and Flame Structures for Advanced Solid Propellants

NASA Technical Reports Server (NTRS)

Beckstead, M. W.

2000-01-01

Ammonium perchlorate (AP) and cyclotretamethylenetetranitramine (HMX) are two solid ingredients often used in modern solid propellants. Although these two ingredients have very similar burning rates as monopropellants, they lead to significantly different characteristics when combined with binders to form propellants. Part of the purpose of this paper is to relate the observed combustion characteristics to the postulated flame structures and mechanisms for AP and HMX propellants that apparently lead to these similarities and differences. For AP composite, the primary diffusion flame is more energetic than the monopropellant flame, leading to an increase in burning rate over the monopropellant rate. In contrast the HMX primary diffusion flame is less energetic than the HMX monopropellant flame and ultimately leads to a propellant rate significantly less than the monopropellant rate in composite propellants. During the past decade the search for more energetic propellants and more environmentally acceptable propellants is leading to the development of propellants based on ingredients other than AP and HMX. The objective of this paper is to utilize the more familiar combustion characteristics of AP and HMX containing propellants to project the combustion characteristics of propellants made up of more advanced ingredients. The principal conclusion reached is that most advanced ingredients appear to burn by combustion mechanisms similar to HMX containing propellants rather than AP propellants.

Imbedded Thermocouples as a Solid Propellant Combustion Probe

DTIC Science & Technology

1985-04-01

IMBEDDED THERMOCOUPLES AS A SOLID PROPELLANT COMBUSTION PROBE Martin S. Miller Terence P. Coffee Anthony J. Kotlar April 1985 APPROVEO FOR PUBUC...COMPLETING FORM RECIPIENT’S CATALOG NUMBER 4. TITLE (and Subtitle) IMBEDDED THERMOCOUPLES AS A SOLID PROPELLANT COMBUSTION PROBE 7. AuTHORf...this report were presented at the 1984 JANNAF Combustion Meeting 19 KEY WOROS (Continue on reveree aide tl neceeemry end Identity by block number

Regression rate behaviors of HTPB-based propellant combinations for hybrid rocket motor

NASA Astrophysics Data System (ADS)

Sun, Xingliang; Tian, Hui; Li, Yuelong; Yu, Nanjia; Cai, Guobiao

2016-02-01

The purpose of this paper is to characterize the regression rate behavior of hybrid rocket motor propellant combinations, using hydrogen peroxide (HP), gaseous oxygen (GOX), nitrous oxide (N2O) as the oxidizer and hydroxyl-terminated poly-butadiene (HTPB) as the based fuel. In order to complete this research by experiment and simulation, a hybrid rocket motor test system and a numerical simulation model are established. Series of hybrid rocket motor firing tests are conducted burning different propellant combinations, and several of those are used as references for numerical simulations. The numerical simulation model is developed by combining the Navies-Stokes equations with the turbulence model, one-step global reaction model, and solid-gas coupling model. The distribution of regression rate along the axis is determined by applying simulation mode to predict the combustion process and heat transfer inside the hybrid rocket motor. The time-space averaged regression rate has a good agreement between the numerical value and experimental data. The results indicate that the N2O/HTPB and GOX/HTPB propellant combinations have a higher regression rate, since the enhancement effect of latter is significant due to its higher flame temperature. Furthermore, the containing of aluminum (Al) and/or ammonium perchlorate(AP) in the grain does enhance the regression rate, mainly due to the more energy released inside the chamber and heat feedback to the grain surface by the aluminum combustion.

Register of specialized sources for information on selected fuels and oxidizers. [rocket propellants, bibliographies

NASA Technical Reports Server (NTRS)

Ludtke, P. R.

1975-01-01

Thirty-eight (38) organizations are listed and described that catalog and file information in their data systems on fuel and oxidizers. The fuels include hydrogen, methane and hydrazine-type fuels; the oxidizers include oxygen, fluorine, flox, nitrogen tetroxide and ozone. The type of available information covers thermophysical properties, propellant systems, propellant fires-control-extinguishment, propellant explosions, propellant combustion, propellant safety, and fluorine chemistry. These organizations have assembled and collated their information so that it will be useful in the solution of engineering problems.

Technical Evaluation Motor no. 5 (TEM-5)

NASA Technical Reports Server (NTRS)

Cook, M.

1990-01-01

Technical Evaluation Motor No. 5 (TEM-5) was static test fired at the Thiokol Corporation Static Test Bay T-97. TEM-5 was a full scale, full duration static test fire of a high performance motor (HPM) configuration solid rocket motor (SRM). The primary purpose of TEM static tests is to recover SRM case and nozzle hardware for use in the redesigned solid rocket motor (RSRM) flight program. Inspection and instrumentation data indicate that the TEM-5 static test firing was successful. The ambient temperature during the test was 41 F and the propellant mean bulk temperature (PMBT) was 72 F. Ballistics performance values were within the specified requirements. The overall performance of the TEM-5 components and test equipment was nominal. Dissembly inspection revealed that joint putty was in contact with the inner groove of the inner primary seal of the ignitor adapter-to-forward dome (inner) joint gasket; this condition had not occurred on any previous static test motor or flight RSRM. While no qualification issues were addressed on TEM-5, two significant component changes were evaluated. Those changes were a new vented assembly process for the case-to-nozzle joint and the installation of two redesigned field joint protection systems. Performance of the vented case-to-nozzle joint assembly was successful, and the assembly/performance differences between the two field joint protection system (FJPS) configurations were compared.

Studies of the exhaust products from solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Dawbarn, R.; Kinslow, M.

1976-01-01

This study was undertaken to determine the feasibility of conducting environmental chamber tests on the physical processes which occur when a solid rocket motor exhaust mixes with the ambient atmosphere. Of particular interest was the interaction between hydrogen chloride, aluminum oxide, and water vapor. The program consisted of three phases: (1) building a small rocket motor and using it to provide the exhaust species in a controlled environment; (2) evaluating instruments used to detect and measure HCl concentrations and if possible determining whether the HCl existed in the gaseous state or as an acid aerosol; (3) monitoring a series of 6.4-percent scale space shuttle motor tests and comparing the results to the environmental chamber studies. Eighteen firings were conducted in an environmental chamber with the initial ambient relative humidity set at values from 29 to 100 percent. Two additional firings were made in a large shed, and four were made on an open concrete apron. Six test firings at MSFC were monitored, and the ground level concentrations are reported. Evidence is presented which shows that the larger Al2O3 (5 to 50 micrometers) particles from the rocket motor can act as condensation nuclei. Under appropriate ambient conditions where there is sufficient water vapor this results in the formation of an acid aerosol. Droplets of this acid were detected both in the environmental chamber and in the scaled shuttle engine tests.

Ignition of a granular propellant bed

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

Wildegger-Gaissmaier, A.E.; Johnston, I.R.

1996-08-01

An experimental and theoretical study is reported on the ignition process of a low vulnerability ammunition (LOVA) propellant bed in a 127-mm (5-in) bore gun charge. The theoretical investigation was with a two-phase flow interior ballistics code and the model predictions showed the marked influence the igniter system can have on pressure wave development, flame spreading, and the overall interior ballistics performance. A number of different igniter systems were investigated in an empty and propellant-filled gun simulator. Pressure, flame spreading, and high-speed film records were used to analyze the ignition/combustion event. The model predictions for flame spreading were confirmed qualitativelymore » by the experimental data. Full-scale instrumented gun firings were conducted with the optimized igniter design. Pressure waves were not detected in the charge during the firings. Model predictions on overall interior ballistics performance agreed well with the firing data.« less

Dynamic characterization of solid rockets

NASA Technical Reports Server (NTRS)

1973-01-01

The structural dynamics of solid rockets in-general was studied. A review is given of the modes of vibration and bending that can exist for a solid propellant rocket, and a NASTRAN computer model is included. Also studied were the dynamic properties of a solid propellant, polybutadiene-acrylic acid-acrylonitrile terpolymer, which may be used in the space shuttle rocket booster. The theory of viscoelastic materials (i.e, Poisson's ratio) was employed in describing the dynamic properties of the propellant. These studies were performed for an eventual booster stage development program for the space shuttle.

This overview displays the concentration of JPL solid propellant production ...

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

This overview displays the concentration of JPL solid propellant production buildings as seen looking directly north (6 degrees) from the roof of the Administration Building (4231-E-32). The structures closest to the camera contain the equipment for weighing, grinding, mixing, and casting solid propellant grain for motors. Structures in the distance generally house curing or inspection activities. - Jet Propulsion Laboratory Edwards Facility, Edwards Air Force Base, Boron, Kern County, CA

Analysis of capillary drainage from a flat solid strip

NASA Astrophysics Data System (ADS)

Ramé, Enrique; Zimmerli, Gregory A.

2014-06-01

A long and narrow solid strip coated with a thin liquid layer is used as a model of a generic fluid mass probe in a spacecraft propellant tank just after a small thruster firing. The drainage dynamics of the initial coating layer into the settled bulk fluid affects the interpretation of probe measurements as the sensors' signal depends strongly on whether a sensor is in contact with vapor or with liquid. We analyze the drainage under various conditions of zero-gravity (i.e., capillary drainage) and with gravity aligned with the strip length, corresponding to the thruster acceleration. Long-time analytical solutions are found for zero and non-zero gravity. In the case with gravity, an approximate solution is found using matched asymptotics. Estimates show that a thrust of 10-3g0 significantly reduces drainage times.

Effect of Chamber Pressurization Rate on Combustion and Propagation of Solid Propellant Cracks

NASA Astrophysics Data System (ADS)

Yuan, Wei-Lan; Wei, Shen; Yuan, Shu-Shen

2002-01-01

area of the propellant grain satisfies the designed value. But cracks in propellant grain can be generated during manufacture, storage, handing and so on. The cracks can provide additional surface area for combustion. The additional combustion may significantly deviate the performance of the rocket motor from the designed conditions, even lead to explosive catastrophe. Therefore a thorough study on the combustion, propagation and fracture of solid propellant cracks must be conducted. This paper takes an isolated propellant crack as the object and studies the effect of chamber pressurization rate on the combustion, propagation and fracture of the crack by experiment and theoretical calculation. deformable, the burning inside a solid propellant crack is a coupling of solid mechanics and combustion dynamics. In this paper, a theoretical model describing the combustion, propagation and fracture of the crack was formulated and solved numerically. The interaction of structural deformation and combustion process was included in the theoretical model. The conservation equations for compressible fluid flow, the equation of state for perfect gas, the heat conducting equation for the solid-phase, constitutive equation for propellant, J-integral fracture criterion and so on are used in the model. The convective burning inside the crack and the propagation and fracture of the crack were numerically studied by solving the set of nonlinear, inhomogeneous gas-phase governing equations and solid-phase equations. On the other hand, the combustion experiments for propellant specimens with a precut crack were conducted by RTR system. Predicted results are in good agreement with experimental data, which validates the reasonableness of the theoretical model. Both theoretical and experimental results indicate that the chamber pressurization rate has strong effects on the convective burning in the crack, crack fracture initiation and fracture pattern.

Environmentally Compliant Disposal Method for Heavy Metal Containing Propellants

NASA Technical Reports Server (NTRS)

Decker, M. W.; Erickson, E. D.; Byrd, E. R.; Crispin, K. W. R.; Ferguson, B. W.

2000-01-01

ABSTRACT An environmentally friendly, cost effective technology has been developed and demonstrated by a team of Naval Air Warfare Center and Lockheed Martin personnel to dispose of Shillelagh solid rocket motor propellants. The Shillelagh is a surface to surface anti-tank weapon approaching the end of its service life. The current demilitarization process employs open detonation, but the presence of lead stearate in the N5 propellant grain motivated the need for the development of an environmentally friendly disposal method. Contained burning of the propellant followed by propellant exhaust processing was chosen as the disposal methodology. The developmental test bed, completed in February 1998, is inexpensive and transportable. Contained burning of Shillelagh propellants posed two technical hurdles: 1) removal of the sub micron lead and cadmium particulate generated during combustion, and 2) secondary combustion of the significant quantifies of carbon monoxide and hydrogen. A firing chamber with a stepped nozzle, air injection, and active ignition was developed to combust the carbon monoxide and hydrogen in real time. The hot gases and particulates from the combustion process are completely contained within a gas holder. The gases are subsequently cooled and routed through a treatment facility to remove the heavy metal particulate. Results indicate that the lead and cadmium particulates are removed below their respective detection limits (2 micro-g/cu m & 0.2 micro-g/cu m) of the analytical procedures employed and that the carbon monoxide and hydrogen levels have been reduced well below the lower flammability limits. Organic concentrations, principally benzene, are I ppm or less. A semi-automated machine has been developed which can rapidly prepare Shillelagh missiles for the contained burn facility. This machine allows the contained burn technology to be more competitive with current open bum open detonation disposal rates.

Development of strand burner for solid propellant burning rate studies

NASA Astrophysics Data System (ADS)

Aziz, A.; Mamat, R.; Ali, W. K. Wan

2013-12-01

It is well-known that a strand burner is an apparatus that provides burning rate measurements of a solid propellant at an elevated pressure in order to obtain the burning characteristics of a propellant. This paper describes the facilities developed by author that was used in his studies. The burning rate characteristics of solid propellant have be evaluated over five different chamber pressures ranging from 1 atm to 31 atm using a strand burner. The strand burner has a mounting stand that allows the propellant strand to be mounted vertically. The strand was ignited electrically using hot wire, and the burning time was recorded by electronic timer. Wire technique was used to measure the burning rate. Preliminary results from these techniques are presented. This study shows that the strand burner can be used on propellant strands to obtain accurate low pressure burning rate data.

Effect of propellant deformation on ignition and combustion processes in solid propellant cracks

NASA Technical Reports Server (NTRS)

Kumar, M.; Kuo, K. K.

1980-01-01

A comprehensive theoretical model was formulated to study the development of convective burning in a solid propellant crack which continually deforms due to burning and pressure loading. In the theoretical model, the effect of interrelated structural deformation and combustion processes was taken into account by considering (1) transient, one dimensional mass, momentum, and energy conservation equations in the gas phase; (2) a transient, one dimensional heat conduction equation in the solid phase; and (3) quasi-static deformation of the two dimensional, linear viscoelastic propellant crack caused by pressure loading. Partial closures may generate substantial local pressure peaks along the crack, implying a strong coupling between chamber pressurization, crack combustion, and propellant deformation, especially when the cracks are narrow and the chamber pressurization rates high. The maximum pressure in the crack cavity is generally higher than that in the chamber. The initial flame-spreading process is not affected by propellant deformation.

Invited article: Time accurate mass flow measurements of solid-fueled systems.

PubMed

Olliges, Jordan D; Lilly, Taylor C; Joslyn, Thomas B; Ketsdever, Andrew D

2008-10-01

A novel diagnostic method is described that utilizes a thrust stand mass balance (TSMB) to directly measure time-accurate mass flow from a solid-fuel thruster. The accuracy of the TSMB mass flow measurement technique was demonstrated in three ways including the use of an idealized numerical simulation, verifying a fluid mass calibration with high-speed digital photography, and by measuring mass loss in more than 30 hybrid rocket motor firings. Dynamic response of the mass balance was assessed through weight calibration and used to derive spring, damping, and mass moment of inertia coefficients for the TSMB. These dynamic coefficients were used to determine the mass flow rate and total mass loss within an acrylic and gaseous oxygen hybrid rocket motor firing. Intentional variations in the oxygen flow rate resulted in corresponding variations in the total propellant mass flow as expected. The TSMB was optimized to determine mass losses of up to 2.5 g and measured total mass loss to within 2.5% of that calculated by a NIST-calibrated digital scale. Using this method, a mass flow resolution of 0.0011 g/s or 2% of the average mass flow in this study has been achieved.

Invited Article: Time accurate mass flow measurements of solid-fueled systems

NASA Astrophysics Data System (ADS)

Olliges, Jordan D.; Lilly, Taylor C.; Joslyn, Thomas B.; Ketsdever, Andrew D.

2008-10-01

A novel diagnostic method is described that utilizes a thrust stand mass balance (TSMB) to directly measure time-accurate mass flow from a solid-fuel thruster. The accuracy of the TSMB mass flow measurement technique was demonstrated in three ways including the use of an idealized numerical simulation, verifying a fluid mass calibration with high-speed digital photography, and by measuring mass loss in more than 30 hybrid rocket motor firings. Dynamic response of the mass balance was assessed through weight calibration and used to derive spring, damping, and mass moment of inertia coefficients for the TSMB. These dynamic coefficients were used to determine the mass flow rate and total mass loss within an acrylic and gaseous oxygen hybrid rocket motor firing. Intentional variations in the oxygen flow rate resulted in corresponding variations in the total propellant mass flow as expected. The TSMB was optimized to determine mass losses of up to 2.5 g and measured total mass loss to within 2.5% of that calculated by a NIST-calibrated digital scale. Using this method, a mass flow resolution of 0.0011 g/s or 2% of the average mass flow in this study has been achieved.

Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

NASA Astrophysics Data System (ADS)

Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.

2014-11-01

Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

Survivability of a Propellant Fire Inside a Simulated Military Vehicle Crew Compartment: Part 1 - Baseline Study

DTIC Science & Technology

2013-06-01

representative of those used in particular armoured military vehicles, were considered in this study: a top zone propelling charge module (TCM), an...representative of that used in the trial The layout of the hull of a representative armoured vehicle that was simulated in the trial is depicted in...AFESS) are almost universally employed in armoured vehicle crew compartments. Typically the fire suppressant used is a fluorocarbon- based chemical. As

Solid rocket technology advancements for space tug and IUS applications

NASA Technical Reports Server (NTRS)

Ascher, W.; Bailey, R. L.; Behm, J. W.; Gin, W.

1975-01-01

In order for the shuttle tug or interim upper stage (IUS) to capture all the missions in the current mission model for the tug and the IUS, an auxiliary or kick stage, using a solid propellant rocket motor, is required. Two solid propellant rocket motor technology concepts are described. One concept, called the 'advanced propulsion module' motor, is an 1800-kg, high-mass-fraction motor, which is single-burn and contains Class 2 propellent. The other concept, called the high energy upper stage restartable solid, is a two-burn (stop-restartable on command) motor which at present contains 1400 kg of Class 7 propellant. The details and status of the motor design and component and motor test results to date are presented, along with the schedule for future work.

The Effect of Propellant Optical Properties on Composite Solid Propellant Combustion

DTIC Science & Technology

1991-01-01

i a J’i A tkkkeport of Research to NOffice of Naval Research "The Effect of Propellant Optical Properties on Composite Solid Propellant Combustion...87-0547 _ Period (original): July 1987 - June 1990 (with extension): July 1987- December 1990 January 1991 19 . 2 04 090 a Summary of Research ...Results The results of this research program are summarized below in five categories. Only a brief synopsis of the results and their significance are given

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

NASA Technical Reports Server (NTRS)

1972-01-01

A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.

The cohesive law of particle/binder interfaces in solid propellants

NASA Astrophysics Data System (ADS)

Tan, H.

2011-10-01

Solid propellants are treated as composites with high volume fraction of particles embedded in the polymeric binder. A micromechanics model is developed to establish the link between the microscopic behavior of particle/binder interfaces and the macroscopic constitutive information. This model is then used to determine the tension/shearing coupled interface cohesive law of a redesigned solid rocket motor propellant, based on the experimental data of the stress-strain and dilatation-strain curves for the material under slow rate uniaxial tension.

Application of the endochronic theory of viscoplasticity to solid propellants and sandasphalt concrete

NASA Technical Reports Server (NTRS)

Peng, S. T. J.; Valanis, K. C.

1977-01-01

Solid propellants, sand-asphalt concrete and hard plastics showed rate sensitive mechanical behavior which, in addition, indicated that these materials have a permanent memory of the strain (or loading) path by which their present state was attained. A constitutive equation was formulated in general three dimensional tensorial form by means of irreversible thermodynamics. By using a very simple analytical form, it was shown that the mechanical behavior of solid propellants and sand-asphalt concrete can be readily described.

Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the solid propellant rocket engines for use with the space shuttle booster was conducted. A definition of the specific solid propellant rocket engine stage designs, development program requirements, production requirements, launch requirements, and cost data for each program phase were developed.

Development of a miniature solid propellant rocket motor for use in plume simulation studies

NASA Technical Reports Server (NTRS)

Baran, W. J.

1974-01-01

A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

Fire Safety in Extraterrestrial Environments

NASA Technical Reports Server (NTRS)

Friedman, Robert

1998-01-01

Despite rigorous fire-safety policies and practices, fire incidents are possible during lunar and Martian missions. Fire behavior and hence preventive and responsive safety actions in the missions are strongly influenced by the low-gravity environments in flight and on the planetary surfaces. This paper reviews the understanding and key issues of fire safety in the missions, stressing flame spread, fire detection, suppression, and combustion performance of propellants produced from Martian resources.

Scaled Rocket Testing in Hypersonic Flow

NASA Technical Reports Server (NTRS)

Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

2015-01-01

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

Characteristics of a non-volatile liquid propellant in liquid-fed ablative pulsed plasma thrusters

NASA Astrophysics Data System (ADS)

Ling, William Yeong Liang; Schönherr, Tony; Koizumi, Hiroyuki

2017-02-01

In the past several decades, the use of electric propulsion in spacecraft has experienced tremendous growth. With the increasing adoption of small satellites in the kilogram range, suitable propulsion systems will be necessary in the near future. Pulsed plasma thrusters (PPTs) were the first form of electric propulsion to be deployed in orbit, and are highly suitable for small satellites due to their inherent simplicity. However, their lifetime is limited by disadvantages such as carbon deposition leading to thruster failure, and complicated feeding systems required due to the conventional use of solid propellants (usually polytetrafluoroethylene (PTFE)). A promising alternative to solid propellants has recently emerged in the form of non-volatile liquids that are stable in vacuum. This study presents a broad comparison of the non-volatile liquid perfluoropolyether (PFPE) and solid PTFE as propellants on a PPT with a common design base. We show that liquid PFPE can be successfully used as a propellant, and exhibits similar plasma discharge properties to conventional solid PTFE, but with a mass bit that is an order of magnitude higher for an identical ablation area. We also demonstrate that the liquid PFPE propellant has exceptional resistance to carbon deposition, completely negating one of the major causes of thruster failure, while solid PTFE exhibited considerable carbon build-up. Energy dispersive X-ray spectroscopy was used to examine the elemental compositions of the surface deposition on the electrodes and the ablation area of the propellant (or PFPE encapsulator). The results show that based on its physical characteristics and behavior, non-volatile liquid PFPE is an extremely promising propellant for use in PPTs, with an extensive scope available for future research and development.

Upper Stage Tank Thermodynamic Modeling Using SINDA/FLUINT

NASA Technical Reports Server (NTRS)

Schallhorn, Paul; Campbell, D. Michael; Chase, Sukhdeep; Piquero, Jorge; Fortenberry, Cindy; Li, Xiaoyi; Grob, Lisa

2006-01-01

Modeling to predict the condition of cryogenic propellants in an upper stage of a launch vehicle is necessary for mission planning and successful execution. Traditionally, this effort was performed using custom, in-house proprietary codes, limiting accessibility and application. Phenomena responsible for influencing the thermodynamic state of the propellant have been characterized as distinct events whose sequence defines a mission. These events include thermal stratification, passive thermal control roll (rotation), slosh, and engine firing. This paper demonstrates the use of an off the shelf, commercially available, thermal/fluid-network code to predict the thermodynamic state of propellant during the coast phase between engine firings, i.e. the first three of the above identified events. Results of this effort will also be presented.

On the history of the development of solid-propellant rockets in the Soviet Union

NASA Technical Reports Server (NTRS)

Pobedonostsev, Y. A.

1977-01-01

Pre-World War II Soviet solid-propellant rocket technology is reviewed. Research and development regarding solid composite preparations of pyroxyline TNT powder is described, as well as early work on rocket loading calculations, problems of flight stability, and aircraft rocket launching and ground rocket launching capabilities.

The Effect of Propellant Composition on Secondary Muzzle Blast Overpressure

DTIC Science & Technology

1983-04-01

LOVA propellants evaluated included PU/HMX, CTBN /HMX, HTPB/HMX, CAB/RDX, CA/RDX, Kraton/RDX, and EC/NC/RDX. Details of the propellant compositions...RDX tests. Secondary flash was observed for all the firings of all the other candidates, even CTBN /HMX, which had some suppressant. All of these...Propellant Flame Temp (K) Intensity (Mcd) Observations of Flash Kraton/RDX 2283 18.2 ± 1 . 2 11 11 CTBN /HMX 2379 13.8 ± • 72 8 8 HTPB/HMX 2363 10.5

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.

Effects of aluminum and iron nanoparticle additives on composite AP/HTPB solid propellant regression rate

NASA Astrophysics Data System (ADS)

Styborski, Jeremy A.

This project was started in the interest of supplementing existing data on additives to composite solid propellants. The study on the addition of iron and aluminum nanoparticles to composite AP/HTPB propellants was conducted at the Combustion and Energy Systems Laboratory at RPI in the new strand-burner experiment setup. For this study, a large literature review was conducted on history of solid propellant combustion modeling and the empirical results of tests on binders, plasticizers, AP particle size, and additives. The study focused on the addition of nano-scale aluminum and iron in small concentrations to AP/HTPB solid propellants with an average AP particle size of 200 microns. Replacing 1% of the propellant's AP with 40-60 nm aluminum particles produced no change in combustive behavior. The addition of 1% 60-80 nm iron particles produced a significant increase in burn rate, although the increase was lesser at higher pressures. These results are summarized in Table 2. The increase in the burn rate at all pressures due to the addition of iron nanoparticles warranted further study on the effect of concentration of iron. Tests conducted at 10 atm showed that the mean regression rate varied with iron concentration, peaking at 1% and 3%. Regardless of the iron concentration, the regression rate was higher than the baseline AP/HTPB propellants. These results are summarized in Table 3.

Design and Fabrication of a 200N Thrust Rocket Motor Based on NH4ClO4+Al+HTPB as Solid Propellant

NASA Astrophysics Data System (ADS)

Wahid, Mastura Ab; Ali, Wan Khairuddin Wan

2010-06-01

The development of rocket motor using potassium nitrate, carbon and sulphur mixture has successfully been developed by researchers and students from UTM and recently a new combination for solid propellant is being created. The new solid propellant will combine a composition of Ammonium perchlorate, NH4ClO4 with aluminium, Al and Hydroxyl Terminated Polybutadiene, HTPB as the binder. It is the aim of this research to design and fabricate a new rocket motor that will produce a thrust of 200N by using this new solid propellant. A static test is done to obtain the thrust produced by the rocket motor and analyses by observation and also calculation will be done. The experiment for the rocket motor is successful but the thrust did not achieve its required thrust.

Characterization of aluminum/RP-1 gel propellant properties

NASA Technical Reports Server (NTRS)

Rapp, Douglas C.; Zurawski, Robert L.

1988-01-01

Research efforts are being conducted by the NASA Lewis Research Center to formulate and characterize the properties of Al/RP-1 and RP-1 gelled propellants for rocket propulsion systems. Twenty four different compositions of gelled fuels were formulated with 5 and 16 micron, atomized aluminum powder in RP-1. The total solids concentration in the propellant varied from 5 to 60 wt percent. Tests were conducted to evaluate the stability and rheological characteristics of the fuels. Physical separation of the solids occurred in fuels with less than 50 wt percent solids concentration. The rheological characteristics of the Al/RP-1 fuels varied with solids concentration. Both thixotropic and rheopectic gel behavior were observed. The unmetallized RP-1 gels, which were formulated by a different technique than the Al/RP-1 gels, were highly viscoelastic. A history of research efforts which were conducted to formulate and characterize the properties of metallized propellants for various applications is also given.

Laboratory test methods for combustion stability properties of solid propellants

NASA Technical Reports Server (NTRS)

Strand, L. D.; Brown, R. S.

1992-01-01

An overview is presented of experimental methods for determining the combustion-stability properties of solid propellants. The methods are generally based on either the temporal response to an initial disturbance or on external methods for generating the required oscillations. The size distribution of condensed-phase combustion products are characterized by means of the experimental approaches. The 'T-burner' approach is shown to assist in the derivation of pressure-coupled driving contributions and particle damping in solid-propellant rocket motors. Other techniques examined include the rotating-valve apparatus, the impedance tube, the modulated throat-acoustic damping burner, and the magnetic flowmeter. The paper shows that experimental methods do not exist for measuring the interactions between acoustic velocity oscillations and burning propellant.

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 2 2011-01-01 2011-01-01 false Agricultural and fire fighting airplanes... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 2 2013-01-01 2013-01-01 false Agricultural and fire fighting airplanes... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 2 2012-01-01 2012-01-01 false Agricultural and fire fighting airplanes... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 2 2010-01-01 2010-01-01 false Agricultural and fire fighting airplanes... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

Hybrid propulsion technology program. Volume 2: Technology definition package

NASA Technical Reports Server (NTRS)

Jensen, Gordon E.; Holzman, Allen L.; Leisch, Steven O.; Keilbach, Joseph; Parsley, Randy; Humphrey, John

1989-01-01

A concept design study was performed to configure two sizes of hybrid boosters; one which duplicates the advanced shuttle rocket motor vacuum thrust time curve and a smaller, quarter thrust level booster. Two sizes of hybrid boosters were configured for either pump-fed or pressure-fed oxygen feed systems. Performance analyses show improved payload capability relative to a solid propellant booster. Size optimization and fuel safety considerations resulted in a 4.57 m (180 inch) diameter large booster with an inert hydrocarbon fuel. The preferred diameter for the quarter thrust level booster is 2.53 m (96 inches). The demonstration plan would culminate with test firings of a 3.05 m (120 inch) diameter hybrid booster.

One-Dimensional Modelling of Internal Ballistics

NASA Astrophysics Data System (ADS)

Monreal-González, G.; Otón-Martínez, R. A.; Velasco, F. J. S.; García-Cascáles, J. R.; Ramírez-Fernández, F. J.

2017-10-01

A one-dimensional model is introduced in this paper for problems of internal ballistics involving solid propellant combustion. First, the work presents the physical approach and equations adopted. Closure relationships accounting for the physical phenomena taking place during combustion (interfacial friction, interfacial heat transfer, combustion) are deeply discussed. Secondly, the numerical method proposed is presented. Finally, numerical results provided by this code (UXGun) are compared with results of experimental tests and with the outcome from a well-known zero-dimensional code. The model provides successful results in firing tests of artillery guns, predicting with good accuracy the maximum pressure in the chamber and muzzle velocity what highlights its capabilities as prediction/design tool for internal ballistics.

High-Energy Propellant Rocket Firing at the Rocket Lab

NASA Image and Video Library

1955-01-21

A rocket using high-energy propellant is fired from the Rocket Laboratory at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Lab was a collection of ten one-story cinderblock test cells located behind earthen barriers at the western edge of the campus. The rocket engines tested there were comparatively small, but the Lewis researchers were able to study different configurations, combustion performance, and injectors and nozzle design. The rockets were generally mounted horizontally and fired, as seen in this photograph of Test Cell No. 22. A group of fuels researchers at Lewis refocused their efforts after World War II in order to explore high energy propellants, combustion, and cooling. Research in these three areas began in 1945 and continued through the 1960s. The group of rocket researches was not elevated to a division branch until 1952. The early NACA Lewis work led to the development of liquid hydrogen as a viable propellant in the late 1950s. Following the 1949 reorganization of the research divisions, the rocket group began working with high-energy propellants such as diborane, pentaborane, and hydrogen. The lightweight fuels offered high levels of energy but were difficult to handle and required large tanks. In late 1954, Lewis researchers studied the combustion characteristics of gaseous hydrogen in a turbojet combustor. Despite poor mixing of the fuel and air, it was found that the hydrogen yielded more than a 90-percent efficiency. Liquid hydrogen became the focus of Lewis researchers for the next 15 years.

Fast-acting sprinkler system design considerations for propellant manufacture

NASA Astrophysics Data System (ADS)

Matthews, A. L.; Crable, J. M.; Kristoff, P. T.

1984-08-01

Fast-acting sprinkler systems for detection and suppression of fires in propellant operations, which require activation in the millisecond range in order to be effective, can be easily defeated unless particular attention is paid to design and maintenance details. Of primary consideration are detector selection and placement in processes to minimize the effect of environmental influences. Also important are nozzle placement, water flow density, water supply pressure, and pattern and sloping of piping. When all of these design criteria are properly implemented, water application can occur within 100 ms of fire detection.

Boundary cooled rocket engines for space storable propellants

NASA Technical Reports Server (NTRS)

Kesselring, R. C.; Mcfarland, B. L.; Knight, R. M.; Gurnitz, R. N.

1972-01-01

An evaluation of an existing analytical heat transfer model was made to develop the technology of boundary film/conduction cooled rocket thrust chambers to the space storable propellant combination oxygen difluoride/diborane. Critical design parameters were identified and their importance determined. Test reduction methods were developed to enable data obtained from short duration hot firings with a thin walled (calorimeter) chamber to be used quantitatively evaluate the heat absorbing capability of the vapor film. The modification of the existing like-doublet injector was based on the results obtained from the calorimeter firings.

30 CFR 56.4230 - Self-propelled equipment.

Code of Federal Regulations, 2011 CFR

2011-07-01

... other persons in the area, a fire extinguisher shall be on the equipment or within 100 feet of the... size that can extinguish fires of any class in their early stages which could originate from the...

Propagation of a Chemical Reaction through Heterogeneous Lithium- Polytetrafluoroethylene Mixtures

DTIC Science & Technology

1975-12-11

Condensed Phases ........... ............... 9 1.2.1 Lithium-Gas Surface Reactions. .......... 10 1.2.2 Composite Solid Propellant Combustion. . .. 13...f:- the o:cu:=ence _A a surface reaction was developed, but no analyti7al reaction zate model was presented- 1.2.2 Composite S’-lid Propellant...Combustion Composite solid propellants are plastic-like materials consisting of small oxidizer particles embedded in a fuel matrix. Ammonium perchlorate is

Real-Time Inhibitor Recession Measurements in Two Space Shuttle Reusable Solid Rocket Motors

NASA Technical Reports Server (NTRS)

McWhorter, B. B.; Ewing, M. E.; Bolton, D. E.; Albrechtsen, K. U.; Earnest, T. E.; Noble, T. C.; Longaker, M.

2003-01-01

Real-time internal motor insulation char line recession measurements have been evaluated for two full-scale static tests of the Space Shuttle Reusable Solid Rocket Motor (RSRM). These char line recession measurements were recorded on the forward facing propellant grain inhibitors to better understand the thermal performance of these inhibitors. The RSRM propellant grain inhibitors are designed to erode away during motor operation, thus making it difficult to use post-fire observations to determine inhibitor thermal performance. Therefore, this new internal motor instrumentation is invaluable in establishing an accurate understanding of inhibitor recession versus motor operation time. The data for the first test was presented at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (AIAA 2001-3280) in July 2001. Since that time, a second full scale static test has delivered additional real-time data on inhibitor thermal performance. The evaluation of this data is presented in this paper. The second static test, in contrast to the first test, used a slightly different arrangement of instrumentation in the inhibitors. This instrumentation has yielded a better understanding of the inhibitor time dependent inboard tip recession. Graphs of inhibitor recession profiles with time are presented. Inhibitor thermal ablation models have been created from theoretical principals. The model predictions compare favorably with data from both tests. This verified modeling effort is important to support new inhibitor designs for a five segment Space Shuttle solid rocket motor. The internal instrumentation project on RSRM static tests is providing unique opportunities for other real-time internal motor measurements that could not otherwise be directly quantified.

Technique for Evaluating the Erosive Properties of Ablative Internal Insulation Materials

NASA Technical Reports Server (NTRS)

McComb, J. C.; Hitner, J. M.

1989-01-01

A technique for determining the average erosion rate versus Mach number of candidate internal insulation materials was developed for flight motor applications in 12 inch I.D. test firing hardware. The method involved the precision mounting of a mechanical measuring tool within a conical test cartridge fabricated from either a single insulation material or two non-identical materials each of which constituted one half of the test cartridge cone. Comparison of the internal radii measured at nine longitudinal locations and between eight to thirty two azimuths, depending on the regularity of the erosion pattern before and after test firing, permitted calculation of the average erosion rate and Mach number. Systematic criteria were established for identifying erosion anomalies such as the formation of localized ridges and for excluding such anomalies from the calculations. The method is discussed and results presented for several asbestos-free materials developed in-house for the internal motor case insulation in solid propellant rocket motors.

118. #3 SHAFT ALLEY (PROPELLER SHAFT) FORWARD LOOKING AFT ...

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

118. #3 SHAFT ALLEY (PROPELLER SHAFT) - FORWARD LOOKING AFT ON PORT SIDE SHOWING THE SHAFT, SHAFT PACKING GLAND, SHAFT SEAL COOLING WATER LINE AND FIVE INCH FIRE MAIN PIPING. - U.S.S. HORNET, Puget Sound Naval Shipyard, Sinclair Inlet, Bremerton, Kitsap County, WA

A theoretical evaluation of aluminum gel propellant two-phase flow losses on vehicle performance

NASA Technical Reports Server (NTRS)

Mueller, Donn C.; Turns, Stephen R.

1993-01-01

A one-dimensional model of a hydrocarbon/Al/O2(gaseous) fueled rocket combustion chamber was developed to study secondary atomization effects on propellant combustion. This chamber model was coupled with a two dimensional, two-phase flow nozzle code to estimate the two-phase flow losses associated with solid combustion products. Results indicate that moderate secondary atomization significantly reduces propellant burnout distance and Al2O3 particle size; however, secondary atomization provides only moderate decreases in two-phase flow induced I(sub sp) losses. Despite these two-phase flow losses, a simple mission study indicates that aluminum gel propellants may permit a greater maximum payload than the hydrocarbon/O2 bi-propellant combination for a vehicle of fixed propellant volume. Secondary atomization was also found to reduce radiation losses from the solid combustion products to the chamber walls, primarily through reductions in propellant burnout distance.

Effects of Near Field Pyroshock on the Performance of a Nitramine Nitrocellulose Propellant

NASA Technical Reports Server (NTRS)

Baca, Arcenio B.

2016-01-01

The overall purpose of this study is to investigate the effects of a pyroshock environment on the performance characteristics of a propellant used in pyrotechnic devices such as guillotine cutters. Near field pyroshock which is defined by acceleration amplitudes in excess of 10,000g at a frequency of greater than 10,000 Hz is a highly transient environment that has a known potential to cause failure in both structural and electronic components. A heritage pressure cartridge assembly which uses a nitramine nitrocellulose propellant with a known performance baseline will be exposed to a near field pyroshock event. The pressure cartridge will then be fired in an ambient closed bomb firing to collect pressure time history. The two performance characteristics that will be evaluated are the pressure amplitude and time to peak pressure. This data will be compared to the base-lined ambient closed bomb data to evaluate the effects of the shock on the performance of the propellant. It is expected that the pyroshock environment will cause brittle failures of the propellant increasing the surface area of said propellant. This increase of surface area should result in increased combustion rate which should show as an increased pressure peak and decreased time to peak pressure in the pressure time data.

Concept and performance study of turbocharged solid propellant ramjet

NASA Astrophysics Data System (ADS)

Li, Jiang; Liu, Kai; Liu, Yang; Liu, Shichang

2018-06-01

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

Minimization of Roll Firings for Optimal Propellant Maneuvers

NASA Astrophysics Data System (ADS)

Leach, Parker C.

Attitude control of the International Space Station (ISS) is critical for operations, impacting power, communications, and thermal systems. The station uses gyroscopes and thrusters for attitude control, and reorientations are normally assisted by thrusters on docked vehicles. When the docked vehicles are unavailable, the reduction in control authority in the roll axis results in frequent jet firings and massive fuel consumption. To improve this situation, new guidance and control schemes are desired that provide control with fewer roll firings. Optimal control software was utilized to solve for potential candidates that satisfied desired conditions with the goal of minimizing total propellant. An ISS simulation too was then used to test these solutions for feasibility. After several problem reformulations, multiple candidate solutions minimizing or completely eliminating roll firings were found. Flight implementation would not only save massive amounts of fuel and thus money, but also reduce ISS wear and tear, thereby extending its lifetime.

VIABILITY OF BACILLUS SUBTILIS SPORES IN ROCKET PROPELLANTS.

PubMed

GODDING, R M; LYNCH, V H

1965-01-01

The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N(2)O(4), monomethylhydrazine and 1,1-dimethylhydrazine. N(2)O(4) was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components.

Holographic investigation of solid propellant particulates

NASA Astrophysics Data System (ADS)

Gillespie, T. R.

1981-12-01

The investigation completed the development process to establish a technique to obtain holographic recordings of particulate behavior during the combustion process of solid propellants in a two-dimensional rocket motor. Holographic and photographic recordings were taken in a crossflow environment using various compositions of metallized propellants. The reconstructed holograms are used to provide data on the behavior of aluminum/aluminum oxide particulates in a steady state combustion environment as a function of the initial aluminum size cast into the propellant. High speed, high resolution motion pictures were taken to compare the cinematic data with that available from the holograms.

Viability of Bacillus subtilis Spores in Rocket Propellants

PubMed Central

Godding, Rogene M.; Lynch, Victoria H.

1965-01-01

The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N2O4, monomethylhydrazine and 1,1-dimethylhydrazine. N2O4 was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components. PMID:14264838

14 CFR 420.65 - Handling of solid propellants.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Handling of solid propellants. 420.65 Section 420.65 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION... from the closest debris or explosive hazard source in an explosive hazard facility. ...

14 CFR 420.65 - Handling of solid propellants.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Handling of solid propellants. 420.65 Section 420.65 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION... from the closest debris or explosive hazard source in an explosive hazard facility. ...

14 CFR 420.65 - Handling of solid propellants.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Handling of solid propellants. 420.65 Section 420.65 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION... from the closest debris or explosive hazard source in an explosive hazard facility. ...

76 FR 51459 - Office of Commercial Space Transportation (AST); Notice of Availability of the Finding of No...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-08-18

... five solid-propellant strap-on rocket motors to the Atlas V launch vehicle and larger solid- propellant strap-on rocket motors on the Delta IV vehicle. The FAA participated as a cooperating agency in...

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.

Lessons Learned Entry: Hypergolic Propellant Related Spills and Fires

NASA Technical Reports Server (NTRS)

Nufer, Brian

2009-01-01

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.

Mars Ascent Vehicle-Propellant Aging

NASA Technical Reports Server (NTRS)

Dankanich, John; Rousseau, Jeremy; Williams, Jacob

2015-01-01

This project is to develop and test a new propellant formulation specifically for the Mars Ascent Vehicle (MAV) for the robotic Mars Sample Return mission. The project was initiated under the Planetary Sciences Division In-Space Propulsion Technology (ISPT) program and is continuing under the Mars Exploration Program. The two-stage, solid motor-based MAV has been the leading MAV solution for more than a decade. Additional studies show promise for alternative technologies including hybrid and bipropellant options, but the solid motor design has significant propellant density advantages well suited for physical constraints imposed while using the SkyCrane descent stage. The solid motor concept has lower specific impulse (Isp) than alternatives, but if the first stage and payload remain sufficiently small, the two-stage solid MAV represents a potential low risk approach to meet the mission needs. As the need date for the MAV slips, opportunities exist to advance technology with high on-ramp potential. The baseline propellant for the MAV is currently the carboxyl terminated polybutadiene (CTPB) based formulation TP-H-3062 due to its advantageous low temperature mechanical properties and flight heritage. However, the flight heritage is limited and outside the environments, the MAV must endure. The ISPT program competed a propellant formulation project with industry and selected ATK to develop a new propellant formulation specifically for the MAV application. Working with ATK, a large number of propellant formulations were assessed to either increase performance of a CTPB propellant or improve the low temperature mechanical properties of a hydroxyl terminated polybutadiene (HTPB) propellant. Both propellants demonstrated potential to increase performance over heritage options, but an HTPB propellant formulation, TP-H-3544, was selected for production and testing. The test plan includes propellant aging first at high vacuum conditions, representative of the Mars transit, followed by an additional year at simulated Mars surface conditions. The actual Mars surface environment is based on the igloo design, actively maintains the propellant at or above -40 degC, 95% carbon dioxide at Mars surface pressure. The NASA Marshall Space Flight Center (MSFC) Mars environment test facility is shown in figure 1 and located in the East Test area of Redstone Arsenal due to storage of live propellants. The facility consists of a vacuum chamber placed inside a large freezer unit. The facility includes pressure and temperature monitoring equipment in addition to a vacuum quality monitoring system spectrometer to record any outgassing products.

New Propellants and Cryofuels

NASA Technical Reports Server (NTRS)

Palasezski, Bryan; Sullivan, Neil S.; Hamida, Jaha; Kokshenev, V.

2006-01-01

The proposed research will investigate the stability and cryogenic properties of solid propellants that are critical to NASA s goal of realizing practical propellant designs for future spacecraft. We will determine the stability and thermal properties of a solid hydrogen-liquid helium stabilizer in a laboratory environment in order to design a practical propellant. In particular, we will explore methods of embedding atomic species and metallic nano-particulates in hydrogen matrices suspended in liquid helium. We will also measure the characteristic lifetimes and diffusion of atomic species in these candidate cryofuels. The most promising large-scale advance in rocket propulsion is the use of atomic propellants; most notably atomic hydrogen stabilized in cryogenic environments, and metallized-gelled liquid hydrogen (MGH) or densified gelled hydrogen (DGH). The new propellants offer very significant improvements over classic liquid oxygen/hydrogen fuels because of two factors: (1) the high energy-release, and (ii) the density increase per unit energy release. These two changes can lead to significant reduced mission costs and increased payload to orbit weight ratios. An achievable 5 to 10 percent improvement in specific impulse for the atomic propellants or MGH fuels can result in a doubling or tripling of system payloads. The high-energy atomic propellants must be stored in a stabilizing medium such as solid hydrogen to inhibit or delay their recombination into molecules. The goal of the proposed research is to determine the stability and thermal properties of the solid hydrogen-liquid helium stabilizer. Magnetic resonance techniques will be used to measure the thermal lifetimes and the diffusive motions of atomic species stored in solid hydrogen grains. The properties of metallic nano-particulates embedded in hydrogen matrices will also be studied and analyzed. Dynamic polarization techniques will be developed to enhance signal/noise ratios in order to be able to detect low concentrations of the introduced species. The required lifetimes for atomic hydrogen and other species can only be realized at low temperatures to avoid recombination of atoms before use as a fuel.

Characterization of the non axial thrust generated by large solid propellant rocket motors in three axis stabilized ascent

NASA Technical Reports Server (NTRS)

Kosmann, W. J.; Dionne, E. R.; Klemetson, R. W.

1978-01-01

Nonaxial thrusts produced by solid rocket motors during three-axis stabilized attitude control have been determined from ascent experience on twenty three Burner II, Burner IIA and Block 5D-1 upper stage vehicles. A data base representing four different rocket motor designs (three spherical and one extended spherical) totaling twenty five three-axis stabilized firings is generated. Solid rocket motor time-varying resultant and lateral side force vector magnitudes, directions and total impulses, and roll torque couple magnitudes, directions, and total impulses are tabulated in the appendix. Population means and three sigma deviations are plotted. Existing applicable ground test side force and roll torque magnitudes and total impulses are evaluated and compared to the above experience data base. Within the spherical motor population, the selected AEDC ground test data consistently underestimated experienced motor side forces, roll torques and total impulses. Within the extended spherical motor population, the selected AEDC test data predicted experienced motor side forces, roll torques, and total impulses, with surprising accuracy considering the very small size of the test and experience populations.

Investigation of the NACA 4-(3)(8)-045 Two-blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Compressibility and Solidity on Performance

NASA Technical Reports Server (NTRS)

Stack, John; Draley, Eugene C; Delano, James B; Feldman, Lewis

1950-01-01

As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.

Experimental investigation of the combustion products in an aluminised solid propellant

NASA Astrophysics Data System (ADS)

Liu, Zhu; Li, Shipeng; Liu, Mengying; Guan, Dian; Sui, Xin; Wang, Ningfei

2017-04-01

Aluminium is widely used as an important additive to improve ballistic and energy performance in solid propellants, but the unburned aluminium does not contribute to the specific impulse and has both thermal and momentum two-phase flow losses. So understanding of aluminium combustion behaviour during solid propellant burning is significant when improving internal ballistic performance. Recent developments and experimental results reported on such combustion behaviour are presented in this paper. A variety of experimental techniques ranging from quenching and dynamic measurement, to high-speed CCD video recording, were used to study aluminium combustion behaviour and the size distribution of the initial agglomerates. This experimental investigation also provides the size distribution of the condensed phase products. Results suggest that the addition of an organic fluoride compound to solid propellant will generate smaller diameter condensed phase products due to sublimation of AlF3. Lastly, a physico-chemical picture of the agglomeration process was also developed based on the results of high-speed CCD video analysis.

Structural Assessment of Solid Propellant Grains (l’Evaluation structurale des blocs de poudre a’ propergol solide)

DTIC Science & Technology

1997-12-01

bonds) This technique is based on the observation of the reflection and attenuation of an ultrasonic wave traversing an object, and is used to check...Nearly all present day composite propellants for tactical rocket motors use hydroxy-terminated polybutadiene ( HTPB ) as a binder as this offers the...polyurethane as a binder. The inferior mechanical properties of these propellants compared to HTPB limited their use. In large space booster and

Multiple-wavelength transmission measurements in rocket motor plumes

NASA Astrophysics Data System (ADS)

Kim, Hong-On

1991-09-01

Multiple-wavelength light transmission measurements were used to measure the mean particle size (d(sub 32)), index of refraction (m), and standard deviation of the small particles in the edge of the plume of a small solid propellant rocket motor. The results have shown that the multiple-wavelength light transmission measurement technique can be used to obtain these variables. The technique was shown to be more sensitive to changes in d(sub 32) and standard deviation (sigma) than to m. A GAP/AP/4.7 percent aluminum propellant burned at 25 atm produced particles with d32 = 0.150 +/- 0.006 microns, standard deviation = 1.50 +/- 0.04 and m = 1.63 +/- 0.13. The good correlation of the data indicated that only submicron particles were present in the edge of the plume. In today's budget conscious industry, the solid propellant rocket motor is an ideal propulsion system due to its low cost and simplicity. The major obstacle for solid rocket motors, however, is their limited specific impulse compared to airbreathing motors. One way to help overcome this limitation is to utilize metal fuel additives. Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Aluminum propellants also increase propellant densities and suppress transverse modes of combustion oscillations by damping the oscillations with the aluminum agglomerates in the combustion chamber.

Flame Temperatures and Internal Pressures of Pyrotechnic Igniters Used in Liquid Propellant Gun Firings

DTIC Science & Technology

1982-03-01

IP AT 655 ~~I . . . . . 45 7 I. INTRODUCTION The lack of quantitative ignition design criteria in liquid propellant gun firings requires the...Meeting~ CPIA PubUaation No. :300~ VoZ . I, AppUed Physias Laboratory~ SiZver Spring~ MD~ p. :39:3 (19?9). 26 REFERENCES 1. J. D. Knapton, I. C. Stobie...T9E6 Igniter and a Booster Charge of M30 and Eimite !I I ll[[l 1!13 IP -111 .. Sll tiiiiML I RRX-P.D. !1252 I ~· s 1: 31~ 211 z II Figure B2

Reduced-Smoke Solid Propellant Combustion Products Analysis. Development of a Micromotor Combustor Technique.

DTIC Science & Technology

1976-10-01

A low-cost micromotor combustor technique has been devised to support the development of reduced-smoke solid propellant formulations. The technique...includes a simple, reusable micromotor capable of high chamber pressures, a combustion products collection system, and procedures for analysis of

76 FR 51459 - Office of Commercial Space Transportation (AST); Notice of Availability of the Record of Decision...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-08-18

... impacts of up to five solid-propellant strap-on rocket motors (SRMs) on the Atlas V medium lift vehicle... Proposed Action in the 2000 SEIS, up to five solid- propellant strap-on rocket motors (SRMs) would be added...

The Delta launch vehicle Model 2914 series

NASA Technical Reports Server (NTRS)

Gunn, C. R.

1973-01-01

Description of a new, medium-class Delta launch-vehicle configuration, the three-stage Model 2914. The first stage of this vehicle is composed of a liquid-propellant core which is thrust-augmented with up to nine strap-on solid-propellant motors. The second stage, recently uprated with a strap-down inertial guidance system, is now being modified to adapt the liquid-propellant descent engine from the Apollo Lunar Excursion Module. The third stage is a spin-stabilized solid-propellant motor. The Model 2914 is capable of injecting 2040 kg into low earth orbit, 705 kg into geosynchronous transfer orbit, or 455 kg into an escape trajectory.

Modal survey of the space shuttle solid rocket motor using multiple input methods

NASA Technical Reports Server (NTRS)

Brillhart, Ralph; Hunt, David L.; Jensen, Brent M.; Mason, Donald R.

1987-01-01

The ability to accurately characterize propellant in a finite element model is a concern of engineers tasked with studying the dynamic response of the Space Shuttle Solid Rocket Motor (SRM). THe uncertainties arising from propellant characterization through specimem testing led to the decision to perform a model survey and model correlation of a single segment of the Shuttle SRM. Multiple input methods were used to excite and define case/propellant modes of both an inert segment and, later, a live propellant segment. These tests were successful at defining highly damped, flexible modes, several pairs of which occured with frequency spacing of less than two percent.

The effects of solid rocket motor effluents on selected surfaces and solid particle size, distribution, and composition for simulated shuttle booster separation motors

NASA Technical Reports Server (NTRS)

Jex, D. W.; Linton, R. C.; Russell, W. M.; Trenkle, J. J.; Wilkes, D. R.

1976-01-01

A series of three tests was conducted using solid rocket propellants to determine the effects a solid rocket plume would have on thermal protective surfaces (TPS). The surfaces tested were those which are baselined for the shuttle vehicle. The propellants used were to simulate the separation solid rocket motors (SSRM) that separate the solid rocket boosters (SRB) from the shuttle launch vehicle. Data cover: (1) the optical effects of the plume environment on spacecraft related surfaces, and (2) the solid particle size, distribution, and composition at TPS sample locations.

Introduction to the problem

NASA Technical Reports Server (NTRS)

Ramohalli, Kumar

1989-01-01

Solid propellant rockets were used extensively in space missions ranging from large boosters to orbit-raising upper stages. The smaller motors find exclusive use in various earth-based applications. The advantage of the solids include simplicity, readiness, volumetric efficiency, and storability. Important recent progress in related fields (combustion, rheology, micro-instrumentation/diagnostics, and chaos theory) can be applied to solid rockets to derive maximum advantage and avoid waste. Main objectives of research in solid propellants include: to identify critical parameters, to establish specification rules, and to develop quantitative criteria.

Design assurance test of the Thiokol Te-M-521-5 apogee kick motor tested in the spin mode at simulated altitude conditions

NASA Technical Reports Server (NTRS)

Cimino, A. A.

1973-01-01

One Thiokol Chemical Corporation TE-M-521-5 solid-propellant apogee rocket motor was successfully fired at an average simulated altitude of about 108,000 ft while spinning at 46 rpm. The general program objectives were to verify compliance of motor performance with the manufacturer's specifications. Specific primary objectives were to determine vacuum ballistic performance of the motor after prefire vibration conditioning and temperature conditioning at 40F, altitude ignition characteristics, motor structural integrity, and motor temperature-time history during and after motor operation. Additional objectives were to measure the lateral (nonaxial) thrust component during motor operation and to measure radiation heat flux in the vicinity of the nozzle exit plane.

Low acid producing solid propellants

NASA Technical Reports Server (NTRS)

Bennett, Robert R.

1995-01-01

The potential environmental effects of the exhaust products of conventional rocket propellants have been assessed by various groups. Areas of concern have included stratospheric ozone, acid rain, toxicity, air quality and global warming. Some of the studies which have been performed on this subject have concluded that while the impacts of rocket use are extremely small, there are propellant development options which have the potential to reduce those impacts even further. This paper discusses the various solid propellant options which have been proposed as being more environmentally benign than current systems by reducing HCI emissions. These options include acid neutralized, acid scavenged, and nonchlorine propellants. An assessment of the acid reducing potential and the viability of each of these options is made, based on current information. Such an assessment is needed in order to judge whether the potential improvements justify the expenditures of developing the new propellant systems.

Propellant grain dynamics in aft attach ring of shuttle solid rocket booster

NASA Technical Reports Server (NTRS)

Verderaime, V.

1979-01-01

An analytical technique for implementing simultaneously the temperature, dynamic strain, real modulus, and frequency properties of solid propellant in an unsymmetrical vibrating ring mode is presented. All dynamic parameters and sources are defined for a free vibrating ring-grain structure with initial displacement and related to a forced vibrating system to determine the change in real modulus. Propellant test data application is discussed. The technique was developed to determine the aft attach ring stiffness of the shuttle booster at lift-off.

Seismic vibration source

NASA Technical Reports Server (NTRS)

Dowler, W. L.; Varsi, G.; Yang, L. C. (Inventor)

1979-01-01

A system for vibrating the earth in a location where seismic mapping is to take place is described. A relatively shallow hole formed in the earth, such as a hole 10 feet deep, placing a solid propellant in the hole, sealing a portion of the hole above the solid propellant with a device that can rapidly open and close to allow a repeatedly interrupted escape of gas. The propellant is ignited so that high pressure gas is created which escapes in pulses to vibrate the earth.

Assessment of analytical techniques for predicting solid propellant exhaust plumes and plume impingement environments

NASA Technical Reports Server (NTRS)

Tevepaugh, J. A.; Smith, S. D.; Penny, M. M.

1977-01-01

An analysis of experimental nozzle, exhaust plume, and exhaust plume impingement data is presented. The data were obtained for subscale solid propellant motors with propellant Al loadings of 2, 10 and 15% exhausting to simulated altitudes of 50,000, 100,000 and 112,000 ft. Analytical predictions were made using a fully coupled two-phase method of characteristics numerical solution and a technique for defining thermal and pressure environments experienced by bodies immersed in two-phase exhaust plumes.

Development and Testing of a Green-Propellant Micro-Hybrid Thruster with Electrostatic Ignition

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Judson, Michael D.

2012-01-01

As early as 1937 German scientists at Peenemunde experimented with highly unstable fuel blends of nitrous oxide (N2O) and ethanol. These early tests mostly resulted in explosions and destroyed rocket engines. More recently several companies have developed experimental nitrous oxide fuel blends (NOFB) with Isp exceeding 300 sec. Although NOFBx has recently been cleared for tests on the International Space Station, this propellant remains highly experimental and has not been cleared for commercial transport by the US DOT. Recent work by Karabeyoglu et al. has raised concerns about the safety risks of mixing hydrocarbons with N2O. Liquid oxidizer/fuel blends are highly explosive and require extreme care in transport and servicing. By adding small amounts of a liquid organic fuel such as alcohol or a hydrocarbon, the odds of an explosive decomposition event are significantly increased.iv The proposed solution mitigates the explosion hazards of NOFB by separating the oxidizer from the hydrocarbon fuel formed as of a small cylindrical section of ABS thermoplastic. As N2O vapor flows across the grain segment, current enters a 1000 VDC high-tension lead in the ABS fuel grain and produces an inductive spark that vaporizes a small amount of the material. The ablated fuel vapor plus residual energy from the spark seed a localized exothermic N2O dissociation that produces sufficient heat to initiate combustion. The process is also effective when gaseous oxygen is used. A low TRL (2-3) prototype demonstrating the feasibility of controlled hydrocarbon-seeding was recently tested at Utah State University.v The unit features a miniature 2.5 cm ABS fuel grain fabricated using a Stratasys Dimension 3-D printer. The 9-N thruster was pulse-fired up to 27 consecutive times on a single ABS grain segment. Ignition was achieved by as little as 12-15 Joules energy input. This value is contrasted with the typical 30-minute pre-heat requirement for the ECAPS LMP-103S ADN-based monopropellant, requiring an energy input of 14,850 Joules for catalytic dissociation. The hydrocarbon-seeded micro-hybrid was also adapted as a non-pyrotechnic ignitor for a 900 N (200-lbf) thrust hybrid motor. The motor was successfully ignited 4 consecutive times with no hardware swaps or propellant additions. The amount of ABS seed material that can be fit into the injector cap is the only limit to the number of available repeat firings. This series of tests marks the first time a hybrid motor was ever ignited by other than a solid-propellant pyrotechnic charge or bi-propellant flame ignitor. Nitrous oxide hybrid motors are typically difficult to ignite and usually require multiple solid-propellant charges to initiate combustion, so this nonpyrotechnic ignition is a significant accomplishment. The controlled hydrocarbon-seeding approach is fundamentally different from all other green propellant solutions offered by the aerospace industry. Although the proposed system is more correctly a hybrid technology; the system retains all the simple features of a monopropellant design. To date no optimization study has been performed to identify the best grain geometry for electrostatic ignition. Fortunately, because the grain segments are fabricated using rapid-prototyping technology, changing the grain geometry is as simple as modifying the 3-D printer CAD-file. Vacuum Isp exceeding 270 seconds has been demonstrated (Ref v), a value significantly higher than those offered by competing green monopropellant options. The propellants of choice, N2O/GOX and ABS are 100% non-toxic, non-explosive, and environmentally benign. Because the inert oxidizer and fuel components are mixed only within the combustion chamber, the system retains the inherent safety of a hybrid rocket and can be piggy-backed as a secondary payload with no overall mission risk increase to the primary payload, an excellent characteristic for secondary launch systems.

Space shuttle propellant constitutive law verification tests

NASA Technical Reports Server (NTRS)

Thompson, James R.

1995-01-01

As part of the Propellants Task (Task 2.0) on the Solid Propulsion Integrity Program (SPIP), a database of material properties was generated for the Space Shuttle Redesigned Solid Rocket Motor (RSRM) PBAN-based propellant. A parallel effort on the Propellants Task was the generation of an improved constitutive theory for the PBAN propellant suitable for use in a finite element analysis (FEA) of the RSRM. The outcome of an analysis with the improved constitutive theory would be more reliable prediction of structural margins of safety. The work described in this report was performed by Materials Laboratory personnel at Thiokol Corporation/Huntsville Division under NASA contract NAS8-39619, Mod. 3. The report documents the test procedures for the refinement and verification tests for the improved Space Shuttle RSRM propellant material model, and summarizes the resulting test data. TP-H1148 propellant obtained from mix E660411 (manufactured February 1989) which had experienced ambient igloo storage in Huntsville, Alabama since January 1990, was used for these tests.

Development of a solid propellant viscoelastic dynamic model

NASA Technical Reports Server (NTRS)

Hufferd, W. L.; Fitzgerald, J. E.

1976-01-01

The results of a one year study to develop a dynamic response model for the Space Shuttle Solid Rocket Motor (SRM) propellant are presented. An extensive literature survey was conducted, from which it was concluded that the only significant variables affecting the dynamic response of the SRM propellant are temperature and frequency. Based on this study, and experimental data on propellants related to the SRM propellant, a dynamic constitutive model was developed in the form of a simple power law with temperature incorporated in the form of a modified power law. A computer program was generated which performs a least-squares curve-fit of laboratory data to determine the model parameters and it calculates dynamic moduli at any desired temperature and frequency. Additional studies investigated dynamic scaling laws and the extent of coupling between the SRM propellant and motor cases. It was found, in agreement with other investigations, that the propellant provides all of the mass and damping characteristics whereas the case provides all of the stiffness.

Erosive Burning of Composite Solid Propellants: Experimental and Modeling Studies

DTIC Science & Technology

1978-08-01

of Crossflow on Solid Pro- appears that an additional mechanism(s) of erosive pallant Combustion: Interior Ballistic Design burning will have to be...Orlondo, Florida, July , 1977, AIAA Paper 77-930. 14. Lengelle,G., "Model Describing the Erosive Com- bustion and Velocity Response of Composite Pro...Propulsion Conference, Orlando, Florida, July , 1977. 17. Beddini, R.A., A Reacting Turbulent Boundary Layer Approach to Solid Propellant Erosive Burning, AFOSR

Locating spatial variation in the association between wildland fire risk and social vulnerability across six southern states

Treesearch

Neelam Poudyal; Cassandra Johnson Gaither; Scott Goodrick; J.M. Bowker; Jianbang Gan

2012-01-01

Wildland fire in the South commands considerable attention, given the expanding wildland urban interface (WUI) across the region. Much of this growth is propelled by higher income retirees and others desiring natural amenity residential settings. However, population growth in the WUI increases the likelihood of wildfire fire ignition caused by people, as humans account...

Modeling of gun barrel surface erosion: Historic perspective

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

Buckingham, A.C.

1996-08-01

Results and interpretations of numerical simulations of some dominant processes influencing gun barrel propellant combustion and flow-induced erosion are presented. Results include modeled influences of erosion reduction techniques such as solid additives, vapor phase chemical modifications, and alteration of surface solid composition through use of thin coatings. Precedents and historical perspective are provided with predictions from traditional interior ballistics compared to computer simulations. Accelerating reactive combustion flow, multiphase and multicomponent transport, flow-to-surface thermal/momentum/phase change/gas-surface chemical exchanges, surface and micro-depth subsurface heating/stress/composition evolution and their roles in inducing surface cracking, spall, ablation, melting, and vaporization are considered. Recognition is given tomore » cyclic effects of previous firing history on material preconditioning. Current perspective and outlook for future are based on results of a US Army-LLNL erosion research program covering 7 y in late 1970s. This is supplemented by more recent research on hypervelocity electromagnetic projectile launchers.« less

Space Launch System Base Heating Test: Experimental Operations & Results

NASA Technical Reports Server (NTRS)

Dufrene, Aaron; Mehta, Manish; MacLean, Matthew; Seaford, Mark; Holden, Michael

2016-01-01

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Test methodology and conditions are presented, and base heating results from 76 runs are reported in non-dimensional form. Regions of high heating are identified and comparisons of various configuration and conditions are highlighted. Base pressure and radiometer results are also reported.

Fluid dynamics of the unsteady two phase processes leading to DDT in granular solid propellants

NASA Technical Reports Server (NTRS)

Krier, H.; Butler, P. B.; Lembeck, M. F.

1980-01-01

Deflagration to Detonation (DDT) was predicted to occur in porous beds of high-energy solid propellants by solving the unsteady fluid mechanical convective heat transfer from hot gas products, obtained from the rapid burning at high pressures, provides the impetus to develop a narrow combustion zone and a resulting strong shock. A parametric study clearly indicates that DDT occurs only when a combination of the solids loading fraction, the burning rate constants, the propellant chemical energy, and the particle size provide for critical energy and gas release to support a detonation wave. Predictions for the run-up length to detonation as a function of these parameters are presented.

Assessment of analytical techniques for predicting solid propellant exhaust plumes

NASA Technical Reports Server (NTRS)

Tevepaugh, J. A.; Smith, S. D.; Penny, M. M.

1977-01-01

The calculation of solid propellant exhaust plume flow fields is addressed. Two major areas covered are: (1) the applicability of empirical data currently available to define particle drag coefficients, heat transfer coefficients, mean particle size and particle size distributions, and (2) thermochemical modeling of the gaseous phase of the flow field. Comparisons of experimentally measured and analytically predicted data are made. The experimental data were obtained for subscale solid propellant motors with aluminum loadings of 2, 10 and 15%. Analytical predictions were made using a fully coupled two-phase numerical solution. Data comparisons will be presented for radial distributions at plume axial stations of 5, 12, 16 and 20 diameters.

Launch Vehicles

NASA Image and Video Library

2007-09-09

Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA's Constellation Program. This transportation system will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. The Ares I effort includes multiple project element teams at NASA centers and contract organizations around the nation, and is managed by the Exploration Launch Projects Office at NASA's Marshall Space Flight Center (MFSC). ATK Launch Systems near Brigham City, Utah, is the prime contractor for the first stage booster. ATK's subcontractor, United Space Alliance of Houston, is designing, developing and testing the parachutes at its facilities at NASA's Kennedy Space Center in Florida. NASA's Johnson Space Center in Houston hosts the Constellation Program and Orion Crew Capsule Project Office and provides test instrumentation and support personnel. Together, these teams are developing vehicle hardware, evolving proven technologies, and testing components and systems. Their work builds on powerful, reliable space shuttle propulsion elements and nearly a half-century of NASA space flight experience and technological advances. Ares I is an inline, two-stage rocket configuration topped by the Crew Exploration Vehicle, its service module, and a launch abort system. The launch vehicle's first stage is a single, five-segment reusable solid rocket booster derived from the Space Shuttle Program's reusable solid rocket motor that burns a specially formulated and shaped solid propellant called polybutadiene acrylonitrile (PBAN). The second or upper stage will be propelled by a J-2X main engine fueled with liquid oxygen and liquid hydrogen. This HD video image depicts a test firing of a 40k subscale J2X injector at MSFC's test stand 115. (Highest resolution available)

Use of Atomic Fuels for Rocket-Powered Launch Vehicles Analyzed

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1999-01-01

At the NASA Lewis Research Center, the launch vehicle gross lift-off weight (GLOW) was analyzed for solid particle feed systems that use high-energy density atomic propellants (ref. 1). The analyses covered several propellant combinations, including atoms of aluminum, boron, carbon, and hydrogen stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents for the liquid carrier were investigated, and the GLOW values of vehicles using the solid particle feed systems were compared with that of a conventional oxygen/hydrogen (O2/H2) propellant vehicle. Atomic propellants, such as boron, carbon, and hydrogen, have an enormous potential for high specific impulse Isp operation, and their pursuit has been a topic of great interest for decades. Recent and continuing advances in the understanding of matter, the development of new technologies for simulating matter at its most basic level, and manipulations of matter through microtechnology and nanotechnology will no doubt create a bright future for atomic propellants and an exciting one for the researchers exploring this technology.

Modeling of Nonlinear Combustion Instability in Solid Propellant Rocket Motors

DTIC Science & Technology

1984-02-01

34. .. .°. .., . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . .... . . . . ..°.... . .°-""... ’o.’ . . °o: :--, - .:" . "" . °° - - 54. Flandro , 0. A., "Solid Propellant Acoustic Admittance...such as those due to Gary , 2 1) Gourlay and Morris ( 2 2 ) and Mas- (23)son are more involved, both from a program development, and computational

Development of high temperature materials for solid propellant rocket nozzle applications

NASA Technical Reports Server (NTRS)

Manning, C. R., Jr.; Lineback, L. D.

1974-01-01

Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

14 CFR 91.815 - Agricultural and fire fighting airplanes: Noise operating limitations.

Code of Federal Regulations, 2014 CFR

2014-01-01

...: Noise operating limitations. 91.815 Section 91.815 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... RULES Operating Noise Limits § 91.815 Agricultural and fire fighting airplanes: Noise operating limitations. (a) This section applies to propeller-driven, small airplanes having standard airworthiness...

Hazard Studies for Solid Propellant Rocket Motors (Etude des Risque pour les Moteurs-Fusees a Propergols Solides)

DTIC Science & Technology

1990-09-01

RESEARCH AND DEVELOPMENT (ORGANISATION DU TRAITE DE LATIANTIOUF NORD) AGARDograph No.3 16 Hazard Studies for Solid Propellant Rocket Motors (Etudes de...member nations to use their research and development capabilities for the common benefit of the NATO community; - Providing scientific and technical...advice and assistance to the Military Committee in the field of aerospace research and development (with particular regard to its military application

Process for the leaching of AP from propellant

NASA Technical Reports Server (NTRS)

Shaw, G. C.; Mcintosh, M. J. (Inventor)

1980-01-01

A method for the recovery of ammonium perchlorate from waste solid rocket propellant is described wherein shredded particles of the propellant are leached with an aqueous leach solution containing a low concentration of surface active agent while stirring the suspension.

Survivability of a Propellant Fire inside a Simulated Military Vehicle Crew Compartment: Part 2 - Hazard Mitigation Strategies and Their Effectiveness

DTIC Science & Technology

2013-06-01

Weapons Propulsion Group where his work initially focussed on R&D relating to cast- composite rocket motors. The emphasis of his work then shifted to gun...Relative humidity RHS Rectangular Hollow Section t Time (s) T1 Ambient room temperature, ceiling-height (K) T2 Ambient room temperature...propellant and a centre- core igniter train. The BCM and UNCLASSIFIED DSTO-RR-0393 UNCLASSIFIED 2 TCM contain the same propellant formulation and

KENNEDY SPACE CENTER, FLA. - Seen from below and through a solid rocket booster segment mockup, Jeff Thon, an SRB mechanic with United Space Alliance, tests the feasibility of a vertical solid rocket booster propellant grain inspection technique. The inspection of segments is required as part of safety analysis.

NASA Image and Video Library

2003-09-11

KENNEDY SPACE CENTER, FLA. - Seen from below and through a solid rocket booster segment mockup, Jeff Thon, an SRB mechanic with United Space Alliance, tests the feasibility of a vertical solid rocket booster propellant grain inspection technique. The inspection of segments is required as part of safety analysis.

Ultrasonic method for inspection of the propellant grain in the space shuttle solid rocket booster

NASA Astrophysics Data System (ADS)

Doyle, T. E.; Degtyar, A. D.; Sorensen, K. P.; Kelso, M. J.; Berger, T. A.

2000-05-01

Defects in solid rocket propellant may affect the safe operation of a space launch vehicle. The Space Shuttle reusable solid rocket motor (RSRM) is therefore routinely inspected with radiography for voids, cracks, and inclusions. Ultrasonic methods can be used to supplement radiography when an indication is difficult to interpret due to the projection geometry or low contrast. Such a method was developed to inspect a local region of propellant in an RSRM forward segment for a suspect inclusion. The method used a through-transmission approach, with a stationary transmitter on the propellant grain inside the segment and a receiving transducer scanned over the case surface. Low frequency (⩽250 kHz) pulses were propagated through 10-12 inches of propellant, 0.5 inches of NBR insulation, and 0.5 inches of steel case. Through-transmission images were constructed using time-of-flight analysis of the waveforms. The ultrasonic inspections supported results from extended radiographic studies, showing that the indication was not an inclusion but an artifact resulting from liner thickness variations and a low X-ray projection angle in the segment's dome region. This work demonstrated the feasibility of using ultrasonics for inspection of propellant grain in steel-cased rocket motors.

Characterization of Emissions from Liquid Fuel and Propane Open Burns

EPA Science Inventory

The comparative combustion emissions of using jet propellant (JP-5) liquid fuel pools or a propane manifold grid to simulate the effects of accidental fires was investigated. A helium-filled tethered aerostat was used to maneuver an instrument package into the open fire plumes ...

14 CFR 36.1583 - Noncomplying agricultural and fire fighting airplanes.

Code of Federal Regulations, 2013 CFR

2013-01-01

... airplanes. 36.1583 Section 36.1583 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... Limitations and Information § 36.1583 Noncomplying agricultural and fire fighting airplanes. (a) This section applies to propeller-driven, small airplanes that— (1) Are designed for “agricultural aircraft operations...

14 CFR 36.1583 - Noncomplying agricultural and fire fighting airplanes.

Code of Federal Regulations, 2011 CFR

2011-01-01

... airplanes. 36.1583 Section 36.1583 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... Limitations and Information § 36.1583 Noncomplying agricultural and fire fighting airplanes. (a) This section applies to propeller-driven, small airplanes that— (1) Are designed for “agricultural aircraft operations...

14 CFR 36.1583 - Noncomplying agricultural and fire fighting airplanes.

Code of Federal Regulations, 2012 CFR

2012-01-01

... airplanes. 36.1583 Section 36.1583 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... Limitations and Information § 36.1583 Noncomplying agricultural and fire fighting airplanes. (a) This section applies to propeller-driven, small airplanes that— (1) Are designed for “agricultural aircraft operations...

14 CFR 36.1583 - Noncomplying agricultural and fire fighting airplanes.

Code of Federal Regulations, 2014 CFR

2014-01-01

... airplanes. 36.1583 Section 36.1583 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... Limitations and Information § 36.1583 Noncomplying agricultural and fire fighting airplanes. (a) This section applies to propeller-driven, small airplanes that— (1) Are designed for “agricultural aircraft operations...

14 CFR 36.1583 - Noncomplying agricultural and fire fighting airplanes.

Code of Federal Regulations, 2010 CFR

2010-01-01

... airplanes. 36.1583 Section 36.1583 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... Limitations and Information § 36.1583 Noncomplying agricultural and fire fighting airplanes. (a) This section applies to propeller-driven, small airplanes that— (1) Are designed for “agricultural aircraft operations...

Experimental investigation of solid rocket motors for small sounding rockets

NASA Astrophysics Data System (ADS)

Suksila, Thada

2018-01-01

Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

U.S. Army Armament Research, Development and Engineering Center Grain Evaluation Software to Numerically Predict Linear Burn Regression for Solid Propellant Grain Geometries

DTIC Science & Technology

2017-10-01

ENGINEERING CENTER GRAIN EVALUATION SOFTWARE TO NUMERICALLY PREDICT LINEAR BURN REGRESSION FOR SOLID PROPELLANT GRAIN GEOMETRIES Brian...author(s) and should not be construed as an official Department of the Army position, policy, or decision, unless so designated by other documentation...U.S. ARMY ARMAMENT RESEARCH, DEVELOPMENT AND ENGINEERING CENTER GRAIN EVALUATION SOFTWARE TO NUMERICALLY PREDICT LINEAR BURN REGRESSION FOR SOLID

Solid Propellant Nonlinear Constitutive Theory Extension

DTIC Science & Technology

1984-01-01

Force Rocket Propulsion Laboratory, June 1979. Farris, R. J., Hermann , I. R., Hutchinson, J. R., and Schapery, R. A., "Development of a Solid Rocket...Effect of Stretching on the Properties of Rubber," J. Rub. Res., 16, 275-289, 1947. 28. Oberth , A. E., and Brenner, R. S., "Tear Phenomena Around...34Development of a Solid Rocket Propellant Nonlinear Viscoelastic Constitutive Theory," AFRPL-TR-73-50, June 1973. 30. Hermann , L. R., and Peterson, F. E., "A

AFRL Solid Propellant Laboratory Explosive Siting and Renovation Lessons Learned

DTIC Science & Technology

2010-05-19

AFRL Solid Propellant Laboratory Explosive Siting and Renovation Lessons Learned Daniel F. Schwartz Air Force Research Laboratory ...9. SPONSORING / MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSOR/MONITOR’S ACRONYM(S) Air Force Research Laboratory (AFMC) AFRL /RZS...provide the United States Air Force with advanced rocket propulsion technologies, the Air Force Research

Determination of solid-propellant transient regression rates using a microwave Doppler shift technique

NASA Technical Reports Server (NTRS)

Strand, L. D.; Schultz, A. L.; Reedy, G. K.

1972-01-01

A microwave Doppler shift system, with increased resolution over earlier microwave techniques, was developed for the purpose of measuring the regression rates of solid propellants during rapid pressure transients. A continuous microwave beam is transmitted to the base of a burning propellant sample cast in a metal waveguide tube. A portion of the wave is reflected from the regressing propellant-flame zone interface. The phase angle difference between the incident and reflected signals and its time differential are continuously measured using a high resolution microwave network analyzer and related instrumentation. The apparent propellant regression rate is directly proportional to this latter differential measurement. Experiments were conducted to verify the (1) spatial and time resolution of the system, (2) effect of propellant surface irregularities and compressibility on the measurements, and (3) accuracy of the system for quasi-steady-state regression rate measurements. The microwave system was also used in two different transient combustion experiments: in a rapid depressurization bomb, and in the high-frequency acoustic pressure environment of a T-burner.

Propellant Readiness Level: A Methodological Approach to Propellant Characterization

NASA Technical Reports Server (NTRS)

Bossard, John A.; Rhys, Noah O.

2010-01-01

A methodological approach to defining propellant characterization is presented. The method is based on the well-established Technology Readiness Level nomenclature. This approach establishes the Propellant Readiness Level as a metric for ascertaining the readiness of a propellant or a propellant combination by evaluating the following set of propellant characteristics: thermodynamic data, toxicity, applications, combustion data, heat transfer data, material compatibility, analytical prediction modeling, injector/chamber geometry, pressurization, ignition, combustion stability, system storability, qualification testing, and flight capability. The methodology is meant to be applicable to all propellants or propellant combinations; liquid, solid, and gaseous propellants as well as monopropellants and propellant combinations are equally served. The functionality of the proposed approach is tested through the evaluation and comparison of an example set of hydrocarbon fuels.

14 CFR 25.903 - Engines.

Code of Federal Regulations, 2010 CFR

2010-01-01

... chapter. (2) Each turbine engine must comply with one of the following: (i) Sections 33.76, 33.77 and 33... any engine individually in flight, except that, for turbine engine installations, the means for... might be exposed to fire must be at least fire-resistant. If hydraulic propeller feathering systems are...

Feasibility of an advanced thrust termination assembly for a solid propellant rocket motor

NASA Technical Reports Server (NTRS)

1975-01-01

A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.

Combustion chemistry of solid propellants

NASA Technical Reports Server (NTRS)

Baer, A. D.; Ryan, N. W.

1974-01-01

Several studies are described of the chemistry of solid propellant combustion which employed a fast-scanning optical spectrometer. Expanded abstracts are presented for four of the studies which were previously reported. One study of the ignition of composite propellants yielded data which suggested early ammonium perchlorate decomposition and reaction. The results of a study of the spatial distribution of molecular species in flames from uncatalyzed and copper or lead catalyzed double-based propellants support previously published conclusions concerning the site of action of these metal catalysts. A study of the ammonium-perchlorate-polymeric-fuel-binder reaction in thin films, made by use of infrared absorption spectrometry, yielded a characterization of a rapid condensed-phase reaction which is likely important during the ignition transient and the burning process.

Evaluation of weapons' combustion products in armored vehicles. Final report, 30 September 1986-14 December 1988

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

Menzies, K.T.; Randel, M.A.; Quill, A.L.

1989-01-01

The U.S. Army Biomedical Research and Development Laboratory defined an extensive research program to address the generation of potentially toxic propellant combustion products in crew compartments of armored vehicles during weapons firing. The major objectives of the research were: (1) to determine the presence and concentration of propellant combustion products, (2) to determine potential crew exposure to these combustion products, and (3) to assess the efficacy of field monitoring in armored vehicles. To achieve these goals, air monitoring was conducted in selected armored vehicle types, i.e., M109, M60, M3, M1, at several Army installations. Auxiliary information concerning the specific munitionsmore » fired and the Training and Doctrine Command (TRADOC) or Forces Command (FORSCOM) firing scenarios was collected so that a comparison of pollutant concentrations generated by specific weapons both within vehicle types and between vehicle types could be made.« less

Safety issues of high-concentrated hydrogen peroxide production used as rocket propellant

NASA Astrophysics Data System (ADS)

Romantsova, O. V.; Ulybin, V. B.

2015-04-01

The article dwells on the possibility of production of high-concentrated hydrogen peroxide with the Russian technology of isopropyl alcohol autoxidation. Analysis of fire/explosion hazards and reasons of insufficient quality is conducted for the technology. Modified technology is shown. Non-standard fire/explosion characteristics required for integrated fire/explosion hazards rating for modified hydrogen peroxide production based on the autoxidation of isopropyl alcohol are defined.

KENNEDY SPACE CENTER, FLA. - Jeff Thon, an SRB mechanic with United Space Alliance, tests a technique for vertical solid rocket booster propellant grain inspection. The inspection of segments is required as part of safety analysis.

NASA Image and Video Library

2003-09-11

KENNEDY SPACE CENTER, FLA. - Jeff Thon, an SRB mechanic with United Space Alliance, tests a technique for vertical solid rocket booster propellant grain inspection. The inspection of segments is required as part of safety analysis.

The 17th JANNAF Combustion Meeting, Volume 1

NASA Technical Reports Server (NTRS)

Eggleston, D. S. (Editor)

1980-01-01

The combustion of solid rocket propellants and combustion in ramjets is addressed. Subjects discussed include metal burning, steady-state combustion of composite propellants, velocity coupling and nonlinear instability, vortex shedding and flow effects on combustion instability, combustion instability in solid rocket motors, combustion diagnostics, subsonic and supersonic ramjet combustion, characterization of ramburner flowfields, and injection and combustion of ramjet fuels.

Stable isotopes of nitrate reflect natural attenuation of propellant residues on military training ranges.

PubMed

Bordeleau, Geneviève; Savard, Martine M; Martel, Richard; Smirnoff, Anna; Ampleman, Guy; Thiboutot, Sonia

2013-08-06

Nitroglycerin (NG) and nitrocellulose (NC) are constituents of double-base propellants used notably for firing antitank ammunitions. Nitroglycerin was detected in soil and water samples from the unsaturated zone (pore water) at an active antitank firing position, where the presence of high nitrate (NO3(-)) concentrations suggests that natural attenuation of NG is occurring. However, concentrations alone cannot assess if NG is the source of NO3(-), nor can they determine which degradation processes are involved. To address this issue, isotopic ratios (δ(15)N, δ(18)O) were measured for NO3(-) produced from NG and NC through various controlled degradation processes and compared with ratios measured in field pore water samples. Results indicate that propellant combustion and degradation mediated by soil organic carbon produced the observed NO3(-) in pore water at this site. Moreover, isotopic results are presented for NO3(-) produced through photolysis of propellant constituents, which could be a dominant process at other sites. The isotopic data presented here constitute novel information regarding a source of NO3(-) that was practically not documented before and a basis to study the contamination by energetic materials in different contexts.

Hypervelocity cutting machine and method

DOEpatents

Powell, J.R.; Reich, M.

1996-11-12

A method and machine are provided for cutting a workpiece such as concrete. A gun barrel is provided for repetitively loading projectiles therein and is supplied with a pressurized propellant from a storage tank. A thermal storage tank is disposed between the propellant storage tank and the gun barrel for repetitively receiving and heating propellant charges which are released in the gun barrel for repetitively firing projectiles therefrom toward the workpiece. In a preferred embodiment, hypervelocity of the projectiles is obtained for cutting the concrete workpiece by fracturing thereof. 10 figs.

Extension of a simplified computer program for analysis of solid-propellant rocket motors

NASA Technical Reports Server (NTRS)

Sforzini, R. H.

1973-01-01

A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

NASA Technical Reports Server (NTRS)

Meyer, Mike L.

1993-01-01

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

Polar Satellite Launch Vehicle (PSLV) development programme in India

NASA Astrophysics Data System (ADS)

Janardhana, E.

The design of the Indian Polar Satellite Launch Vehicle (PSLV), for the launching (by 1990) of 1-1.5-tonne payloads into 900-km sun-synchronous orbit, is discussed, and the mission development program is described. The first stage is a solid propellant motor augmented by six solid strap-ons, and the second stage of liquid storable propellant has a high thrust gimballed engine. A high performance solid motor incorporates a flex nozzle for control as the third stage, and the fourth stage is a liquid propulsion system using N204 and MMH propellant with two regeneratively cooled engines. The vehicle equipment bay, housing the inertial guidance and control system, and the TTC system are located around the fourth stage for guidance and tracking with the associated ground segment until spacecraft ejection into orbit.

Space Storable Propellant Performance Gas/Liquid Like-Doublet Injector Characterization

NASA Technical Reports Server (NTRS)

Falk, A. Y.

1972-01-01

A 30-month applied research program was conducted, encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space-storable propellants. The gas/liquid propellant combination selected for study was FLOX (82.6% F2)/ambient temperature gaseous methane. The injector pattern characterized was the like-(self)-impinging doublet. Program effort was apportioned into four basic technical tasks: injector and thrust chamber design, injector and thrust chamber fabrication, performance evaluation testing, and data evaluation and reporting. Analytical parametric combustion analyses and cold flow distribution and atomization experiments were conducted with injector segment models to support design of injector/thrust chamber combinations for hot fire evaluation. Hot fire tests were conducted to: (1) optimize performance of the injector core elements, and (2) provide design criteria for the outer zone elements so that injector/thrust chamber compatibility could be achieved with only minimal performance losses.

46 CFR 167.45-45 - Carbon dioxide fire-extinguishing system requirements.

Code of Federal Regulations, 2010 CFR

2010-10-01

... SCHOOLS PUBLIC NAUTICAL SCHOOL SHIPS Special Firefighting and Fire Prevention Requirements § 167.45-45... school ship propelled by internal combustion engines, the quantity of carbon dioxide required may be... arrangement of the piping shall be such as to give a general and fairly uniform distribution over the entire...

46 CFR 167.45-45 - Carbon dioxide fire-extinguishing system requirements.

Code of Federal Regulations, 2011 CFR

2011-10-01

... SCHOOLS PUBLIC NAUTICAL SCHOOL SHIPS Special Firefighting and Fire Prevention Requirements § 167.45-45... school ship propelled by internal combustion engines, the quantity of carbon dioxide required may be... arrangement of the piping shall be such as to give a general and fairly uniform distribution over the entire...

Fundamental Understanding of Propellant/Nozzle Interaction for Rocket Nozzle Erosion Minimization Under Very High Pressure Conditions

DTIC Science & Technology

2005-08-31

conditions; with X-ray radiography for erosion rate measurements. A vortex combustor was also designed to simulate propellant product species and to...DATES COVERED Interim Progress Report, August 1, 2004 to July 31, 2005 4. TITLE AND SUBTITLE Fundamental Understanding of Propellant /Nozzle...nozzle erosion by solid- propellant combustion products. Several processes can affect the nozzle erosion rate at high pressure and temperature

Environmentally compatible solid rocket propellants

NASA Technical Reports Server (NTRS)

Jacox, James L.; Bradford, Daniel J.

1995-01-01

Hercules' clean propellant development research is exploring three major types of clean propellant: (1) chloride-free formulations (no chlorine containing ingredients), being developed on the Clean Propellant Development and Demonstration (CPDD) contract sponsored by Phillips Laboratory, Edwards Air Force Base, CA; (2) low HCl scavenged formulations (HCl-scavenger added to propellant oxidized with ammonium perchlorate (AP)); and (3) low HCl formulations oxidized with a combination of AN and AP (with or without an HCl scavenger) to provide a significant reduction (relative to current solid rocket boosters) in exhaust HCl. These propellants provide performance approaching that of current systems, with less than 2 percent HCl in the exhaust, a significant reduction (greater than or equal to 70 percent) in exhaust HCl levels. Excellent processing, safety, and mechanical properties were achieved using only readily available, low cost ingredients. Two formulations, a sodium nitrate (NaNO3) scavenged HTPB and a chloride-free hydroxy terminated polyether (HTPE) propellant, were characterized for ballistic, mechanical, and rheological properties. In addition, the hazards properties were demonstrated to provide two families of class 1.3, 'zero-card' propellants. Further characterization is planned which includes demonstration of ballistic tailorability in subscale (one to 70 pound) motors over the range of burn rates required for retrofit into current Hercules space booster designs (Titan 4 SRMU and Delta 2 GEM).

A Study on New Composite Thermoplastic Propellant

NASA Astrophysics Data System (ADS)

Kahara, Takehiro; Nakayama, Masanobu; Hasegawa, Hiroshi; Katoh, Kazushige; Miyazaki, Shigehumi; Maruizumi, Haruki; Hori, Keiichi; Morita, Yasuhiro; Akiba, Ryojiro

Efforts have been paid to realize a new composite propellant using thermoplastics as a fuel binder and lithium as a metallic fuel. Thermoplastics binder makes it possible the storage of solid propellant in small blocks and to provide propellants blocks into rocket motor case at a quantity needed just before use, which enables the production facility of solid propellant at a minimum level, thus, production cost significantly lower. Lithium has been a candidate for a metallic fuel for the ammonium perchlorate based composite propellants owing to its capability to reduce the hydrogen chloride in the exhaust gas, however, never been used because lithium is not stable at room conditions and complex reaction products between oxygen, nitrogen, and water are formed at the surface of particles and even in the core. However, lithium particles whose surface shell structure is well controlled are rather stable and can be stored in thermoplastics for a long period. Evaluation of several organic thermoplastics whose melting temperatures are easily tractable was made from the standpoint of combustion characteristics, and it is shown that thermoplastics propellants can cover wide range of burning rate spectrum. Formation of well-defined surface shell of lithium particles and its kinetics are also discussed.

Failure Investigation of an Intra-Manifold Explosion in a Horizontally-Mounted 870 lbf Reaction Control Thruster

NASA Technical Reports Server (NTRS)

Durning, Joseph G., III; Westover, Shayne C.; Cone, Darren M.

2011-01-01

In June 2010, an 870 lbf Space Shuttle Orbiter Reaction Control System Primary Thruster experienced an unintended shutdown during a test being performed at the NASA White Sands Test Facility. Subsequent removal and inspection of the thruster revealed permanent deformation and misalignment of the thruster valve mounting plate. Destructive evaluation determined that after three nominal firing sequences, the thruster had experienced an energetic event within the fuel (monomethylhydrazine) manifold at the start of the fourth firing sequence. The current understanding of the phenomenon of intra-manifold explosions in hypergolic bipropellant thrusters is documented in literature where it is colloquially referred to as a ZOT. The typical ZOT scenario involves operation of a thruster in a gravitational field with environmental pressures above the triple point pressure of the propellants. Post-firing, when the thruster valves are commanded closed, there remains a residual quantity of propellant in both the fuel and oxidizer (nitrogen tetroxide) injector manifolds known as the "dribble volume". In an ambient ground test configuration, these propellant volumes will drain from the injector manifolds but are impeded by the local atmospheric pressure. The evacuation of propellants from the thruster injector manifolds relies on the fluids vapor pressure to expel the liquid. The higher vapor pressure oxidizer will evacuate from the manifold before the lower vapor pressure fuel. The localized cooling resulting from the oxidizer boiling during manifold draining can result in fuel vapor migration and condensation in the oxidizer passage. The liquid fuel will then react with the oxidizer that enters the manifold during the next firing and may produce a localized high pressure reaction or explosion within the confines of the oxidizer injector manifold. The typical ZOT scenario was considered during this failure investigation, but was ultimately ruled out as a cause of the explosion. Converse to the typical ZOT failure mechanism, the failure of this particular thruster was determined to be the result of liquid oxidizer being present within the fuel manifold.

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

NASA Technical Reports Server (NTRS)

Magill, B. T.; Nauflett, G. W.; Furrow, K. W.

2000-01-01

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.

Hypervelocity Launching and Frozen Fuels as a Major Contribution to Spaceflight

NASA Astrophysics Data System (ADS)

Cocks, F. H.; Harman, C. M.; Klenk, P. A.; Simmons, W. N.

Acting as a virtual first stage, a hypervelocity launch together with the use of frozen hydrogen/frozen oxygen propellant, offers a Single-Stage-To-Orbit (SSTO) system that promises an enormous increase in SSTO mass-ratio. Ram acceleration provides hypervelocity (2 km/sec) to the orbital vehicle with a gas gun supplying the initial velocity required for ram operation. The vehicle itself acts as the center body of a ramjet inside a launch tube, filled with gaseous fuel and oxidizer, acting as an engine cowling. The high acceleration needed to achieve hypervelocity precludes a crew, and it would require greatly increased liquid fuel tank structural mass if a liquid propellant is used for post-launch vehicle propulsion. Solid propellants do not require as much fuel- chamber strengthening to withstand a hypervelocity launch as do liquid propellants, but traditional solid fuels have lower exhaust velocities than liquid hydrogen/liquid oxygen. The shock-stability of frozen hydrogen/frozen oxygen propellant has been experimentally demonstrated. A hypervelocity launch system using frozen hydrogen/frozen oxygen propellant would be a revolutionary new development in spaceflight.

Performance and Cost Evaluation of Cryogenic Solid Propulsion Systems

NASA Astrophysics Data System (ADS)

Adirim, Harry; Lo, Roger; Knecht, Thomas; Reinbold, Georg-Friedrich; Poller, Sascha

2002-01-01

Under the sponsorship of the German Aerospace Center DLR, Cryogenic Solid Propulsion (CSP) is now in its 6th year of R&D. The development proceeds as a joint international university-, small business-, space industry- and professional research effort (Berlin University of Technology / AI: Aerospace Institute, Berlin / Bauman Moscow State Technical University, Russia / ASTRIUM GmbH, Bremen / Fraunhofer Institute for Chemical Technology, Berghausen). This paper aims at introducing CSP as a novel type of chemical propellant that uses frozen liquids as Oxygen (SOX) or Hydrogen Peroxide (SH2O2) inside of a coherent solid Hydrocarbon (PE, PU or HTPB) matrix in solid rocket motors. Theoretically any conceivable chemical rocket propellant combination (including any environmentally benign ,,green propellant") can be used in solid rocket propellant motors if the definition of solids is not restricted to "solid at ambient temperature". The CSP concept includes all suitable high energy propellant combinations, but is not limited to them. Any liquid or hybrid bipropellant combination is (Isp-wise) superior to any conventional solid propellant formulation. While CSPs do share some of the disadvantages of solid propulsion (e.g. lack of cooling fluid and preset thrust-time function), they definitely share one of their most attractive advantages: the low number of components that is the base for high reliability and low cost of structures. In this respect, CSPs are superior to liquid propellant rocket motors with whom, they share the high Isp performance. High performance, low cost, low pollution CSP technology could bring about a near term improvement for chemical Earth-to-orbit high thrust propulsion. In the long run it could surpass conventional chemical propulsion because it is better suited for applying High Energy Density Matter (HEDM) than any other mode of propulsion. So far, ongoing preliminary analyses have not shown any insuperable problems in areas of concern, such as cooling equipment and its operation during fabrication and launch, neither were there problems with thrust to weight ratio of un-cooled but insulated Cryogenic Solid Motors which ascend into their trajectory while leaving the cooling equipment at the launch pad. In performance calculations for new launchers with CSP-replacements of boosters or existing stages, ARIANE 5 and a 3-stage launcher with CSP - 1st stage into GTO serve as examples. For keeping payload-capacity in the reference orbit constant, the modeling of a rocket system essentially requires a process of iteration, in which the propellant mass is varied as central parameter and - with the help of a CSP mass-model - all other dimensions of the booster are derived from mass models etc. accordingly. The process is repeated until the payload resulting from GTO track-optimization corresponds with that of the model ARIANE 5 in sufficient approximation. Under the assumptions made, the application of cryogenic motors lead to a clear reduction of the launch mass. This is essentially caused by the lower propellant mass and secondary by the reduced structure mass. Finally cost calculations have been made by ASTRIUM and demonstrated the cost saving potential of CSP propulsion. For estimating development, production, ground facilities, and operating cost, the parametric cost modeling tool has been used in combination with Cost Estimating Relationships (CER). Parametric cost models only allow comparative analyses, therefore ARIANE 5 in its current (P1) configuration has been estimated using the same mission model as for the CSP launcher. As conclusion of these cost assessment can be stated, that the utilization of cryogenic solid propulsion could offer a considerable cost savings potential. Academic and industrial cooperation is crucial for the challenging R&D work required. It will take the combined capacities of all experts involved to unlock the promises of clean, high Isp CSP propulsion for chemical Earth-to-orbit transportation in next 10 to 15 years to come.

Optical Measurements on Solid Specimens of Solid Rocket Motor Exhaust and Solid Rocket Motor Slag

NASA Technical Reports Server (NTRS)

Roberts, F. E., III

1991-01-01

Samples of aluminum slag were investigated to aid the Earth Science and Applications Division at the Marshall Space Flight Center (MSFC). Alumina from space motor propellant exhaust and space motor propellant slag was examined as a component of space refuse. Thermal emittance and solar absorptivity measurements were taken to support their comparison with reflectance measurements derived from actual debris. To determine the similarity between the samples and space motor exhaust or space motor slag, emittance and absorbance results were correlated with an examination of specimen morphology.

Space Shuttle solid rocket motor exposure monitoring

NASA Technical Reports Server (NTRS)

Brown, S. W.

1993-01-01

During the processing of the Space Shuttle Solid Rocket Booster (SRB), segments at the Kennedy Space Center, an odor was detected around the solid propellant. An Industrial Hygiene survey was conducted to determine the chemical identity of the SRB offgassing constituents. Air samples were collected inside a forward SRB segment and analyzed to determine chemical composition. Specific chemical analysis for suspected offgassing constituents of the propellant indicated ammonia to be present. A gas chromatograph mass spectroscopy (GC/MS) analysis of the air samples detected numerous high molecular weight hydrocarbons.

Detonation command and control

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

Mace, Jonathan Lee; Seitz, Gerald J.; Echave, John A.

The detonation of one or more explosive charges and propellant charges by a detonator in response to a fire control signal from a command and control system comprised of a command center and instrumentation center with a communications link therebetween. The fire control signal is selectively provided to the detonator from the instrumentation center if plural detonation control switches at the command center are in a fire authorization status, and instruments, and one or more interlocks, if included, are in a ready for firing status. The instrumentation and command centers are desirably mobile, such as being respective vehicles.

Detonation command and control

DOEpatents

Mace, Jonathan L.; Seitz, Gerald J.; Echave, John A.; Le Bas, Pierre-Yves

2015-11-10

The detonation of one or more explosive charges and propellant charges by a detonator in response to a fire control signal from a command and control system comprised of a command center and instrumentation center with a communications link therebetween. The fire control signal is selectively provided to the detonator from the instrumentation center if plural detonation control switches at the command center are in a fire authorization status, and instruments, and one or more interlocks, if included, are in a ready for firing status. The instrumentation and command centers are desirably mobile, such as being respective vehicles.

Detonation command and control

DOEpatents

Mace, Jonathan L.; Seitz, Gerald J.; Echave, John A.; Le Bas, Pierre-Yves

2016-05-31

The detonation of one or more explosive charges and propellant charges by a detonator in response to a fire control signal from a command and control system comprised of a command center and instrumentation center with a communications link there between. The fire control signal is selectively provided to the detonator from the instrumentation center if plural detonation control switches at the command center are in a fire authorization status, and instruments, and one or more interlocks, if included, are in a ready for firing status. The instrumentation and command centers are desirably mobile, such as being respective vehicles.

KSC technicians use propellant slump measurement tool on ATA SRM

NASA Technical Reports Server (NTRS)

1988-01-01

Kennedy Space Center (KSC) technicians use new propellant slump measurement tool on the Assembly Test Article (ATA) aft solid rocket motor (SRM). The tool measures any slumping of the top of the solid rocket booster (SRB) solid propellant. Data gathered by this tool and others during the ATA test will be analyzed by SRM engineers. Astronaut Stephen S. Oswald at far right (barely visible) and Morton Thiokol supervisor Howard Fichtl look on during the data gathering process. The month-long ATA test is designed to evaluate the performance of new tools required to put the tighter fitting redesigned SRM joints together. In addition, new procedures are being used and ground crews are receiving training in preparation for stacking the STS-26 flight set of motors. View provided by KSC with alternate number KSC-87PC-956.

JANNAF 36th Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

1999-01-01

Volume 1, the first of three volumes is a compilation of 47 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 36th Combustion Subcommittee held jointly with the 24th Airbreathing Propulsion Subcommittee and 18th Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Solid phase propellant combustion topics covered in this volume include cookoff phenomena in the pre- and post-ignition phases, solid rocket motor and gun propellant combustion, aluminized composite propellant combustion, combustion modeling and combustion instability and instability measurement techniques.

Pressure-coupled combustion response model for solid propellants based on Zeldovich-Novozhilov approach

NASA Technical Reports Server (NTRS)

Harstad, K. G.; Strand, L. D.

1987-01-01

An exact analytical solution is given to the problem of long-time propellant thermal response to a specified pressure oscillation. Coupling to the gas phase is made using the quasisteady Zeldovich-Novozhilov approximation. Explicit linear and lowest order (quadratic) nonlinear expressions for propellant response are obtained from the implicit nonlinear solutions. Using these expressions, response curves are presented for an ammonium perchlorate composite propellant and HMX monopropellant.

Hydrogen-oxygen catalytic ignition and thruster investigation. Volume 2: High pressure thruster evaluations

NASA Technical Reports Server (NTRS)

Johnson, R. J.; Heckert, B.; Burge, H. L.

1972-01-01

A high pressure thruster effort was conducted with the major objective of demonstrating a duct cooling concept with gaseous propellant in a thruster operating at nominally 300 psia and 1500 lbf. The analytical design methods for the duct cooling were proven in a series of tests with both ambient and reduced temperature propellants. Long duration tests as well as pulse mode tests demonstrated the feasibility of the concept. All tests were conducted with a scaling of the raised post triplet injector design previously demonstrated at 900 lbf in demonstration firings. A series of environmental conditioned firings were also conducted to determine the effects of thermal soaks, atmospheric air and high humidity. This volume presents the results of the high pressure thruster evaluations.

Quick look test report: MPT static firing no. 2 test MPT-S2

NASA Technical Reports Server (NTRS)

1978-01-01

The three engine cluster was fired at 70 percent power level for a nominal 15 seconds to evaluate the integrated performance of the main propulsion system. Engine ignition occurred at approximately 1403 with the planned mainstage duration achieved for all three engines. Operation of all systems was as expected with the exception of the recirculation pumps. The pumps were started while the propellant loading was in fast fill, but they cavitated and lost head at the termination of fast fill. The pumps were subsequently restarted after pressurizing the tank and draining back propellant to get good quality. Post test inspection of the engines revealed some discoloration on the inside of the thrust chamber and distorted drain lines for engine #2.

Model for Steady-State Combustion of Unimodal Composite Solid Propellants.

DTIC Science & Technology

1978-01-01

Research and Technology Div.do= * 5390 Cherokee Avenue Alexandria, Virginia 22314 Cw* Contract F49620-78-C-0016 Air Force Office of Scientific Research ...owmaretgli w SW MODEL FOR STEADY-STATE COMBUSTION OF UNIMODAL COMPOSITE SOLID PROPELLANTS* Dr. Merrill K. Kingk* Atlantic Research Corporation...this country today) for pre- model, all flames are considered to occur in flame sheets at discrete distances from the * Research sponsored by the Air

Calculation and design of a ramjet missile

NASA Astrophysics Data System (ADS)

Schubert, Johannes

The fundamentals for the design of a ramjet missile are treated. The chemical fundamentals of the solid rocket propellants used for ramjet missiles are outlined. The determination of the most favorable flying speed is discussed. The thermodynamic fundamentals (calculation of the solid propellant missile, calculation of the mixing procedure and the after burning in the pressure nozzle, and power calculation) are presented. The design specifications of the propulsion system are given.

Some problems of nonlinear waves in solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Culick, F. E. C.

1979-01-01

An approximate technique for analyzing nonlinear waves in solid propellant rocket motors is presented which inexpensively provides accurate results up to amplitudes of ten percent. The connection with linear stability analysis is shown. The method is extended to third order in the amplitude of wave motion in order to study nonlinear stability, or triggering. Application of the approximate method to the behavior of pulses is described.

Numerical study on the influence of aluminum on infrared radiation signature of exhaust plume

NASA Astrophysics Data System (ADS)

Zhang, Wei; Ye, Qing-qing; Li, Shi-peng; Wang, Ning-fei

2013-09-01

The infrared radiation signature of exhaust plume from solid propellant rockets has been widely mentioned for its important realistic meaning. The content of aluminum powder in the propellants is a key factor that affects the infrared radiation signature of the plume. The related studies are mostly on the conical nozzles. In this paper, the influence of aluminum on the flow field of plume, temperature distribution, and the infrared radiation characteristics were numerically studied with an object of 3D quadrate nozzle. Firstly, the gas phase flow field and gas-solid multi phase flow filed of the exhaust plume were calculated using CFD method. The result indicates that the Al203 particles have significant effect on the flow field of plume. Secondly, the radiation transfer equation was solved by using a discrete coordinate method. The spectral radiation intensity from 1000-2400 cm-1 was obtained. To study the infrared radiation characteristics of exhaust plume, an exceptional quadrate nozzle was employed and much attention was paid to the influences of Al203 particles in solid propellants. The results could dedicate the design of the divert control motor in such hypervelocity interceptors or missiles, or be of certain meaning to the improvement of ingredients of solid propellants.

Solid propellant processing factor in rocket motor design

NASA Technical Reports Server (NTRS)

1971-01-01

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

Rheology of composite solid propellants during motor casting

NASA Technical Reports Server (NTRS)

Klager, K.; Rogers, C. J.; Smith, P. L.

1978-01-01

Results of casting studies are reviewed so as to define the viscosity criteria insuring the fabrication of defect-free grains. The rheology of uncured propellants is analyzed showing that a realistic assessment of a propellant's flow properties must include measurement of viscosity as a function of shear stress and time after curing agent. Methods for measuring propellant viscosity are discussed, with particular attention given to the Haake-Rotovisko rotational viscometer. The effects of propellant compositional and processing variables on apparent viscosity are examined, as are results relating rheological behavior to grain defect formation during casting.

Solid rocket propellant waste disposal/ingredient recovery study

NASA Technical Reports Server (NTRS)

Mcintosh, M. J.

1976-01-01

A comparison of facility and operating costs of alternate methods shows open burning to be the lowest cost incineration method of waste propellant disposal. The selection, development, and implementation of an acceptable alternate is recommended. The recovery of ingredients from waste propellant has the probability of being able to pay its way, and even show a profit, when large consistent quantities of composite propellant are available. Ingredients recovered from space shuttle waste propellant would be worth over $1.5 million. Open and controlled burning are both energy wasteful.

Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

NASA Astrophysics Data System (ADS)

Kinefuchi, K.; Funaki, I.; Shimada, T.; Abe, T.

2012-10-01

Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

Monte Carlo investigation of thrust imbalance of solid rocket motor pairs

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.

1976-01-01

The Monte Carlo method of statistical analysis is used to investigate the theoretical thrust imbalance of pairs of solid rocket motors (SRMs) firing in parallel. Sets of the significant variables are selected using a random sampling technique and the imbalance calculated for a large number of motor pairs using a simplified, but comprehensive, model of the internal ballistics. The treatment of burning surface geometry allows for the variations in the ovality and alignment of the motor case and mandrel as well as those arising from differences in the basic size dimensions and propellant properties. The analysis is used to predict the thrust-time characteristics of 130 randomly selected pairs of Titan IIIC SRMs. A statistical comparison of the results with test data for 20 pairs shows the theory underpredicts the standard deviation in maximum thrust imbalance by 20% with variability in burning times matched within 2%. The range in thrust imbalance of Space Shuttle type SRM pairs is also estimated using applicable tolerances and variabilities and a correction factor based on the Titan IIIC analysis.

Identity and distribution of residues of energetic compounds at army live-fire training ranges.

PubMed

Jenkins, Thomas F; Hewitt, Alan D; Grant, Clarence L; Thiboutot, Sonia; Ampleman, Guy; Walsh, Marianne E; Ranney, Thomas A; Ramsey, Charles A; Palazzo, Antonio J; Pennington, Judith C

2006-05-01

Environmental investigations have been conducted at 23 military firing ranges in the United States and Canada. The specific training facilities most frequently evaluated were hand grenade, antitank rocket, and artillery ranges. Energetic compounds (explosives and propellants) were determined and linked to the type of munition used and the major mechanisms of deposition.

Hypervelocity cutting machine and method

DOEpatents

Powell, James R.; Reich, Morris

1996-11-12

A method and machine 14 are provided for cutting a workpiece 12 such as concrete. A gun barrel 16 is provided for repetitively loading projectiles 22 therein and is supplied with a pressurized propellant from a storage tank 28. A thermal storage tank 32,32A is disposed between the propellant storage tank 28 and the gun barrel 16 for repetitively receiving and heating propellant charges which are released in the gun barrel 16 for repetitively firing projectiles 22 therefrom toward the workpiece 12. In a preferred embodiment, hypervelocity of the projectiles 22 is obtained for cutting the concrete workpiece 12 by fracturing thereof.

Some experiments related to L-star instability in rocket motors

NASA Technical Reports Server (NTRS)

Kumar, R. N.; Mcnamara, R. P.

1973-01-01

The influence of condensed phase heterogeneity on the L-star instability of nonmetallized AP/PBAN propellants is explored using four propellants (with monomodal AP particle distributions having 50 per cent weight average points at 11, 39.5, 175, and 350 microns). An economical firing program is used. One-dimensional nature of the Helmholtz mode and the complex nature of the chuff mode are revealed through color movies. The stability boundary on the L-star pressure plot is found to be parabolic. Frequency correlations and many other features reveal the important role of condensed phase details in propellant combustion.

Centaur space vehicle pressurized propellant feed system tests

NASA Technical Reports Server (NTRS)

1972-01-01

Engine firing tests, using a full-scale flight-weight vehicle, were performed to evaluate a pressurized propellant feed system for the Centaur. The pressurant gases used were helium and hydrogen. The system was designed to replace the boost pumps currently used on Centaur. Two liquid oxygen tank pressurization modes were studied: (1) directly into the ullage and (2) below the propellant surface. Test results showed the two Centaur RL10 engines could be started and run over the range of expected flight variables. No system instabilities were encountered. Measured pressurization gas quantities agreed well with analytically predicted values.

Characterization of Firing Range Soil from Camp Edwards, MA, and the Efficacy of Acid and Alkaline Hydrolysis for the Remediation of M1 105mm M67 Propellant

DTIC Science & Technology

2013-06-01

method is intended for trace analysis of explosives and propellant residues by high performance liquid chromatography (HPLC) using an ultraviolet (UV...detector set at 254 nm. The HPLC used for this analysis was a Dionex Summit System with a UV detector equipped with Dionex E1 and E2 columns...Ca(OH)2) and sodium hydroxide (NaOH) were evaluated as sources of hydroxide ion for the alkaline hydrolysis of M1 propellant in soil from Camp

Aerospace Laser Ignition/Ablation Variable High Precision Thruster

NASA Technical Reports Server (NTRS)

Campbell, Jonathan W. (Inventor); Edwards, David L. (Inventor); Campbell, Jason J. (Inventor)

2015-01-01

A laser ignition/ablation propulsion system that captures the advantages of both liquid and solid propulsion. A reel system is used to move a propellant tape containing a plurality of propellant material targets through an ignition chamber. When a propellant target is in the ignition chamber, a laser beam from a laser positioned above the ignition chamber strikes the propellant target, igniting the propellant material and resulting in a thrust impulse. The propellant tape is advanced, carrying another propellant target into the ignition chamber. The propellant tape and ignition chamber are designed to ensure that each ignition event is isolated from the remaining propellant targets. Thrust and specific impulse may by precisely controlled by varying the synchronized propellant tape/laser speed. The laser ignition/ablation propulsion system may be scaled for use in small and large applications.

Materials characterization of propellants using ultrasonics

NASA Technical Reports Server (NTRS)

Workman, Gary L.; Jones, David

1993-01-01

Propellant characteristics for solid rocket motors were not completely determined for its use as a processing variable in today's production facilities. A major effort to determine propellant characteristics obtainable through ultrasonic measurement techniques was performed in this task. The information obtained was then used to determine the uniformity of manufacturing methods and/or the ability to determine non-uniformity in processes.

The Aerodynamic Characteristics of Full-Scale Propellers Having 2, 3, and 4 Blades of Clark Y and R.A.F. 6 Airfoil Sections

NASA Technical Reports Server (NTRS)

Hartman, Edwin P; Biermann, David

1938-01-01

Aerodynamic tests were made of seven full-scale 10-foot-diameter propellers of recent design comprising three groups. The first group was composed of three propellers having Clark y airfoil sections and the second group was composed of three propellers having R.A.F. 6 airfoil sections, the propellers of each group having 2, 3, and 4 blades. The third group was composed of two propellers, the 2-blade propeller taken from the second group and another propeller having the same airfoil section and number of blades but with the width and thickness 50 percent greater. The tests of these propellers reveal the effect of changes in solidity resulting either from increasing the number of blades or from increasing the blade width propeller design charts and methods of computing propeller thrust are included.

Full pillar extraction at the Kathleen Mine with mobile roof supports

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

Grimm, E.S.

1994-12-31

The Voest Alpine Breaker Line Supports (ABLS) resemble self-propelled longwall shields. Each individual unit consists of four hydraulic legs extending from the base of the unit, pressing a solid flat canopy against the mine roof. Each support unit is capable of exerting 606 tons of force against the roof. A chain curtain on the sides and rear protects the interior of the support from falling rock. The internal scissoring lemniscate design allows for parallel movement of the canopy as it is raised or lowered. Each ABLS has 750 feet of 4 AWG trailing cable to supply 480 volts AC tomore » a permissible controller and a 40 hp explosion-proof electrical motor. The hydraulic pump and reservoir are self-contained and protected with an automatic fire suppression system.« less

Near noise field characteristics of Nike rocket motors for application to space vehicle payload acoustic qualification

NASA Technical Reports Server (NTRS)

Hilton, D. A.; Bruton, D.

1977-01-01

Results of a series of noise measurements that were made under controlled conditions during the static firing of two Nike solid propellant rocket motors are presented. The usefulness of these motors as sources for general spacecraft noise testing was assessed, and the noise expected in the cargo bay of the orbiter was reproduced. Brief descriptions of the Nike motor, the general procedures utilized for the noise tests, and representative noise data including overall sound pressure levels, one third octave band spectra, and octave band spectra were reviewed. Data are presented on two motors of different ages in order to show the similarity between noise measurements made on motors having different loading dates. The measured noise from these tests is then compared to that estimated for the space shuttle orbiter cargo bay.

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.

Army and Marine Corps Active Protection System (APS) Efforts

DTIC Science & Technology

2016-08-23

with hard or soft kill capabilities to a variety of threats, including rocket -propelled grenades (RPGs) and anti-tank guided missiles (ATGMs). APS...of threats, including rocket -propelled grenades (RPGs) and anti-tank guided missiles (ATGMs). APS technologies are not new, and a number of nations...training. 1 RPGs are basically single man-portable, shoulder-fired, unguided rockets . RPGs have been widely proliferated but can be mitigated to a

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.

Evaluation of weapons' combustion products in armored vehicles. Appendix C: Summary of descriptive statistics (proc tabulate). Appendix D: Summary of comparative statistics. Final report, 30 September 1986-14 December 1988

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

Menzies, K.T.; Randel, M.A.; Quill, A.L.

1989-01-01

The U.S. Army Biomedical Research and Development Laboratory defined an extensive research program to address the generation of potentially toxic propellant combustion products in crew compartments of armored vehicles during weapons firing. The major objectives of the research were (1) to determine the presence and concentration of propellant combustion products, (2) to determine potential crew exposure to these combustion products, and (3) to assess the efficacy of field monitoring in armored vehicles. To achieve these goals, air monitoring was conducted in selected armored vehicle types, i.e., M109, M60, M3, M1, at several Army installations. Auxiliary information concerning the specific munitionsmore » fired and the Training and Doctrine Command (TRADOC) or Forces Command (FORSCOM) firing scenarios was collected so that a comparison of pollutant concentrations generated by specific weapons both within vehicle types and between vehicle types could be made.« less

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.

Enhanced alkaline hydrolysis and biodegradability studies of nitrocellulose-bearing missile propellant

NASA Technical Reports Server (NTRS)

Sidhoum, Mohammed; Christodoulatos, Christos; Su, Tsan-Liang; Redis, Mercurios

1995-01-01

Large amounts of energetic materials which have been accumulated over the years in various manufacturing and military installations must be disposed of in an environmentally sound manner. Historically, the method of choice for destruction of obsolete or aging energetic materials has been open burning or open detonation (OB/OD). This destruction approach has become undesirable due to air pollution problems. Therefore, there is a need for new technologies which will effectively and economically deal with the disposal of energetic materials. Along those lines, we have investigated a chemical/biological process for the safe destruction and disposal of a double base solid rocket propellant (AHH), which was used in several 8 inch projectile systems. The solid propellant is made of nitrocellulose and nitroglycerin as energetic components, two lead salts which act as ballistic modifiers, triacetin as a plasticizer and 2-Nitrodiphenylamine (2-NDPA) as a stabilizer. A process train is being developed to convert the organic components of the propellant to biodegradable products and remove the lead from the process stream. The solid propellant is first hydrolyzed through an enhanced alkaline hydrolysis process step. Following lead removal and neutralization, the digested liquor rich in nitrates and nitrites is found to be easily biodegradable. The digestion rate of the intact ground propellant as well as the release of nitrite and nitrate groups were substantially increased when ultrasound were supplied to the alkaline reaction medium compared to the conventional alkaline hydrolysis. The effects of reaction time, temperature, sodium hydroxide concentration and other relevant parameters on the digestion efficiency and biodegradability have been studied. The present work indicates that the AHH propellant can be disposed of safely with a combination of physiochemical and biological processes.

Study of aluminum particle combustion in solid propellant plumes using digital in-line holography and imaging pyrometry

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

Chen, Yi; Guildenbecher, Daniel R.; Hoffmeister, Kathryn N. G.

The combustion of molten metals is an important area of study with applications ranging from solid aluminized rocket propellants to fireworks displays. Our work uses digital in-line holography (DIH) to experimentally quantify the three-dimensional position, size, and velocity of aluminum particles during combustion of ammonium perchlorate (AP) based solid-rocket propellants. Additionally, spatially resolved particle temperatures are simultaneously measured using two-color imaging pyrometry. To allow for fast characterization of the properties of tens of thousands of particles, automated data processing routines are proposed. In using these methods, statistics from aluminum particles with diameters ranging from 15 to 900 µm are collectedmore » at an ambient pressure of 83 kPa. In the first set of DIH experiments, increasing initial propellant temperature is shown to enhance the agglomeration of nascent aluminum at the burning surface, resulting in ejection of large molten aluminum particles into the exhaust plume. The resulting particle number and volume distributions are quantified. In the second set of simultaneous DIH and pyrometry experiments, particle size and velocity relationships as well as temperature statistics are explored. The average measured temperatures are found to be 2640 ± 282 K, which compares well with previous estimates of the range of particle and gas-phase temperatures. The novel methods proposed here represent new capabilities for simultaneous quantification of the joint size, velocity, and temperature statistics during the combustion of molten metal particles. The proposed techniques are expected to be useful for detailed performance assessment of metalized solid-rocket propellants.« less

SOLID SOLUTION CARBIDES ARE THE KEY FUELS FOR FUTURE NUCLEAR THERMAL PROPULSION

NASA Technical Reports Server (NTRS)

Panda, Binayak; Hickman, Robert R.; Shah, Sandeep

2005-01-01

Nuclear thermal propulsion uses nuclear energy to directly heat a propellant (such as liquid hydrogen) to generate thrust for space transportation. In the 1960 s, the early Rover/Nuclear Engine for Rocket Propulsion Application (NERVA) program showed very encouraging test results for space nuclear propulsion but, in recent years, fuel research has been dismal. With NASA s renewed interest in long-term space exploration, fuel researchers are now revisiting the RoverMERVA findings, which indicated several problems with such fuels (such as erosion, chemical reaction of the fuel with propellant, fuel cracking, and cladding issues) that must be addressed. It is also well known that the higher the temperature reached by a propellant, the larger the thrust generated from the same weight of propellant. Better use of fuel and propellant requires development of fuels capable of reaching very high temperatures. Carbides have the highest melting points of any known material. Efforts are underway to develop carbide mixtures and solid solutions that contain uranium carbide, in order to achieve very high fuel temperatures. Binary solid solution carbides (U, Zr)C have proven to be very effective in this regard. Ternary carbides such as (U, Zr, X) carbides (where X represents Nb, Ta, W, and Hf) also hold great promise as fuel material, since the carbide mixtures in solid solution generate a very hard and tough compact material. This paper highlights past experience with early fuel materials and bi-carbides, technical problems associated with consolidation of the ingredients, and current techniques being developed to consolidate ternary carbides as fuel materials.

Study of aluminum particle combustion in solid propellant plumes using digital in-line holography and imaging pyrometry

DOE PAGES

Chen, Yi; Guildenbecher, Daniel R.; Hoffmeister, Kathryn N. G.; ...

2017-05-05

The combustion of molten metals is an important area of study with applications ranging from solid aluminized rocket propellants to fireworks displays. Our work uses digital in-line holography (DIH) to experimentally quantify the three-dimensional position, size, and velocity of aluminum particles during combustion of ammonium perchlorate (AP) based solid-rocket propellants. Additionally, spatially resolved particle temperatures are simultaneously measured using two-color imaging pyrometry. To allow for fast characterization of the properties of tens of thousands of particles, automated data processing routines are proposed. In using these methods, statistics from aluminum particles with diameters ranging from 15 to 900 µm are collectedmore » at an ambient pressure of 83 kPa. In the first set of DIH experiments, increasing initial propellant temperature is shown to enhance the agglomeration of nascent aluminum at the burning surface, resulting in ejection of large molten aluminum particles into the exhaust plume. The resulting particle number and volume distributions are quantified. In the second set of simultaneous DIH and pyrometry experiments, particle size and velocity relationships as well as temperature statistics are explored. The average measured temperatures are found to be 2640 ± 282 K, which compares well with previous estimates of the range of particle and gas-phase temperatures. The novel methods proposed here represent new capabilities for simultaneous quantification of the joint size, velocity, and temperature statistics during the combustion of molten metal particles. The proposed techniques are expected to be useful for detailed performance assessment of metalized solid-rocket propellants.« less

Improved Net-Level Filling And Finishing Of Large Castings

NASA Technical Reports Server (NTRS)

Johnson, Erik P.; Brown, Richard F.

1995-01-01

Improved method of vacuum casting of large, generally cylindrical objects to net sizes and shapes reduces amount of direct manual labor by workers in proximity to cast material. Original application for which method devised is fabrication of solid rocket-motor segments containing solid propellant, wherein need to minimize exposure of workers to propellant material being cast. Improved method adaptable to other applications involving large castings of toxic, flammable, or otherwise hazardous materials.

Buckling of thin walled composite cylindrical shell filled with solid propellant

NASA Astrophysics Data System (ADS)

Dash, A. P.; Velmurugan, R.; Prasad, M. S. R.

2017-12-01

This paper investigates the buckling of thin walled composite cylindrical tubes that are partially filled with solid propellant equivalent elastic filler. Experimental investigation is conducted on thin composite tubes made out of S2-glass epoxy, which is made by using filament winding technique. The composite tubes are filled with elastic filler having similar mechanical properties as that of a typical solid propellant used in rocket motors. The tubes are tested for their buckling strength against the external pressure in the presence of the filler. Experimental data confirms the enhancement of external pressure carrying capacity of the composite tubes by up to three times as that of empty tubes for a volumetric loading fraction (VLF) of 0.9. Furthermore, the finite element based geometric nonlinearity analysis predicts the buckling behaviour of the partially filled composite tubes close to the experimental results.

Prediction of crosslink density of solid propellant binders. [curing of elastomers

NASA Technical Reports Server (NTRS)

Marsh, H. E., Jr.

1976-01-01

A quantitative theory is outlined which allows calculation of crosslink density of solid propellant binders from a small number of predetermined parameters such as the binder composition, the functionality distributions of the ingredients, and the extent of the curing reaction. The parameter which is partly dependent on process conditions is the extent of reaction. The proposed theoretical model is verified by independent measurement of effective chain concentration and sol and gel fractions in simple compositions prepared from model compounds. The model is shown to correlate tensile data with composition in the case of urethane-cured polyether and certain solid propellants. A formula for the branching coefficient is provided according to which if one knows the functionality distributions of the ingredients and the corresponding equivalent weights and can measure or predict the extent of reaction, he can calculate the branching coefficient of such a system for any desired composition.

Solid Hydrogen Experiments for Atomic Propellants

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2001-01-01

This paper illustrates experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Solid particles of hydrogen were frozen in liquid helium, and observed with a video camera. The solid hydrogen particle sizes, their molecular structure transitions, and their agglomeration times were estimated. article sizes of 1.8 to 4.6 mm (0.07 to 0. 18 in.) were measured. The particle agglomeration times were 0.5 to 11 min, depending on the loading of particles in the dewar. These experiments are the first step toward visually characterizing these particles, and allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Space Shuttle with rail system and aft thrust structure securing solid rocket boosters to external tank

NASA Technical Reports Server (NTRS)

Vonpragenau, G. L. (Inventor)

1984-01-01

The configuration and relationship of the external propellant tank and solid rocket boosters of space transportation systems such as the space shuttle are described. The space shuttle system with the improved propellant tank is shown. The external tank has a forward pressure vessel for liquid hydrogen and an aft pressure vessel for liquid oxygen. The solid rocket boosters are joined together by a thrust frame which extends across and behind the external tank. The thrust of the orbiter's main rocket engines are transmitted to the aft portion of the external tank and the thrust of the solid rocket boosters are transmitted to the aft end of the external tank.

A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

Backlund, S. J.; Rossen, J. N.

1971-01-01

A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

Combustion characteristics of SMX and SMX based propellants

NASA Astrophysics Data System (ADS)

Reese, David A.

This work investigates the combustion of the new solid nitrate ester 2,3-hydroxymethyl-2,3-dinitro-1,4-butanediol tetranitrate (SMX, C6H 8N6O16). SMX was synthesized for the first time in 2008. It has a melting point of 85 °C and oxygen balance of 0% to CO 2, allowing it to be used as an energetic additive or oxidizer in solid propellants. In addition to its neat combustion characteristics, this work also explores the use of SMX as a potential replacement for nitroglycerin (NG) in double base gun propellants and as a replacement for ammonium perchlorate in composite rocket propellants. The physical properties, sensitivity characteristics, and combustion behaviors of neat SMX were investigated. Its combustion is stable at pressures of up to at least 27.5 MPa (n = 0.81). The observed flame structure is nearly identical to that of other double base propellant ingredients, with a primary flame attached at the surface, a thick isothermal dark zone, and a luminous secondary flame wherein final recombination reactions occur. As a result, the burning rate and primary flame structure can be modeled using existing one-dimensional steady state techniques. A zero gas-phase activation energy approximation results in a good fit between modeled and observed behavior. Additionally, SMX was considered as a replacement for nitroglycerin in a double base propellant. Thermochemical calculations indicate improved performance when compared with the common double base propellant JA2 at SMX loadings above 40 wt-%. Also, since SMX is a room temperature solid, migration may be avoided. Like other nitrate esters, SMX is susceptible to decomposition over long-term storage due to the presence of excess acid in the crystals; the addition of stabilizers (e.g., derivatives of urea) during synthesis should be sufficient to prevent this. the addition of Both unplasticized and plasticized propellants were formulated. Thermal analysis of unplasticized propellant showed a distinct melt-recrystallization curve, which indicates that a solid phase solution is being formed between SMX and NC, and that SMX would not act as plasticizer. Analysis of propellant prepared with diethyleneglycol dinitrate (DEGDN) plasticizer indicates that the SMX is likely dissolved in the DEGDN. The plasticized material also showed similar hardness and modulus to JA2. However, both plasticized and unplasticized propellants exhibited deconsolidated burning at elevated pressures due to the high modulus of the propellant. Increased amounts of plasticizer or improved processing of the nitrocellulose should be investigated to remedy this issue. Safety characterization showed that sensitivity of the plasticized propellant is similar to JA2. In short, replacing NG with SMX results in a new family of propellants with acceptable safety characteristics and which may also offer improved theoretical performance. Finally, composite propellants based on SMX were theoretically and experimentally examined and compared to formulations based on ammonium perchlorate (AP). Thermochemical equilibrium calculations show that aluminized SMX-based formulations can achieve theoretical sea level specific impulse values upwards of 260 s-- slightly lower than an AP-based composite. Both ignition sensitivity (tested via drop weight impact, electro-static discharge, and BAM friction) and physical properties (hardness and thermal properties) are comparable to those of the AP-based formulations. However, the SMX-based formulation could be detonated using a high explosive donor charge in contact with the propellant, as do other low smoke propellants. Differential scanning calorimetry of the SMX-based propellant indicated an exotherm onset of 140 °C, which corresponds to the known decomposition temperature of SMX. The propellant has a high burning rate of 1.57 cm/s at 6.89 MPa, with a pressure exponent of 0.85. This high pressure sensitivity might be addressed using various energetic and/or stabilizing additives. With high density and performance, smokeless combustion products, and stable combustion, SMX appears to be a viable replacement for existing energetic ingredients in a wide variety of propellant, explosive, and pyrotechnic applications.

40 CFR 61.44 - Stack sampling.

Code of Federal Regulations, 2013 CFR

2013-07-01

... EMISSION STANDARDS FOR HAZARDOUS AIR POLLUTANTS National Emission Standard for Beryllium Rocket Motor... within 30 days after samples are taken and before any subsequent rocket motor firing or propellant...

40 CFR 61.44 - Stack sampling.

Code of Federal Regulations, 2014 CFR

2014-07-01

... EMISSION STANDARDS FOR HAZARDOUS AIR POLLUTANTS National Emission Standard for Beryllium Rocket Motor... within 30 days after samples are taken and before any subsequent rocket motor firing or propellant...

40 CFR 61.44 - Stack sampling.

Code of Federal Regulations, 2012 CFR

2012-07-01

... EMISSION STANDARDS FOR HAZARDOUS AIR POLLUTANTS National Emission Standard for Beryllium Rocket Motor... within 30 days after samples are taken and before any subsequent rocket motor firing or propellant...

Effects of magnesium-based hydrogen storage materials on the thermal decomposition, burning rate, and explosive heat of ammonium perchlorate-based composite solid propellant.

PubMed

Liu, Leili; Li, Jie; Zhang, Lingyao; Tian, Siyu

2018-01-15

MgH 2 , Mg 2 NiH 4 , and Mg 2 CuH 3 were prepared, and their structure and hydrogen storage properties were determined through X-ray photoelectron spectroscopy and thermal analyzer. The effects of MgH 2 , Mg 2 NiH 4 , and Mg 2 CuH 3 on the thermal decomposition, burning rate, and explosive heat of ammonium perchlorate-based composite solid propellant were subsequently studied. Results indicated that MgH 2 , Mg 2 NiH 4 , and Mg 2 CuH 3 can decrease the thermal decomposition peak temperature and increase the total released heat of decomposition. These compounds can improve the effect of thermal decomposition of the propellant. The burning rates of the propellant increased using Mg-based hydrogen storage materials as promoter. The burning rates of the propellant also increased using MgH 2 instead of Al in the propellant, but its explosive heat was not enlarged. Nonetheless, the combustion heat of MgH 2 was higher than that of Al. A possible mechanism was thus proposed. Copyright © 2017. Published by Elsevier B.V.

Task 2: Flight prototype system design report, pulsed plasma solid propellant microthruster for the Synchronous Meteorological Satellite

NASA Technical Reports Server (NTRS)

Guman, W. J. (Editor)

1972-01-01

Design details are presented of the solid propellant pulsed plasma microthruster which was analyzed during the Task 1 effort. The design details presented show that the inherent functional simplicity underlying the flight proven LES-6 design can be maintained in the SMS systems design even with minimum weight constraints imposed. A 1293 hour uninterrupted vacuum test with the engineering thermal model, simulating an 18.8 to 33 g environment of the propellant, its feed system and electrode assembly, revealed that program thruster performance requirements could be met. This latter g environment is a more severe environment than will be ever encountered in the SMS spacecraft.

Aircraft Survivability: Vulnerability Reduction, Spring 2006

DTIC Science & Technology

2006-01-01

selected small arms, rocket propelled grenades, and shoulder-fired missiles will be presented. Figure 1 and Figure 2 illustrate previous demonstrations...lethality. Hands-on experience will be provided with threat munitions and missiles , test articles, damaged-air- craft hardware, live fire...non-linear effects of scale and operational environment. Current Efforts In the structures S&T program at the US Army Aviation and Missile Research

Environmental Effects of Space Shuttle Solid Rocket Motor Exhaust Plumes

NASA Technical Reports Server (NTRS)

Hwang, B.; Pergament, H. S.

1976-01-01

The deposition of NOx and HCl in the stratosphere from the space shuttle solid rocket motors (SRM) and exhaust plume is discussed. A detailed comparison between stratospheric deposition rates using the baseline SRM propellant and an alternate propellant, which replaces ammonium perchlorate by ammonium nitrate, shows the total NOx deposition rate to be approximately the same for each propellant. For both propellants the ratio of the deposition rates of NOx to total chlorine-containing species is negligibly small. Rocket exhaust ground cloud transport processes in the troposphere are also examined. A brief critique of the multilayer diffusion models (presently used for predicting pollutant deposition in the troposphere) is presented, and some detailed cloud rise calculations are compared with data for Titan 3C launches. The results show that, when launch time meteorological data are used as input, the model can reasonably predict measured cloud stabilization heights.

2005 40th Annual Armament Systems: Guns - Ammunition - Rockets - Missiles Conference and Exhibition. Volume 3: Wednesday

DTIC Science & Technology

2005-04-28

Lessons Learned, Mr. David F. Fair, US Army ARDEC Propellant Replacement for the 105-mm M67 Propelling Charge, Ms. Adriana L. Eng, US Army ARDEC Lead...Application of Lessons Learned Mr. David F. Fair, US Army ARDEC Propellant Replacement for the 105-mm Artillery Propelling Charge Ms. Adriana L. Eng...high voltage power supply (several kV and kA ) • Solid state Switching device • Appropriate dimensions en properties of: • Exploding foil • Flyer

The pasty propellant rocket engine development

NASA Astrophysics Data System (ADS)

Kukushkin, V. I.; Ivanchenko, A. N.

1993-06-01

The paper describes a newly developed pasty propellant rocket engine (PPRE) and the combustion process and presents results of performance tests. It is shown that, compared with liquid propellant rocket engines, the PPREs can regulate the thrust level within a wider range, are safer ecologically, and have better weight characteristics. Compared with solid propellant rocket engines, the PPREs may be produced with lower costs and more safely, are able to regulate thrust performance within a wider range, and are able to offer a greater scope for the variation of the formulation components and propellant characteristics. Diagrams of the PPRE are included.

Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy.

PubMed

Terry, Brandon C; Sippel, Travis R; Pfeil, Mark A; Gunduz, I Emre; Son, Steven F

2016-11-05

Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (ISP). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal ISP by ∼7s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5±4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption. Copyright © 2016 Elsevier B.V. All rights reserved.

Space Propulsion Hazards Analysis Manual (SPHAM). Volume 1

DTIC Science & Technology

1988-10-01

Wiley, New York, 1983, p.p. 64-68 (11) Martin Marietta MCR 82-800, Rev. B, 29 September 1982, "DOD Safety Review Team Lessons Learned Data Base...FLinaIRe-p,.-t, Martin Marietta Technical Report , Contract F42600-81-D-1379, September 1982. (57) Bader, Donaldson, et. al., Liquid Propellant Rocket Abort...Fire Model, Journal of Astronautics and Aeronautics, December 1971. (58) Banning, D., Propellant_$pill Analysi, Martin Marietta Technical Report , July

Ignition propagation and heat effects of propellant chips embedded in castable inhibitor using a laser flux test bomb

NASA Technical Reports Server (NTRS)

Bolton, Douglas E., Jr.

1993-01-01

A castable inhibitor is applied to the aft face of the Space Shuttle Redesigned Solid Rocket Motor (RSRM) forward segment propellant grain to control propellant surface burn area. During fabrication, the propellant surface is trimmed prior to the inhibitor application. This produces a potential for small propellant chips to remain undetected on the propellant surface and contaminate the inhibitor during application. The concern was that undetected propellant chips in the inhibitor might provide a fuse path for premature propellant ignition underneath the inhibitor. To evaluate the fuse path potential, testing was performed on inhibitor samples with embedded propellant. The internal motor environment was simulated with a calibrated CO2 laser beam directed onto a sample which was placed in a 4100 kPa (600 psi) nitrogen pressurized bomb (laser bomb). The testing showed definitive results pertaining to fuse path formation. Embedded propellant chips did not autoignite until the receding heat affected inhibitor surface reached, or passed, the propellant chip. Samples with embedded propellant chips in alignment did not propagate ignition from one chip to another with separation distances as small as 0.010 cm(0.004 inc) and some as little as 0.0051 cm (0.002 in). Propellant chips with volumes approximately less than 0.025 cu cm (0.0015 cu in) (which did not propagate ignition) did not increase the inhibitor material decomposition depth more than the resulting void cavity of the burned out propellant chip. In addition, the depth of this void cavity did not increase until it was overtaken by the surrounding material decomposition depth. This was due, in part, to the retention of the protective inhibitor char layer. Samples with embedded propellant strings, whose thicknesses were below 0.023 cm (0.009 in), did not propagate ignition. Propellant string thicknesses above 0.038 cm (0.015 in) did propagate ignition. Test sample char and heat affected layer measurements and observations compared well with those from the Space Shuttle Solid Rocket Motor (SRM) Technical Evaluation Motor no. 9(TEM-9).

KENNEDY SPACE CENTER, FLA. - At the Rotation, Processing and Surge Facility stand a mockup of two segments of a solid rocket booster (SRB) being used to test the feasibility of a vertical SRB propellant grain inspection, required as part of safety analysis.

NASA Image and Video Library

2003-09-11

KENNEDY SPACE CENTER, FLA. - At the Rotation, Processing and Surge Facility stand a mockup of two segments of a solid rocket booster (SRB) being used to test the feasibility of a vertical SRB propellant grain inspection, required as part of safety analysis.

Credit PSR. This view shows southeast and southwest facades as ...

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

Credit PSR. This view shows southeast and southwest facades as seen when looking east northeast (70°). This steel frame building is clad in "Transite" board (fire- resistant, pressed asbestos composition board). This structure was built as a back-up to Building 4237/E-38, but no equipment was ever installed. It was equipped instead to conduct tensile tests on propellant samples. In 1984, it was converted into a back-up structure supporting Building 4283/E-84, Propellant Processing Building. Small amounts of HMX propellants were processed and dried here - Jet Propulsion Laboratory Edwards Facility, Oxidizer Dryer Blender Building, Edwards Air Force Base, Boron, Kern County, CA

A two-phase restricted equilibrium model for combustion of metalized solid propellants

NASA Technical Reports Server (NTRS)

Sabnis, J. S.; Dejong, F. J.; Gibeling, H. J.

1992-01-01

An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

Acoustic Measurements for Small Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Vargas, Magda B.; Kenny, R. Jeremy

2010-01-01

Models have been developed to predict large solid rocket motor acoustic loads based on the scaling of small solid rocket motors. MSFC has measured several small solid rocket motors in horizontal and launch configurations to anchor these models. Solid Rocket Test Motor (SRTM) has ballistics similar to the Reusable Solid Rocket Motor (RSRM) therefore a good choice for acoustic scaling. Acoustic measurements were collected during the test firing of the Insulation Configuration Extended Length (ICXL) 7,6, and 8 (in firing order) in order to compare to RSRM horizontal firing data. The scope of this presentation includes: Acoustic test procedures and instrumentation implemented during the three SRTM firings and Data analysis method and general trends observed in the data.

European Scientific Notes. Volume 36, Number 9

DTIC Science & Technology

1982-09-30

studies of super- One of the chief reasons the’ foregoing "conducting tunneling, ultrasonic attenuation , activity was initiated was the historical...paper entitled "The Effect of HTPB propellant and binder. Results from tests Polymer Characteristics on Propellant using 105-mm munitions show that...polybutadiene ( HTPB ) composite solid Dr. A. Iwama (Institute of Space and propellants. The influence of the polymer Astronautical Science, Tokyo, Japan

AFAL (Air Force Astronautics Laboratory) Technical Objective Document FY89.

DTIC Science & Technology

1987-12-01

propellant manufacture that arc s Wu’ ’ " - applications. TITLE: Ultrasonic Nondestructive Evaluation fur Solid ’nu -! , c.. Performance Period: Jul 88...Develop and demonstrate a stabilizer system for (GAP propellants that provides protection equal to that achieved in HTPB propellants. Formulate GAP...monitoring equipment will be used to determine propagation and attenuation effects. 95 TITLE: Advanced Nuclear Propulsion Performance Period: Jan 89

Regulatory fire test requirements for plutonium air transport packages : JP-4 or JP-5 vs. JP-8 aviation fuel.

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

Figueroa, Victor G.; Lopez, Carlos; Nicolette, Vernon F.

2010-10-01

For certification, packages used for the transportation of plutonium by air must survive the hypothetical thermal environment specified in 10CFR71.74(a)(5). This regulation specifies that 'the package must be exposed to luminous flames from a pool fire of JP-4 or JP-5 aviation fuel for a period of at least 60 minutes.' This regulation was developed when jet propellant (JP) 4 and 5 were the standard jet fuels. However, JP-4 and JP-5 currently are of limited availability in the United States of America. JP-4 is very hard to obtain as it is not used much anymore. JP-5 may be easier to getmore » than JP-4, but only through a military supplier. The purpose of this paper is to illustrate that readily-available JP-8 fuel is a possible substitute for the aforementioned certification test. Comparisons between the properties of the three fuels are given. Results from computer simulations that compared large JP-4 to JP-8 pool fires using Sandia's VULCAN fire model are shown and discussed. Additionally, the Container Analysis Fire (CAFE) code was used to compare the thermal response of a large calorimeter exposed to engulfing fires fueled by these three jet propellants. The paper then recommends JP-8 as an alternate fuel that complies with the thermal environment implied in 10CFR71.74.« less

Solid Propellant Test Article (SPTA) Test Stand

NASA Technical Reports Server (NTRS)

1991-01-01

This photograph shows the Solid Propellant Test Article (SPTA) test stand with the Modified Nasa Motor (M-NASA) test article at the Marshall Space Flight Center (MSFC). The SPTA test stand, 12-feet wide by 12-feet long by 24-feet high, was built in 1989 to provide comparative performance data on nozzle and case insulation material and to verify thermostructural analysis models. A modified NASA 48-inch solid motor (M-NASA motor) with a 12-foot blast tube and 10-inch throat makes up the SPTA. The M-NASA motor is being used to evaluate solid rocket motor internal non-asbestos insulation materials, nozzle designs, materials, and new inspection techniques. New internal motor case instrumentation techniques are also being evaluated.

46 CFR 28.160 - Portable fire extinguishers.

Code of Federal Regulations, 2012 CFR

2012-10-01

... Feet (19.8 Meters) or More in Length Space Classification Quantity and location Safety areas... spaces; Internal combustion propelling machinery B-II 1 for each 1,000 brake horsepower or fraction...

46 CFR 28.160 - Portable fire extinguishers.

Code of Federal Regulations, 2014 CFR

2014-10-01

... Feet (19.8 Meters) or More in Length Space Classification Quantity and location Safety areas... spaces; Internal combustion propelling machinery B-II 1 for each 1,000 brake horsepower or fraction...

46 CFR 28.160 - Portable fire extinguishers.

Code of Federal Regulations, 2011 CFR

2011-10-01

... Feet (19.8 Meters) or More in Length Space Classification Quantity and location Safety areas... spaces; Internal combustion propelling machinery B-II 1 for each 1,000 brake horsepower or fraction...

46 CFR 28.160 - Portable fire extinguishers.

Code of Federal Regulations, 2013 CFR

2013-10-01

... Feet (19.8 Meters) or More in Length Space Classification Quantity and location Safety areas... spaces; Internal combustion propelling machinery B-II 1 for each 1,000 brake horsepower or fraction...

46 CFR 160.035-5 - Construction of steel motor-propelled lifeboats with and without radio cabin.

Code of Federal Regulations, 2010 CFR

2010-10-01

... reinforced plastic, it shall be made of fire retardant material. The top of the engine box shall be fitted... tanks shall have a thickness of not less than 0.187 inch. The resins used shall be of a fire retardant... cloth shall be used. Tank laminates shall not be constructed exclusively with fibrous glass fabrics. An...

JANNAF 35th Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor); Rognan, Melanie (Editor)

1998-01-01

Volume 1, the first of two volumes is a compilation of 63 unclassified/unlimited distribution technical papers presented at the 35th meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Combustion Subcommittee (CS) held jointly with the 17th Propulsion Systems Hazards Subcommittee (PSHS) and Airbreathing Propulsion Subcommittee (APS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include solid gun propellant processing, ignition and combustion, charge concepts, barrel erosion and flash, gun interior ballistics, kinetics and molecular modeling, ETC gun modeling, simulation and diagnostics, and liquid gun propellant combustion; solid rocket motor propellant combustion, combustion instability fundamentals, motor instability, and measurement techniques; and liquid and hybrid rocket combustion.

Recommendations for Safe Separation Distances from the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB) Using a Heat-Flux-Based Analytical Approach (Abridged)

NASA Technical Reports Server (NTRS)

Cragg, Clinton H.; Bowman, Howard; Wilson, John E.

2011-01-01

The NASA Engineering and Safety Center (NESC) was requested to provide computational modeling to support the establishment of a safe separation distance surrounding the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB). The two major objectives of the study were 1) establish a methodology based on thermal flux to determine safe separation distances from the Kennedy Space Center's (KSC's) Vehicle Assembly Building (VAB) with large numbers of solid propellant boosters containing hazard division 1.3 classification propellants, in case of inadvertent ignition; and 2) apply this methodology to the consideration of housing eight 5-segment solid propellant boosters in the VAB. The results of the study are contained in this report.

Developing the World's Most Powerful Solid Booster

NASA Technical Reports Server (NTRS)

Priskos, Alex S.; Frame, Kyle L.

2016-01-01

NASA's Journey to Mars has begun. Indicative of that challenge, this will be a multi-decadal effort requiring the development of technology, operational capability, and experience. The first steps are underway with more than 15 years of continuous human operations aboard the International Space Station (ISS) and development of commercial cargo and crew transportation capabilities. NASA is making progress on the transportation required for deep space exploration - the Orion crew spacecraft and the Space Launch System (SLS) heavy-lift rocket that will launch Orion and large components such as in-space stages, habitat modules, landers, and other hardware necessary for deep-space operations. SLS is a key enabling capability and is designed to evolve with mission requirements. The initial configuration of SLS - Block 1 - will be capable of launching more than 70 metric tons (t) of payload into low Earth orbit, greater mass than any other launch vehicle in existence. By enhancing the propulsion elements and larger payload fairings, future SLS variants will launch 130 t into space, an unprecedented capability that simplifies hardware design and in-space operations, reduces travel times, and enhances two solid propellant five-segment boosters, both based on space shuttle technologies. This paper will focus on development of the booster, which will provide more than 75 percent of total vehicle thrust at liftoff. Each booster is more than 17 stories tall, 3.6 meters (m) in diameter and weighs 725,000 kilograms (kg). While the SLS booster appears similar to the shuttle booster, it incorporates several changes. The additional propellant segment provides additional booster performance. Parachutes and other hardware associated with recovery operations have been deleted and the booster designated as expendable for affordability reasons. The new motor incorporates new avionics, new propellant grain, asbestos-free case insulation, a redesigned nozzle, streamlined manufacturing processes, and new inspection techniques. New materials and processes provide improved performance, safety, and affordability but also have led to challenges for the government/industry development team. The team completed its first full-size qualification motor test firing in early 2015. The second is scheduled for mid-2016. This paper will discuss booster accomplishments to date, as well as challenges and milestones ahead.

High-Pressure Burning Rate Studies of Solid Rocket Propellants

DTIC Science & Technology

2013-01-01

monopropellant burning rate. The self-de§agration rates of neat AP are plotted in Fig. 2 for both pressed pellets and single crystals. There is agreement...rate data from various investigators: 1 ¡ [2]; pressed pellets : 2 ¡ [3], 3 ¡ [4], and 4 ¡ [2]; and single crystals: 5 ¡ [5], and 6 ¡ [6]. Line ¡ AP...7]. Strand or window burners have had more use in the solid propellant community. There are numerous types and styles of combustion vessels, but they

Energetic Materials - Technology, Manufacturing and Processing, 27th International Annual Conference of ICT June 25 - June 28, 1996 Karisruhe, Federal Republic of Germany.

DTIC Science & Technology

1997-02-01

PROPELLANTS WITH VARYING COMPOSITION L.F. Dimaranan, I. Lee, F.E. Hudson III V12 A COMBUSTION MODEL FOR AN/HTPB-IPDI COMPOSITE SOLID PROPELLANTS P... COMPOSITE PROPELLANTS WITH A LOW PRESSURE EXPONENT SUITABLE FOR NOZZLELESS BOOSTER MOTORS G.J. van Zyl V21 PROPERTIES OF AN AND PSAN/GAP-PROPELLANTS K...APPLICATION B.N. Kondrikov SENSITIVITY TO PROJECTILE IMPACT OF PRE-HEATED EXPLOSIVE COMPOSITIONS H.Cherin, D. Lemoine, L. Gautier VULNERABILITY TESTING OF

SRM propellant, friction/ESD testing

NASA Technical Reports Server (NTRS)

Campbell, L. A.

1989-01-01

Following the Pershing 2 incident in 1985 and the Peacekeeper ignition during core removal in 1987, it was found that propellant can be much more sensitive to Electrostatic Discharges (ESD) than ever before realized. As a result of the Peacekeeper motor near miss incident, a friction machine was designed and fabricated, and used to determine friction hazards during core removal. Friction testing with and electrical charge being applied across the friction plates resulted in propellant ignitions at low friction pressures and extremely low ESD levels. The objective of this test series was to determine the sensitivity of solid rocket propellant to combined friction pressure and electrostatic stimuli and to compare the sensitivity of the SRM propellant to Peacekeeper propellant. The tests are fully discussed, summarized and conclusions drawn.

Report on JANNAF panel on shotgun/relative quickness testing

NASA Technical Reports Server (NTRS)

Gould, R. A.

1980-01-01

As the need for more energetic solid propellants continues, a number of problems arises. One of these is the tendency of high energy propellants to transition from burning (deflagration) to detonation in regions where the propellant is present in small particle sizes; e.g., in case bonding areas of a motor after a rapid depressurization causes a shear zone at the bond interface as the stressed propellant and motor case relax at different rates. In an effort to determine the susceptibility of propellants to high strain rate break up (friability), and subsequent DDT, the propulsion community uses the shotgun/relative quickness test as one of a number of screening tests for new propellant formulations. Efforts to standardize test techniques and equipment are described.

Space Launch System Booster Passes Major Ground Test

NASA Image and Video Library

2015-03-11

The largest, most powerful rocket booster ever built successfully fired up Wednesday for a major-milestone ground test in preparation for future missions to help propel NASA’s Space Launch System (SLS) rocket and Orion spacecraft to deep space destinations, including an asteroid and Mars. The booster fired for two minutes, the same amount of time it will fire when it lifts the SLS off the launch pad, and produced about 3.6 million pounds of thrust. The test was conducted at the Promontory, Utah test facility of commercial partner Orbital ATK.

Measurement of Solid Rocket Propellant Burning Rate Using X-ray Imaging

NASA Astrophysics Data System (ADS)

Denny, Matthew D.

The burning rate of solid propellants can be difficult to measure for unusual burning surface geometries, but X-ray imaging can be used to measure burning rate. The objectives of this work were to measure the baseline burning rate of an electrically-controlled solid propellant (ESP) formulation with real-time X-ray radiography and to determine the uncertainty of the measurements. Two edge detection algorithms were written to track the burning surface in X-ray videos. The edge detection algorithms were informed by intensity profiles of simulated 2-D X-ray images. With a 95% confidence level, the burning rates measured by the Projected-Slope Intersection algorithm in the two combustion experiments conducted were 0.0839 in/s +/-2.86% at an average pressure of 407 psi +/-3.6% and 0.0882 in/s +/-3.04% at 410 psi +/-3.9%. The uncertainty percentages were based on the statistics of a Monte Carlo analysis on burning rate.

Materials for Liquid Propulsion Systems. Chapter 12

NASA Technical Reports Server (NTRS)

Halchak, John A.; Cannon, James L.; Brown, Corey

2016-01-01

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

Modeling and testing of a tube-in-tube separation mechanism of bodies in space

NASA Astrophysics Data System (ADS)

Michaels, Dan; Gany, Alon

2016-12-01

A tube-in-tube concept for separation of bodies in space was investigated theoretically and experimentally. The separation system is based on generation of high pressure gas by combustion of solid propellant and restricting the expansion of the gas only by ejecting the two bodies in opposite directions, in such a fashion that maximizes generated impulse. An interior ballistics model was developed in order to investigate the potential benefits of the separation system for a large range of space body masses and for different design parameters such as geometry and propellant. The model takes into account solid propellant combustion, heat losses, and gas phase chemical reactions. The model shows that for large bodies (above 100 kg) and typical separation velocities of 5 m/s, the proposed separation mechanism may be characterized by a specific impulse of 25,000 s, two order of magnitude larger than that of conventional solid rockets. It means that the proposed separation system requires only 1% of the propellant mass that would be needed for a conventional rocket for the same mission. Since many existing launch vehicles obtain such separation velocities by using conventional solid rocket motors (retro-rockets), the implementation of the new separation system design can reduce dramatically the mass of the separation system and increase safety. A dedicated experimental setup was built in order to demonstrate the concept and validate the model. The experimental results revealed specific impulse values of up to 27,000 s and showed good correspondence with the model.

A research on polyether glycol replaced APCP rocket propellant

NASA Astrophysics Data System (ADS)

Lou, Tianyou; Bao, Chun Jia; Wang, Yiyang

2017-08-01

Ammonium perchlorate composite propellant (APCP) is a modern solid rocket propellant used in rocket vehicles. It differs from many traditional solid rocket propellants by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing it with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry. For traditional APCP, ingredients normally used are ammonium peroxide, aluminum, Hydroxyl-terminated polybutadiene(HTPB), curing agency and other additives, the greatest disadvantage is that the fuel is too expensive. According to the price we collected in our country, a single kilogram of this fuel will cost 200 Yuan, which is about 35 dollars, for a fan who may use tons of the fuel in a single year, it definitely is a great deal of money. For this reason, we invented a new kind of APCP fuel. Changing adhesive agency from cross-linked htpb to cross linked polyether glycol gives a similar specific thrust, density and mechanical property while costs a lower price.

Alternate propellants for the space shuttle solid rocket booster motors. [for reducing environmental impact of launches

NASA Technical Reports Server (NTRS)

1973-01-01

As part of the Shuttle Exhaust Effects Panel (SEEP) program for fiscal year 1973, a limited study was performed to determine the feasibility of minimizing the environmental impact associated with the operation of the solid rocket booster motors (SRBMs) in projected space shuttle launches. Eleven hypothetical and two existing limited-experience propellants were evaluated as possible alternates to a well-proven state-of-the-art reference propellant with respect to reducing emissions of primary concern: namely, hydrogen chloride (HCl) and aluminum oxide (Al2O3). The study showed that it would be possible to develop a new propellant to effect a considerable reduction of HCl or Al2O3 emissions. At the one extreme, a 23% reduction of HCl is possible along with a ll% reduction in Al2O3, whereas, at the other extreme, a 75% reduction of Al2O3 is possible, but with a resultant 5% increase in HCl.

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.

Check Firing of Master and Reference Propellants

DTIC Science & Technology

2014-12-08

estimate amount and color of smoke generated. s. Visually estimate amount and color of muzzle flash generated. TOP 04-2-607A 8 December 2014 10...Stargauge. TOP 04-2-607A 8 December 2014 3 e. Pressure gauge (piezoelectric preferred, and/or crusher). f. Muzzle velocity radar unit...firing: a. Provide electronics personnel with the weapon caliber and type, the weight and model of projectile, and expected muzzle velocities. b

Hot-Fire Testing of 100 LB(sub F) LOX/LCH4 Reaction Control Engine at Altitude Conditions

NASA Technical Reports Server (NTRS)

Marshall, William M.; Kleinhenz, Julie E.

2010-01-01

Liquid oxygen/liquid methane (LO2/LCH4 ) has recently been viewed as a potential green propulsion system for both the Altair ascent main engine (AME) and reaction control system (RCS). The Propulsion and Cryogenic Advanced Development Project (PCAD) has been tasked by NASA to develop these green propellant systems to enable safe and cost effective exploration missions. However, experience with LO2/LCH4 as a propellant combination is limited, so testing of these systems is critical to demonstrating reliable ignition and performance. A test program of a 100 lb f reaction control engine (RCE) is underway at the Altitude Combustion Stand (ACS) of the NASA Glenn Research Center, with a focus on conducting tests at altitude conditions. These tests include a unique propellant conditioning feed system (PCFS) which allows for the inlet conditions of the propellant to be varied to test warm to subcooled liquid propellant temperatures. Engine performance, including thrust, c* and vacuum specific impulse (I(sub sp,vac)) will be presented as a function of propellant temperature conditions. In general, the engine performed as expected, with higher performance at warmer propellant temperatures but better efficiency at lower propellant temperatures. Mixture ratio effects were inconclusive within the uncertainty bands of data, but qualitatively showed higher performance at lower ratios.

A Study of Fluid Interface Configurations in Exploration Vehicle Propellant Tanks

NASA Technical Reports Server (NTRS)

Zimmerli, Gregory A.; Asipauskas, Marius; Chen, Yongkang; Weislogel, Mark M.

2010-01-01

The equilibrium shape and location of fluid interfaces in spacecraft propellant tanks while in low-gravity is of interest to system designers, but can be challenging to predict. The propellant position can affect many aspects of the spacecraft such as the spacecraft center of mass, response to thruster firing due to sloshing, liquid acquisition, propellant mass gauging, and thermal control systems. We use Surface Evolver, a fluid interface energy minimizing algorithm, to investigate theoretical equilibrium liquid-vapor interfaces for spacecraft propellant tanks similar to those that have been considered for NASA's new class of Exploration vehicles. The choice of tank design parameters we consider are derived from the NASA Exploration Systems Architecture Study report. The local acceleration vector employed in the computations is determined by estimating low-Earth orbit (LEO) atmospheric drag effects and centrifugal forces due to a fixed spacecraft orientation with respect to the Earth or Moon, and rotisserie-type spacecraft rotation. Propellant/vapor interface positions are computed for the Earth Departure Stage and Altair lunar lander descent and ascent stage tanks for propellant loads applicable to LEO and low-lunar orbit. In some of the cases investigated the vapor ullage bubble is located at the drain end of the tank, where propellant management device hardware is often located.

Numerical and experimental analysis of heat transfer in injector plate of hydrogen peroxide hybrid rocket motor

NASA Astrophysics Data System (ADS)

Cai, Guobiao; Li, Chengen; Tian, Hui

2016-11-01

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

Study of the Deposition of Ammonium Perchlorate Following the Static Firing of MK-58 Rocket Motors

DTIC Science & Technology

2008-10-01

hyperthyroidism , gas generators, electrolytes for lithium cells, and as chemical reagents. The occurrence of perchlorate in the environment is...contain 19.6 kg of cross-linked double based propellants, with 7.8 % by weight of binder, 62 % of RDX, 25.88 % of plasticizers, and a few percent...compound is present at less than 1% by weight in the propellant but it might be considered as a health hazard since it has adverse toxicological impacts

Designing Small Propellers for Optimum Efficiency and Low Noise Footprint

DTIC Science & Technology

2015-06-26

each one. The GUI contains input boxes for all of the necessary data in order to run QMIL, QPROP, NAFNoise, and to produce Visual Basic ( VBA ) code... VBA macros that will automatically place reference planes for each airfoil section and insert the splined airfoils to their respective reference...Figure 24. Solid propeller exa mple. Figure 25. Hub and spoke propeller design. Figure 26. Alumninum hub design. accessed on May 12, 2015. DC, August

Experimental Investigation of the Interaction of Electrothermal Plasmas with Solid Propellants

DTIC Science & Technology

2007-09-14

formation increases propellant burning rate (Koleczko, et al . 2001). The experiments described here were designed to create time and spatially resolved...Pesce-Rodriguez 2004, Koleczko, et al . 2001). Most tests involving plasma propellant interactions involve higher plasma energies than the 3.1 kJ of...product that scatters light. The large jump in pressurization seen in closed bomb plasma ignition tests (Lieb, et al . 2001) during the plasma discharge

A stop-restart solid propellant study with salt quench

NASA Technical Reports Server (NTRS)

Kumar, R. N.

1976-01-01

Experiments were conducted to gain insight into the unsatisfactory performance of the salt quench system of solid propellants in earlier studies. Nine open-air salt spray tests were conducted and high-speed cinematographic coverage was obtained of the events. It is shown that the salt spray by the detonator is generally a two-step process yielding two different fractions. The first fraction consists of finely powdered salt and moves practically unidirectionally at a high velocity (thousand of feet per second) while the second fraction consists of coarse particles and moves randomly at a low velocity (a few feet per second). Further investigation is required to verify the speculation that a lower quench charge ratio (weight of salt/propellant burning area) than previously employed may lead to an efficient quench

Analysis of pressure blips in aft-finocyl solid rocket motor

NASA Astrophysics Data System (ADS)

Di Giacinto, M.; Favini, B.; Cavallini, E.

2016-07-01

Ballistic anomalies have frequently occurred during the firing of several solid rocket motors (SRMs) (Inertial Upper Stage, Space Shuttle Redesigned SRM (RSRM) and Titan IV SRM Upgrade (SRMU)), producing even relevant and unexpected variations of the SRM pressure trace from its nominal profile. This paper has the purpose to provide a numerical analysis of the following possible causes of ballistic anomalies in SRMs: an inert object discharge, a slag ejection, and an unexpected increase in the propellant burning rate or in the combustion surface. The SRM configuration under investigation is an aft-finocyl SRM with a first-stage/small booster design. The numerical simulations are performed with a quasi-one-dimensional (Q1D) unsteady model of the SRM internal ballistics, properly tailored to model each possible cause of the ballistic anomalies. The results have shown that a classification based on the head-end pressure (HEP) signature, relating each other the HEP shape and the ballistic anomaly cause, can be made. For each cause of ballistic anomalies, a deepened discussion of the parameters driving the HEP signatures is provided, as well as qualitative and quantitative assessments of the resultant pressure signals.

A Monte Carlo investigation of thrust imbalance of solid rocket motor pairs

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.; Johnson, J. S., Jr.

1974-01-01

A technique is described for theoretical, statistical evaluation of the thrust imbalance of pairs of solid-propellant rocket motors (SRMs) firing in parallel. Sets of the significant variables, determined as a part of the research, are selected using a random sampling technique and the imbalance calculated for a large number of motor pairs. The performance model is upgraded to include the effects of statistical variations in the ovality and alignment of the motor case and mandrel. Effects of cross-correlations of variables are minimized by selecting for the most part completely independent input variables, over forty in number. The imbalance is evaluated in terms of six time - varying parameters as well as eleven single valued ones which themselves are subject to statistical analysis. A sample study of the thrust imbalance of 50 pairs of 146 in. dia. SRMs of the type to be used on the space shuttle is presented. The FORTRAN IV computer program of the analysis and complete instructions for its use are included. Performance computation time for one pair of SRMs is approximately 35 seconds on the IBM 370/155 using the FORTRAN H compiler.

Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

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

Kinefuchi, K.; Funaki, I.; Shimada, T.

Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model.more » The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.« less

46 CFR 76.05-20 - Fixed fire extinguishing systems.

Code of Federal Regulations, 2010 CFR

2010-10-01

... all self-propelled vessels and on all barges with sleeping accommodations for more than six persons. Previously approved installations may be retained as long as they are maintained in good condition to the...

46 CFR 76.05-20 - Fixed fire extinguishing systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

... all self-propelled vessels and on all barges with sleeping accommodations for more than six persons. Previously approved installations may be retained as long as they are maintained in good condition to the...

46 CFR 76.05-20 - Fixed fire extinguishing systems.

Code of Federal Regulations, 2011 CFR

2011-10-01

... all self-propelled vessels and on all barges with sleeping accommodations for more than six persons. Previously approved installations may be retained as long as they are maintained in good condition to the...

46 CFR 76.05-20 - Fixed fire extinguishing systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

... all self-propelled vessels and on all barges with sleeping accommodations for more than six persons. Previously approved installations may be retained as long as they are maintained in good condition to the...

46 CFR 76.05-20 - Fixed fire extinguishing systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... all self-propelled vessels and on all barges with sleeping accommodations for more than six persons. Previously approved installations may be retained as long as they are maintained in good condition to the...

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.

The Influence of Glove Type on Simulated Wheelchair Racing Propulsion: A Pilot Study.

PubMed

Rice, I; Dysterheft, J; Bleakney, A W; Cooper, R A

2016-01-01

Our purpose was to examine the influence of glove type on kinetic and spatiotemporal parameters at the handrim in elite wheelchair racers. Elite wheelchair racers (n=9) propelled on a dynamometer in their own racing chairs with a force and moment sensing wheel attached. Racers propelled at 3 steady state speeds (5.36, 6.26 & 7.60 m/s) and performed one maximal effort sprint with 2 different glove types (soft & solid). Peak resultant force, peak torque, impulse, contact angle, braking torque, push time, velocity, and stroke frequency were recorded for steady state and sprint conditions. Multiple nonparametric Wilcoxon matched pair's tests were used to detect differences between glove types, while effect sizes were calculated based on Cohen's d. During steady state trials, racers propelled faster, using more strokes and larger contact angle, while applying less impulse with solid gloves compared to soft gloves. During the sprint condition, racers achieved greater top end velocities, applying larger peak force, with less braking torque with solid gloves compared to soft gloves. Use of solid gloves may provide some performance benefits to wheelchair racers during steady state and top end velocity conditions. © Georg Thieme Verlag KG Stuttgart · New York.

Space Shuttle Projects

NASA Image and Video Library

1977-12-01

The solid rocket booster (SRB) structural test article is being installed in the Solid Rocket Booster Test Facility for the structural and load verification test at the Marshall Space Flight Center (MSFC). The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment.

Solid Rocket Booster Structural Test Article

NASA Technical Reports Server (NTRS)

1978-01-01

The structural test article to be used in the solid rocket booster (SRB) structural and load verification tests is being assembled in a high bay building of the Marshall Space Flight Center (MSFC). The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment.

Solid Hydrogen Experiments for Atomic Propellants: Image Analyses

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2002-01-01

This paper presents the results of detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Solid particles of hydrogen were frozen in liquid helium, and observed with a video camera. The solid hydrogen particle sizes, their agglomerates, and the total mass of hydrogen particles were estimated. Particle sizes of 1.9 to 8 mm (0.075 to 0.315 in.) were measured. The particle agglomerate sizes and areas were measured, and the total mass of solid hydrogen was computed. A total mass of from 0.22 to 7.9 grams of hydrogen was frozen. Compaction and expansion of the agglomerate implied that the particles remain independent particles, and can be separated and controlled. These experiment image analyses are one of the first steps toward visually characterizing these particles, and allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Operational Concept Evaluation of Solid Oxide Fuel Cells for Space Vehicle Applications

NASA Technical Reports Server (NTRS)

Poast, Kenneth I.

2011-01-01

With the end of the Space Shuttle Program, NASA is evaluating many different technologies to support future missions. Green propellants, like liquid methane and liquid oxygen, have potential advantages for some applications. A Lander propelled with LOX/methane engines is one such application. When the total vehicle design and infrastructure are considered, the advantages of the integration of propulsion, heat rejection, life support and power generation become attractive for further evaluation. Scavenged residual propellants from the propulsion tanks could be used to generate needed electric power, heat and water with a Solid Oxide Fuel Cell(SOFC). In-Situ Resource Utilization(ISRU) technologies may also generate quantities of green propellants to refill these tanks and/or supply these fuel cells. Technology demonstration projects such as the Morpheus Lander are currently underway to evaluate the practicality of such designs and operational concepts. Tethered tests are currently in progress on this vertical test bed to evaluate the propulsion and avionics systems. Evaluation of the SOFC seeks to determine the feasibility of using these green propellants to supply power and identify the limits to the integration of this technology into a space vehicle prototype.

Development of Erosive Burning Models for CFD Predictions of Solid Rocket Motor Internal Environments

NASA Technical Reports Server (NTRS)

Wang, Qun-Zhen

2003-01-01

Four erosive burning models, equations (11) to (14). are developed in this work by using a power law relationship to correlate (1) the erosive burning ratio and the local velocity gradient at propellant surfaces; (2) the erosive burning ratio and the velocity gradient divided by centerline velocity; (3) the erosive burning difference and the local velocity gradient at propellant surfaces; and (4) the erosive burning difference and the velocity gradient divided by centerline velocity. These models depend on the local velocity gradient at the propellant surface (or the velocity gradient divided by centerline velocity) only and, unlike other empirical models, are independent of the motor size. It was argued that, since the erosive burning is a local phenomenon occurring near the surface of the solid propellant, the erosive burning ratio should be independent of the bore diameter if it is correlated with some local flow parameters such as the velocity gradient at the propellant surface. This seems to be true considering the good results obtained by applying these models, which are developed from the small size 5 inch CP tandem motor testing, to CFD simulations of much bigger motors.

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.

America's first long-range-missile and space exploration program: The ORDCIT project of the Jet Propulsion Laboratory, 1943 - 1946: A memoir

NASA Technical Reports Server (NTRS)

Malina, F. J.

1977-01-01

Research and achievements of the wartime Jet Propulsion Laboratory are outlined. Accomplishments included development of the solid-propellant Private A and private R rockets and the liquid-propellant nitric acid-aniline WAC Corporal rocket.

Experimental investigation of a solid rocket combustion simulator

NASA Technical Reports Server (NTRS)

Frederick, Robert A., Jr.

1991-01-01

The response of solid rocket motor materials to high-temperature corrosive gases is usually accomplished by testing the materials in a subscale solid rocket motor. While this imposes the proper thermal and chemical environment, a solid rocket motor does not provide practical features that would enhance systematic evaluations such as: the ability to throttle for margin testing, on/off capability, low test cost, and a low-hazards test article. Solid Rocket Combustion Simulators (SRCS) are being evaluated by NASA to test solid rocket nozzle materials and incorporate these essential practical features into the testing of rocket materials. The SRCS is designed to generate the thermochemical environment of a solid rocket. It uses hybrid rocket motor technology in which gaseous oxygen (Gox) is injected into a chamber containing a solid fuel grain. Specific chemicals are injected in the aft mixing chamber so that the gases entering the test section match the temperature and a non-dimensional erosion factor B' to insure similarity with a solid motor. Because the oxygen flow can be controlled, this approach allows margin testing, the ability to throttle, and an on/off capability. The fuel grains are inert which makes the test article very safe to handle. The objective of this work was to establish the baseline operating characteristics of a Labscale Solid Rocket Combustion Simulator (LSRCS). This included establishing the baseline burning rates of plexiglass fuels and the evaluation of a combustion instability for hydroxy-terminated polybutadyene (HTPB) propellants. The scope of the project included: (1) activation of MSFC Labscale Hybrid Combustion Simulator; (2) testing of plexiglass fuel at Gox ranges from 0.025 to 0.200 lb/s; (3) burning HTPB fuels at a Gox rate of 0.200 lb/s using four different mixing chamber configurations; and (4) evaluating the fuel regression and chamber pressure responses of each firing.

Fine-Water-Mist Multiple-Orientation-Discharge Fire Extinguisher

NASA Technical Reports Server (NTRS)

Butz, James R.; Turchi, Craig S.; Kimball, Amanda; McKinnon, Thomas; Riedel, Edward

2010-01-01

A fine-water-mist fire-suppression device has been designed so that it can be discharged uniformly in any orientation via a high-pressure gas propellant. Standard fire extinguishers used while slightly tilted or on their side will not discharge all of their contents. Thanks to the new design, this extinguisher can be used in multiple environments such as aboard low-gravity spacecraft, airplanes, and aboard vehicles that may become overturned prior to or during a fire emergency. Research in recent years has shown that fine water mist can be an effective alternative to Halons now banned from manufacture. Currently, NASA uses carbon dioxide for fire suppression on the International Space Station (ISS) and Halon chemical extinguishers on the space shuttle. While each of these agents is effective, they have drawbacks. The toxicity of carbon dioxide requires that the crew don breathing apparatus when the extinguishers are deployed on the ISS, and Halon use in future spacecraft has been eliminated because of international protocols on substances that destroy atmospheric ozone. A major advantage to the new system on occupied spacecraft is that the discharged system is locally rechargeable. Since the only fluids used are water and nitrogen, the system can be recharged from stores of both carried aboard the ISS or spacecraft. The only support requirement would be a pump to fill the water and a compressor to pressurize the nitrogen propellant gas. This system uses a gaseous agent to pressurize the storage container as well as to assist in the generation of the fine water mist. The portable fire extinguisher hardware works like a standard fire extinguisher with a single storage container for the agents (water and nitrogen), a control valve assembly for manual actuation, and a discharge nozzle. The design implemented in the proof-of-concept experiment successfully extinguished both open fires and fires in baffled enclosures.

Design and use of a servo-controlled high pressure window bomb in spectroscopic studies of solid propellant combustion

NASA Technical Reports Server (NTRS)

Goetz, F.; Mann, D. M.

1980-01-01

The feasibility of using a high pressure window bomb as a laboratory scale model of actual motor conditions. The design and operation of a modified high pressure window bomb is discussed. An optical servocontrol mechanism has been designed to hold the burning surface of a propellant strand at a fixed position within the bomb chamber. This mechanism permits the recording of visible and infrared emission spectra from various propellants. Preliminary visible emission spectra of a nonmetalized and metalized propellant are compared with spectra recorded using the modified bomb.

The Damage Law of HTPB Propellant under Thermomechanical Loading

NASA Astrophysics Data System (ADS)

Liu, Cheng-wu; Yang, Jian-hong; Wang, Xian-meng; Ma, Yong-kang

2016-01-01

By way of measuring the acoustic emission (AE) signals of Hydroxyl-terminated polybutadiene (HTPB) propellant in condition of uniform speed, and combined with the scanning electron microscopy (SEM) fracture surface observation, the damage law of HTPB composite solid propellant under thermomechanical loading was studied. The results show that the effects of thermomechanical loading on HTPB propellant are related to the time and can be divided into three different stages. In the first stage, thermal air aging dominates; in the second stage, interface damage is dominant; and in the third stage, thermal air aging is once again dominant.

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.

Combustion engine for solid and liquid fuels

NASA Technical Reports Server (NTRS)

Pabst, W.

1986-01-01

A combustion engine having no piston, a single cylinder, and a dual-action, that is applicable for solid and liquid fuels and propellants, and that functions according to the principle of annealing point ignition is presented. The invention uses environmentally benign amounts of fuel and propellants to produce gas and steam pressure, and to use a simple assembly with the lowest possible consumption and constant readiness for mixing and burning. The advantage over conventional combustion engines lies in lower consumption of high quality igniting fluid in the most cost effective manner.

Optimum Distribution of Metal Particles in the Solid-Propellant Charge in the Approximation of a One-Dimensional Flow Field in a Cylindrical Channel

NASA Astrophysics Data System (ADS)

Min'kov, L. L.; Shrager, É. R.

2015-03-01

A study has been made of ways of optimum distribution of particles of dispersed metal in the solid-propellant charge with a cylindrical central channel, which is firmly fastened to the case. The efficiency of combustion of this metal has been analyzed. Consideration has been given to the influence of the dynamic nonequilibrium of two-phase flow on the optimum distribution of metal particles in the indicated charge in the approximation of one-dimensionality of the flow field.

Research on combustion instability and application to solid propellant rocket motors. II.

NASA Technical Reports Server (NTRS)

Culick, F. E. C.

1972-01-01

Review of the current state of analyses of combustion instability in solid-propellant rocket motors, citing appropriate measurements and observations. The work discussed has become increasingly important, both for the interpretation of laboratory data and for predicting the transient behavior of disturbances in full-scale motors. Two central questions are considered - namely, linear stability and nonlinear behavior. Several classes of problems are discussed as special cases of a general approach to the analysis of combustion instability. Application to motors, and particularly the limitations presently understood, are stressed.

Modeling and simulation of combustion chamber and propellant dynamics and issues in active control of combustion instabilities

NASA Astrophysics Data System (ADS)

Isella, Giorgio Carlo

A method for a comprehensive approach to analysis of the dynamics of an actively controlled combustion chamber, with detailed analysis of the combustion models for the case of a solid rocket propellant, is presented here. The objective is to model the system as interconnected blocks describing the dynamics of the chamber, combustion and control. The analytical framework for the analysis of the dynamics of a combustion chamber is based on spatial averaging, as introduced by Culick. Combustion dynamics are analyzed for the case of a solid propellant. Quasi-steady theory is extended to include the dynamics of the gas-phase and also of a surface layer. The models are constructed so that they produce a combustion response function for the solid propellant that can be immediately introduced in the our analytical framework. The principal objective mechanisms responsible for the large sensitivity, observed experimentally, of propellant response to small variations. We show that velocity coupling, and not pressure coupling, has the potential to be the mechanism responsible for that high sensitivity. We also discuss the effect of particulate modeling on the global dynamics of the chamber and revisit the interpretation of the intrinsic stability limit for burning of solid propellants. Active control is also considered. Particular attention is devoted to the effect of time delay (between sensing and actuation); several methods to compensate for it are discussed, with numerical examples based on the approximate analysis produced by our framework. Experimental results are presented for the case of a Dump Combustor. The combustor exhibits an unstable burning mode, defined through the measurement of the pressure trace and shadowgraph imaging. The transition between stable and unstable modes of operation is characterized by the presence of hysteresis, also observed in other experimental works, and hence not a special characteristic of this combustor. Control is introduced in the form of pulsed secondary fuel. We show the capability of forcing the transition from unstable to stable burning, hence extending the stable operating regime of the combustor. The transition, characterized by the use of a shadowgraph movie sequence, is attributed to a combined fluid-mechanic and combustion mechanism.

Diagnostic developments for velocity and temperature measurements in uni-element rocket environments

NASA Astrophysics Data System (ADS)

Philippart, Kenneth D.

1995-08-01

Velocity and temperature measurements were taken within a uni-element rocket combustion chamber for hydrogen-oxygen propellants using laser Doppler velocimetry, thermocouples, and a thermocouple-based temperature rake developed for this effort. Velocity and turbulence profiles were obtained for firings with a gaseous oxygen (GO2)/gaseous hydrogen (GH2) coaxial shear injector at axial locations of 1.6 mm (0.063 in.), 6.4 mm (0.25 in.), 12.7 mm (0.5 in.), 25.4 mm (1 in.) and 50.8 mm (2 in.). Aluminum oxide particles of various sizes seeded the flow in an attempt to explain the discrepancies. While cold-flow simulations were promising, hot-fire results for the various particles were virtually identical and still lower than earlier data. The hot-firings were self-consistent and question the reproducibility of the previous data. Velocity measurements were made closer to the injector than the preceding work. Asymmetries were noted in all profiles. The shear layer displayed high turbulence levels. The central flow near the injector resembled turbulent pipe flow. Recirculation zones existed at the chamber walls and became smaller as the flow evolved downstream. The combusting flow region expanded with increasing axial distance. A thermocouple-instrumented coaxial injector was fired with GO2/GH2 propellants. The injector exit plane boundary conditions were determined. The feasibility of a thermocouple-based temperature rake was established. Tests at three axial positions for air/GM2 firings revealed asymmetric profiles. Temperatures increased with increasing axial distance.

Experimental Evaluation of a Subscale Gaseous Hydrogen/gaseous Oxygen Coaxial Rocket Injector

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert

2002-01-01

The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

Solid Propellant Test Motor Scaling

DTIC Science & Technology

2001-09-01

50 Figure 40. Comparison of Measured and Calculated Strand and Small Motor Burning Rates for Fundamental Studies of HTPB /AP Smokeless...Propellants...................................... 51 Figure 41. Agreement Between 2x4 Motor and Strand Burning Rate Data for Non-aluminized HTPB /AP...58 Figure 51. Comparison Between Results Obtained with Ultrasonic Method and Standard

Behavior of Aluminum in Solid Propellant Combustion

DTIC Science & Technology

1982-06-01

dry pressing 30% Valley Met H- 30 aluminum, 7% carnauba wax , and 63% 100 P AP. One sample was prepared using as received H-30, a second sample used pre...34propellant" formulations. The formulations included dry pressed AP/AI, and AP/AI/ Wax samples. Sandwiches were also prepared consisting of an aluminum...Binder flame instead of by aluminum exposure during accumulate break-up. Combustion of AP/AI/ Wax Samples A set of propellant samples were prepared by

Studies on Decomposition and Combustion Mechanism of Solid Fuel Rich Propellants

DTIC Science & Technology

2010-08-30

thrust to cruise at supersonic speed. This was followed by the test of large diameter ramjet called burner test vehicle (BTV). Advanced low volume...propellant surface. Vernekar et al (43) found that in pressed AP-Al pellets , maximum burn rate is obtained at intermediate metal content. Jain et al...conjunction with high pressure window strand burner . They found that the propellant combustion was irregular and regression rate varied from 0.3 to 3

Technical Report for the Period 10 January 1959 to 30 June 1960

DTIC Science & Technology

1960-08-22

boon started to determine the efficacy of various drying procedures for polyesters. Water contents are being determined by the Karl Fischer method to an...CHARGES 17 XX.4 Inspection Methods 17 XXI SOLID PROPELLANTS FOR ROCKETS 18 XXI.1 Colloidal Propellants - Extruded 18 XXI.2 Colloidal Propellants - Cast...derivatives can be made more durable and, in particular, more resistant to heat. The method used has consisted in the preparation of crotonyl derivatives of

Test Planning, Collection, and Analysis of Pressure Data Resulting from Army Weapon Systems. Volume IV. Data Analysis of the M198 and M109 May 1979 Firings.

DTIC Science & Technology

1980-05-01

the M203 charge during May 1979 at Aberdeen Proving Ground . The data collection and analysis effort is part of a continuing program undertaken by...May to 18 May 1979 the M198 towed howitzer and the M109 self- propelled howitzer were fired with the 14203 charge at the Aberdeen Proving Grounds ...howitzer and the M109 self- propeiled howitzer were fired with the M203 charge at the Aberdeen Proving Grounds . This section of the report gives the

Integrated thruster assembly program

NASA Technical Reports Server (NTRS)

1973-01-01

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

Solid State Digital Propulsion "Cluster Thrusters" For Small Satellites Using High Performance Electrically Controlled Extinguishable Solid Propellants (ECESP)

NASA Technical Reports Server (NTRS)

Sawka, Wayne N.; Katzakian, Arthur; Grix, Charles

2005-01-01

Electrically controlled extinguishable solid propellants (ESCSP) are capable of multiple ignitions, extinguishments and throttle control by the application of electrical power. Both core and end burning no moving parts ECESP grains/motors to three inches in diameter have now been tested. Ongoing research has led to a newer family of even higher performance ECESP providing up to 10% higher performance, manufacturing ease, and significantly higher electrical conduction. The high conductivity was not found to be desirable for larger motors; however it is ideal for downward scaling to micro and pico- propulsion applications with a web thickness of less than 0.125 inch/ diameter. As a solid solution propellant, this ECESP is molecularly uniform, having no granular structure. Because of this homogeneity and workable viscosity it can be directly cast into thin layers or vacuum cast into complex geometries. Both coaxial and grain stacks have been demonstrated. Combining individual propellant coaxial grains and/or grain stacks together form three-dimensional arrays yield modular cluster thrusters. Adoption of fabless manufacturing methods and standards from the electronics industry will provide custom, highly reproducible micro-propulsion arrays and clusters at low costs. These stack and cluster thruster designs provide a small footprint saving spacecraft surface area for solar panels and/or experiments. The simplicity of these thrusters will enable their broad use on micro-pico satellites for primary propulsion, ACS and formation flying applications. Larger spacecraft may find uses for ECESP thrusters on extended booms, on-orbit refueling, pneumatic actuators, and gas generators.

Safety Practices Followed in ISRO Launch Complex- An Overview

NASA Astrophysics Data System (ADS)

Krishnamurty, V.; Srivastava, V. K.; Ramesh, M.

2005-12-01

The spaceport of India, Satish Dhawan Space Centre (SDSC) SHAR of Indian Space Research Organisation (ISRO), is located at Sriharikota, a spindle shaped island on the east coast of southern India.SDSC SHAR has a unique combination of facilities, such as a solid propellant production plant, a rocket motor static test facility, launch complexes for different types of rockets, telemetry, telecommand, tracking, data acquisition and processing facilities and other support services.The Solid Propellant Space Booster Plant (SPROB) located at SDSC SHAR produces composite solid propellant for rocket motors of ISRO. The main ingredients of the propellant produced here are ammonium perchlorate (oxidizer), fine aluminium powder (fuel) and hydroxyl terminated polybutadiene (binder).SDSC SHAR has facilities for testing solid rocket motors, both at ambient conditions and at simulated high altitude conditions. Other test facilities for the environmental testing of rocket motors and their subsystems include Vibration, Shock, Constant Acceleration and Thermal / Humidity.SDSC SHAR has the necessary infrastructure for launching satellites into low earth orbit, polar orbit and geo-stationary transfer orbit. The launch complexes provide complete support for vehicle assembly, fuelling with both earth storable and cryogenic propellants, checkout and launch operations. Apart from these, it has facilities for launching sounding rockets for studying the Earth's upper atmosphere and for controlled reentry and recovery of ISRO's space capsule reentry missions.Safety plays a major role at SDSC SHAR right from the mission / facility design phase to post launch operations. This paper presents briefly the infrastructure available at SDSC SHAR of ISRO for launching sounding rockets, satellite launch vehicles, controlled reentry missions and the built in safety systems. The range safety methodology followed as a part of the real time mission monitoring is presented. The built in safety systems provided onboard the launch vehicle are automatic shut off the propulsion system based on real time mission performance and a passivation system incorporated in the orbit insertion stage are highlighted.

Experimental investigation of atomization characteristics of swirling spray by ADN gelled propellant

NASA Astrophysics Data System (ADS)

Guan, Hao-Sen; Li, Guo-Xiu; Zhang, Nai-Yuan

2018-03-01

Due to the current global energy shortage and increasingly serious environmental issues, green propellants are attracting more attention. In particular, the ammonium dinitramide (ADN)-based monopropellant thruster is gaining world-wide attention as a green, non-polluting and high specific impulse propellant. Gel propellants combine the advantages of liquid and solid propellants, and are becoming popular in the field of spaceflight. In this paper, a swirling atomization experimental study was carried out using an ADN aqueous gel propellant under different injection pressures. A high-speed camera and a Malvern laser particle size analyzer were used to study the spray process. The flow coefficient, cone angle of swirl atomizing spray, breakup length of spray membrane, and droplet size distribution were analyzed. Furthermore, the effects of different injection pressures on the swirling atomization characteristics were studied.

Recent Developments in Chemically Reactive Sensors for Propellants

NASA Technical Reports Server (NTRS)

Davis, Dennis D.; Mast, Dion J.; Baker, David L.; Fries, Joseph (Technical Monitor)

1999-01-01

Propellant system leaks can pose a significant hazard in aerospace operations. For example, a leak in the hydrazine supply system of the shuttle auxiliary power unit (APU) has resulted in hydrazine ignition and fire in the aft compartment of the shuttle. Sensors indicating the location of a leak could provide valuable information required for operational decisions. WSTF has developed a small, single-use sensor for detection of propellant leaks. The sensor is composed of a thermistor bead coated with a substance which is chemically reactive with the propellant. The reactive thermistor is one of a pair of closely located thermistors, the other being a reference. On exposure to the propellant, the reactive coating responds exothermically to it and increases the temperature of the coated-thermistor by several degrees. The temperature rise is sensed by a resistive bridge circuit, and an alarm is registered by data acquisition software. The concept is general and has been applied to sensors for hydrazine, monomethylhydrazine, unsym-dimethylhydrazine, ammonia, hydrogen peroxide, ethanol, and dinitrogen tetroxide. Responses of these sensors to humidity, propellant concentration, distance from the liquid leak, and ambient pressure levels arc presented. A multi-use sensor has also been developed for hydrazine based on its catalytic reactivity with noble metals.

Combustion in an acceleration field: A survey of Soviet literature

NASA Technical Reports Server (NTRS)

Radloff, S. J.; Osborn, J. R.

1980-01-01

The effect of an acceleration field on the burning rate of a solid propellant was measured from -900g's to +1000g's using both double base and ammonium perchlorate based propellants. The acceleration fields were simulated using a centrifuge device and the burning rate was recorded. Both metalized and non-metalized variations of each propellant were tested and it was found that acceleration fields affect the burning rate. For the most part the theoretical predictions and the experimental results agreed.

Laser Ignition of Nitramine Composite Propellants and Crack Propagation and Branching in Burning Solid Propellants

DTIC Science & Technology

1987-10-01

34 Proceedings of the 16th JANNAF Com- bustion Meeting, Sept. 1979, Vol. II, pp. 13-34. 44. Schroeder , M. A., " Critical Analysis of Nitramine Decomposition...34 Proceedings of the 19th JANNAF Combustion Meeting, Oct. 1982. 47. Schroeder , M. A., " Critical Analysis of Nitramine Decomposition Data: Ac- tivation...the surface of the propellant. This is consis- tent with the decomposition mechanism considered by Boggs[48] and Schroeder [43J. They concluded that the

Validation of numerical simulations for nano-aluminum composite solid propellants

NASA Astrophysics Data System (ADS)

Yan, Allen H.

2011-12-01

Nano-aluminum is of interest as an energetic additive in composite solid propellant formulations for its demonstrated ability to increase combustion efficiency and burning rate. However, due to the current cost of nano-aluminum and the associated safety risks associated with propellant testing, it may not always be practical to spend the time and effort to mix, cast, and thoroughly evaluate the burning rate of a new formulation. To provide an alternative method of determining this parameter, numerical methods have been developed to predict the performance of nano-aluminum composite propellants, but these codes still require thorough validation before application. For this purpose, six propellant compositions were formulated, fully characterized, and burn rates were measured at several pressures between 34.0 and 129.3 atmospheres at room temperature, 20°C, and at an elevated temperature of 71.1°C in order to test the code's ability to predict pressure dependent burn rate and temperature sensitivity. To ensure the most accurate model possible, special emphasis was placed on characterizing the size distribution of the constituent nano-aluminum and ammonium perchlorate powders through optical diffraction or optical imaging techniques. Experimental burn rate is compared to the propellant combustion model and shows excellent agreement within 5% for a range of formulations and pressures, however under other conditions the model deviates by as much as 21%. An analysis of the results suggests that the current framework of the numerical model is unable to accurately simulate all the combustion physics of high aluminum content propellants, and suggestions for improvements are identified.

Acceleration effects in solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Langhenry, M. T.

1986-01-01

The performance variations due to acceleration loads imposed on spinning solid propellant rocket motors are investigated. The four potentially most significant modes of acceleration-induced phenomena are identified from a study of the literature and modeled. The four modes are a mechanical mode which deals with deformations of the propellant and case: a thermodynamic mode which covers acceleration-induced combustion phenomena; a stress mode which covers the stressed propellant's effect on burn rate; and a gas dynamic mode which deals with changes in gas flow in the chamber and through the nozzle. Simplified models of each mode are developed or taken from the literature and are added to an internal ballistics evaluation computer program. The resulting analysis is the first to include all of the modes. In order to do this an original analysis of the mechanical and stress modes was necessary. However, the analysis shows that the stress mode is not important for the circular perforated grains studied. The other effects are shown to have a significant influence on solid rocket motor performance. The magnitude of the different mode effects are such that one may not be ignored over the others as has been done in the past. The results of the analysis are compared to published rocket motor data. The comparisons indicate an erosive burning effect that is a function of spin rate. A qualitative explanation of the erosive effect is presented.

A Summary of NASA and USAF Hypergolic Propellant Related Spills and Fires

NASA Technical Reports Server (NTRS)

Nufer, Brian M.

2009-01-01

Several unintentional hypergolic fluid related spills, fires, and explosions from the Apollo Program, the Space Shuttle Program, the Titan Program, and a few others have occurred over the past several decades. Spill sites include the following government facilities: Kennedy Space Center (KSC), Johnson Space Center (JSC), White Sands Test Facility (WSTF), Vandenberg Air Force Base (VAFB), Cape Canaveral Air Force Station (CCAFS), Edwards Air Force Base (EAFB), Little Rock AFB, and McConnell AFB. Until now, the only method of capturing the lessons learned from these incidents has been "word of mouth" or by studying each individual incident report. The root causes and consequences of the incidents 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 internal or external human performance shaping factors 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 materials cleanliness or compatibility and chemical reactivity can result in fires or explosions (8) Improper established "back-out" and/or emergency safing procedures can escalate an event The items listed above are only a short list of the issues that should be recognized prior to handling hypergolic fluids or processing vehicles containing hypergolic propellants. The summary of incidents in this report is intended to cover many more issues than those listed above.

Analysis of rocket engine injection combustion processes

NASA Technical Reports Server (NTRS)

Salmon, J. W.

1976-01-01

A critique is given of the JANNAF sub-critical propellant injection/combustion process analysis computer models and application of the models to correlation of well documented hot fire engine data bases. These programs are the distributed energy release (DER) model for conventional liquid propellants injectors and the coaxial injection combustion model (CICM) for gaseous annulus/liquid core coaxial injectors. The critique identifies model inconsistencies while the computer analyses provide quantitative data on predictive accuracy. The program is comprised of three tasks: (1) computer program review and operations; (2) analysis and data correlations; and (3) documentation.

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.

Ammonium nitrate as an oxidizer in solid composite propellants

NASA Astrophysics Data System (ADS)

Manelis, G. B.; Lempert, D. B.

2009-09-01

Despite the fact that ammonium nitrate (AN) has the highest hydrogen content and fairly high oxygen balance (compared to other oxidizers), its extremely low formation enthalpy and relatively low density makes it one of the worst power oxidizers in solid composite propellants (SCP). Nevertheless, AN has certain advantages - the combustion of the compositions containing AN is virtually safe, its combustion products are ecologically clean, it is very accessible and cheap, and also very thermostable (far more stable than ammonium dinitramide (ADN)). Besides, its low density stops being a disadvantage if the propellant has to be used in deep space and therefore, must be carried there with other rocket carriers. The low cost of AN may also become a serious advantage in the AN application even in lower stages of multistage space launchers as well as in one-stage space launchers with low mass fraction of the propellant. The main specific features relevant to the creation of AN-based SCPs with the optimal energetic characteristics are discussed. The use of metals and their hydrides and proper fuel-binders as well as the recent successes in phase stabilization of AN are described.

Fire performance and decay resistance of solid wood and plywood treated with quaternary ammonia compounds and common fire retardants

Treesearch

Evren Terzi; S. Nami Kartal; Robert White; Katsumi Shinoda; Yuji Imamura

2010-01-01

In this study, the fire performance and decay resistance of solid wood and plywood treated with quaternary ammonia compounds (didecyl dimethyl ammonium chloride (DDAC) and didecyl dimethyl ammonium tetrafluoroborate (DBF)) were compared with the performance of untreated control specimens and specimens treated with common fire retardants ((monoammonium phosphate (MAP),...

JANNAF 36th Combustion Subcommittee Meeting. Volume 2

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

1999-01-01

Volume 11, the second of three volumes is a compilation of 33 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 36th Combustion Subcommittee held jointly with the 24 Airbreathing Propulsion Subcommittee and 18th Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered include gun solid propellant ignition and combustion, Electrothermal Chemical (ETC) propulsion phenomena, liquid propellant gun combustion and barrel erosion, gas phase propellant combustion, kinetic and decomposition phenomena and liquid and hybrid propellant combustion behavior.

Propellant Residues Deposition from Firing of AT4 Rockets

DTIC Science & Technology

2009-12-01

and 254 nm (cell path 1 cm), and a Finnigan SpectraSYSTEM AS300 autosampler. Samples were introduced with a 100-μL sample loop . Separations were...analytical laboratory. The remaining particle samples were left in sealed jars and stored on site in a refrigerator, and the snow sample was stored in a...Ranney. 1998. Characterization of antitank firing ranges at CFB Valcartier. WATC Wainwright, and CFAD Dundurn. DREV-R-9809. Val- Bélair, QC: DRDC

Development of Fuel Neutralizing Agents to Prevent Flashback on Aircraft Fires

DTIC Science & Technology

1991-05-01

Past efforts were surveyed and two approaches decided upon. In one, improved separation of the fuel and oxidizer was attempted in two ways...in a sense " shampoo " chemistry, has been used to permit the water to form an effective cover to aid in extinguishing hydrocarbon fires. What we...Propellants. ASD Technical Report, pp. 61-143, Wright-Patterson Air Force Base, Ohio, July 1961. 2. Skinner, G.B., Survey of Chemical Aspects of Flamo

40 CFR Table 3 to Subpart Ddddd of... - Work Practice Standards

Code of Federal Regulations, 2013 CFR

2013-07-01

...: natural gas, synthetic natural gas, propane, distillate oil, syngas, ultra-low sulfur diesel, fuel oil... start firing coal/solid fossil fuel, biomass/bio-based solids, heavy liquid fuel, or gas 2 (other) gases....While firing coal/solid fossil fuel, biomass/bio-based solids, heavy liquid fuel, or gas 2 (other) gases...

40 CFR Table 3 to Subpart Ddddd of... - Work Practice Standards

Code of Federal Regulations, 2014 CFR

2014-07-01

...: natural gas, synthetic natural gas, propane, distillate oil, syngas, ultra-low sulfur diesel, fuel oil... start firing coal/solid fossil fuel, biomass/bio-based solids, heavy liquid fuel, or gas 2 (other) gases....While firing coal/solid fossil fuel, biomass/bio-based solids, heavy liquid fuel, or gas 2 (other) gases...

Altitude Starting Tests of a 1000-Pound-Thrust Solid-Propellant Rocket

NASA Technical Reports Server (NTRS)

Sloop, John L.; Rollbuhler, R. James; Krawczonek, Eugene M.

1957-01-01

Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

KENNEDY SPACE CENTER, FLA. - Jeff Thon, an SRB mechanic with United Space Alliance, is fitted with a harness to test a vertical solid rocket booster propellant grain inspection technique. Thon will be lowered inside a mockup of two segments of the SRBs. The inspection of segments is required as part of safety analysis.

NASA Image and Video Library

2003-09-11

KENNEDY SPACE CENTER, FLA. - Jeff Thon, an SRB mechanic with United Space Alliance, is fitted with a harness to test a vertical solid rocket booster propellant grain inspection technique. Thon will be lowered inside a mockup of two segments of the SRBs. The inspection of segments is required as part of safety analysis.

Flight Investigation of the Performance of a Two-stage Solid-propellant Nike-deacon (DAN) Meteorological Sounding Rocket

NASA Technical Reports Server (NTRS)

Heitkotter, Robert H

1956-01-01

A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

Propellant combustion product analyses on an M16 rifle and a 105 mm caliber gun

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

Ase, P.; Eisenberg, W.; Gordon, S.

1985-01-01

Some of the propellant combustion products (particulates and gases) that are formed on firing an M16 rifle and 105 mm caliber gun have been subjected to qualitative, and to a more limited extent, quantitative chemical analyses. For both weapons, large numbers of trace gas species, 90 to 70 respectively, were identified in the combustion effluents from the small large bore weapons. Quantifiable data were obtained for 15 of these species in terms of mass of compound formed per unit mass of propellant burned. Polynuclear aromatic hydrocarbons, 11 and 4 respectively, were identified and quantified in the combustion products from themore » small and large bore weapons. Metal particulates in the respirable range in the combustion products from the M16 rifle were analyzed and quantified. Many of the chemical species identified in the study have known toxicological properties. Although the data base is limited, it appears that within the confines of the different propellants' stoichiometries, the amounts of combustion products formed are approximately directly proportional to the masses of propellant burned.« less

The development of an erosive burning model for solid rocket motors using direct numerical simulation

NASA Astrophysics Data System (ADS)

McDonald, Brian A.

A method for developing an erosive burning model for use in solid propellant design-and-analysis interior ballistics codes is described and evaluated. Using Direct Numerical Simulation, the primary mechanisms controlling erosive burning (turbulent heat transfer, and finite rate reactions) have been studied independently through the development of models using finite rate chemistry, and infinite rate chemistry. Both approaches are calibrated to strand burn rate data by modeling the propellant burning in an environment with no cross-flow, and adjusting thermophysical properties until the predicted regression rate matches test data. Subsequent runs are conducted where the cross-flow is increased from M = 0.0 up to M = 0.8. The resulting relationship of burn rate increase versus Mach Number is used in an interior ballistics analysis to compute the chamber pressure of an existing solid rocket motor. The resulting predictions are compared to static test data. Both the infinite rate model and the finite rate model show good agreement when compared to test data. The propellant considered is an AP/HTPB with an average AP particle size of 37 microns. The finite rate model shows that as the cross-flow increases, near wall vorticity increases due to the lifting of the boundary caused by the side injection of gases from the burning propellant surface. The point of maximum vorticity corresponds to the outer edge of the APd-binder flame. As the cross-flow increases, the APd-binder flame thickness becomes thinner; however, the point of highest reaction rate moves only slightly closer to the propellant surface. As such, the net increase of heat transfer to the propellant surface due to finite rate chemistry affects is small. This leads to the conclusion that augmentation of thermal transport properties and the resulting heat transfer increase due to turbulence dominates over combustion chemistry in the erosive burning problem. This conclusion is advantageous in the development of future models that can be calibrated to heat transfer conditions without the necessity for finite rate chemistry. These results are considered applicable for propellants with small, evenly distributed AP particles where the assumption of premixed APd-binder gases is reasonable.

Space Shuttle Projects

NASA Image and Video Library

1978-11-01

The structural test article to be used in the solid rocket booster (SRB) structural and load verification tests is being assembled in a high bay building of the Marshall Space Flight Center (MSFC). The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment.

Solid Hydrogen Experiments for Atomic Propellants: Particle Formation Energy and Imaging Analyses

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2002-01-01

This paper presents particle formation energy balances and detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium during the Phase II testing in 2001. Solid particles of hydrogen were frozen in liquid helium and observed with a video camera. The solid hydrogen particle sizes and the total mass of hydrogen particles were estimated. The particle formation efficiency is also estimated. Particle sizes from the Phase I testing in 1999 and the Phase II testing in 2001 were similar. Though the 2001 testing created similar particles sizes, many new particle formation phenomena were observed. These experiment image analyses are one of the first steps toward visually characterizing these particles and it allows designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Iodine Hall Thruster Propellant Feed System for a CubeSat

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of reservoir temperature/pressure, the flow resistors, and the setting of the PFCV. The calibration is performed using independent flow control monitoring techniques, providing an in situ measure of the flowrate as a function of controllable parameters. The reservoir temperature controls the iodine sublimation rate, providing propellant to ths thruster by pressurizing the propellant feed system to approx.1-2 psi. Control of the temperature and the PFCV are used to maintain reservoir pressure and keep the thruster discharge current constant.

Mini-cavity plasma core reactors for dual-mode space nuclear power/propulsion systems. M.S. Thesis

NASA Technical Reports Server (NTRS)

Chow, S.

1976-01-01

A mini-cavity plasma core reactor is investigated for potential use in a dual-mode space power and propulsion system. In the propulsive mode, hydrogen propellant is injected radially inward through the reactor solid regions and into the cavity. The propellant is heated by both solid driver fuel elements surrounding the cavity and uranium plasma before it is exhausted out the nozzle. The propellant only removes a fraction of the driver power, the remainder is transferred by a coolant fluid to a power conversion system, which incorporates a radiator for heat rejection. Neutronic feasibility of dual mode operation and smaller reactor sizes than those previously investigated are shown to be possible. A heat transfer analysis of one such reactor shows that the dual-mode concept is applicable when power generation mode thermal power levels are within the same order of magnitude as direct thrust mode thermal power levels.

Hydrodynamic capture of microswimmers into sphere-bound orbits.

PubMed

Takagi, Daisuke; Palacci, Jérémie; Braunschweig, Adam B; Shelley, Michael J; Zhang, Jun

2014-03-21

Self-propelled particles can exhibit surprising non-equilibrium behaviors, and how they interact with obstacles or boundaries remains an important open problem. Here we show that chemically propelled micro-rods can be captured, with little change in their speed, into close orbits around solid spheres resting on or near a horizontal plane. We show that this interaction between sphere and particle is short-range, occurring even for spheres smaller than the particle length, and for a variety of sphere materials. We consider a simple model, based on lubrication theory, of a force- and torque-free swimmer driven by a surface slip (the phoretic propulsion mechanism) and moving near a solid surface. The model demonstrates capture, or movement towards the surface, and yields speeds independent of distance. This study reveals the crucial aspects of activity–driven interactions of self-propelled particles with passive objects, and brings into question the use of colloidal tracers as probes of active matter.

Test data from small solid propellant rocket motor plume measurements (FA-21)

NASA Technical Reports Server (NTRS)

Hair, L. M.; Somers, R. E.

1976-01-01

A program is described for obtaining a reliable, parametric set of measurements in the exhaust plumes of solid propellant rocket motors. Plume measurements included pressures, temperatures, forces, heat transfer rates, particle sampling, and high-speed movies. Approximately 210,000 digital data points and 15,000 movie frames were acquired. Measurements were made at points in the plumes via rake-mounted probes, and on the surface of a large plate impinged by the exhaust plume. Parametric variations were made in pressure altitude, propellant aluminum loading, impinged plate incidence angle and distance from nozzle exit to plate or rake. Reliability was incorporated by continual use of repeat runs. The test setup of the various hardware items is described along with an account of test procedures. Test results and data accuracy are discussed. Format of the data presentation is detailed. Complete data are included in the appendix.

Saturn Apollo Program

NASA Image and Video Library

1965-01-01

The Saturn V first stages were test fired at the Mississippi Test Facility and at the Marshall Space Flight Center (MSFC). Five F-1 engines powered the first stage, each developing 1.5 million pounds of thrust. The first stage, known as the S-IC stage, burned over 15 tons of propellant per second during its 2.5 minutes of operation to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour. The stage was 138 feet long and 33 feet in diameter. This photograph shows the test firing of an F-1 engine at the MSFC's S-IC Static Test Firing Facility.

Preliminary Results of Solid Gas Generator Micropropulsion

NASA Technical Reports Server (NTRS)

deGroot, Wilhelmus A.; Reed, Brian D.; Brenizer, Marshall

1999-01-01

A decomposing solid thruster concept, which creates a more benign thermal and chemical environment than solid propellant combustion, while maintaining, performance similar to solid combustion, is described. A Micro-Electro-Mechanical (MEMS) thruster concept with diode laser and fiber-optic initiation is proposed, and thruster components fabricated with MEMS technology are presented. A high nitrogen content solid gas generator compound is evaluated and tested in a conventional axisymmetric thrust chamber with nozzle throat area ratio of 100. Results show incomplete decomposition of this compound in both low pressure (1 kPa) and high pressure (1 MPa) environments, with decomposition of up to 80% of the original mass. Chamber pressures of 1.1 MPa were obtained, with maximum calculated thrust of approximately 2.7 N. Resistively heated wires and resistively heated walls were used to initiate decomposition. Initiation tests using available lasers were unsuccessful, but infrared spectra of the compound show that the laser initiation tests used inappropriate wavelengths for optimal propellant absorption. Optimal wavelengths for laser ignition were identified. Data presented are from tests currently in progress. Alternative solid gas generator compounds are being evaluated for future tests.

Spectral Studies of Solid Propellant Combustion. 3. Emission and Absorption Results for HMX2 Propellant

DTIC Science & Technology

1991-05-01

funding from the Anry Productivity Capital Investment Program. vii INTENTioNALLY LEFT BL~ANK viH I. INTRODUCTION During the last several years we...G-23 Dahlgren, VA 22448-5000 1 OSD/SDIO/ IST ATTN: L. Caveny 2 Commander Pentagon Naval Surface Warfare Center Washington, DC 20301-7100 ATTN: R

Rapid response deluge system

NASA Astrophysics Data System (ADS)

Mille, J. R.

1984-08-01

The development of a rapid response deluge system by the Ammunition Equipment Directorate (AED) for use in suppressing propellant fires during demilitarization shows great promise. Prototype systems have been tested and data acquired on their efficiencies. Present system vs previous generations and lessons learned are discussed.

Detonator-activated ball shutter

DOEpatents

McWilliams, Roy A.; von Holle, William G.

1983-01-01

A detonator-activated ball shutter for closing an aperture in about 300.mu. seconds. The ball shutter containing an aperture through which light, etc., passes, is closed by firing a detonator which propels a projectile for rotating the ball shutter, thereby blocking passage through the aperture.

NASA's Space Launch System: Developing the World's Most Powerful Solid Booster

NASA Technical Reports Server (NTRS)

Priskos, Alex

2016-01-01

NASA's Journey to Mars has begun. Indicative of that challenge, this will be a multi-decadal effort requiring the development of technology, operational capability, and experience. The first steps are under way with more than 15 years of continuous human operations aboard the International Space Station (ISS) and development of commercial cargo and crew transportation capabilities. NASA is making progress on the transportation required for deep space exploration - the Orion crew spacecraft and the Space Launch System (SLS) heavy-lift rocket that will launch Orion and large components such as in-space stages, habitat modules, landers, and other hardware necessary for deep-space operations. SLS is a key enabling capability and is designed to evolve with mission requirements. The initial configuration of SLS - Block 1 - will be capable of launching more than 70 metric tons (t) of payload into low Earth orbit, greater mass than any other launch vehicle in existence. By enhancing the propulsion elements and larger payload fairings, future SLS variants will launch 130 t into space, an unprecedented capability that simplifies hardware design and in-space operations, reduces travel times, and enhances the odds of mission success. SLS will be powered by four liquid fuel RS-25 engines and two solid propellant five-segment boosters, both based on space shuttle technologies. This paper will focus on development of the booster, which will provide more than 75 percent of total vehicle thrust at liftoff. Each booster is more than 17 stories tall, 3.6 meters (m) in diameter and weighs 725,000 kilograms (kg). While the SLS booster appears similar to the shuttle booster, it incorporates several changes. The additional propellant segment provides additional booster performance. Parachutes and other hardware associated with recovery operations have been deleted and the booster designated as expendable for affordability reasons. The new motor incorporates new avionics, new propellant grain, asbestos-free case insulation, a redesigned nozzle, streamlined manufacturing processes, and new inspection techniques. New materials and processes provide improved performance, safety, and affordability but also have led to challenges for the government/industry development team. The team completed its first full-size qualification motor test firing in early 2015. The second is scheduled for mid-2016. This paper will discuss booster accomplishments to date, as well as challenges and milestones ahead.

Study of Dynamic Behavior and Hazards of VHBR, LOVA, And PSS Propellants under Rapid Ignition And Combustion Conditions by Real-Time X-Ray Radiography

DTIC Science & Technology

1991-12-01

formulation . He also discussed the recent develop - ments in VHBR research and suggested a plan for future research to answer some of the questions. METHOD OF...tests. Before VHBR propellants were developed in the early 1970’s, the only solid propellants that burned aster than 05 m/s were explosives, and...and become highly reactive free radicals. This would enhance the combustion process in the gas phase, which would in turn, increase the heat transfer

Workshop on ESD (Electrostatic Discharge) Ignition of Composite Solid Propellants Held on April 18-19, 1989 in Nashville, Tennessee

DTIC Science & Technology

1990-01-01

field) leads to microarcs, involving local breakdowns of At 2 0 3 layers on the At particles within the propellant. Cracks appear at this point, and...Propellants with AP and binder, causes a) 17. Point Breakdown an avalanche effect at a high E-field point, through) 18. Alumina Layers also angstroms, which...E-field point, through) 18. Alumina Layers (also angstroms, which then goes on to create a) 19. Discharge Path (which, given the correct conditions) D

An Investigation to Improve Quality Evaluations of Primers and Propellant for 20mm Munitions

NASA Technical Reports Server (NTRS)

Bement, L. J.; Holmes, C.; McGrory, J.; Schimmel, M. L.

1997-01-01

To reduce the frequency of electrically initiated, 20mm munition hangfires (delayed ignitions), a joint Army/NASA investigation was conducted to recommend quality evaluation improvements for acceptance of both primers and gun propellant. This effort focused only on evaluating ignition and combustion performance as potential causes of hangfires: poor electrical initiation of the primer, low output performance of the primer, low ignition sensitivity of the gun propellant, and the effects of cold temperature. The goal was to determine the "best" of the Army and NASA test methods to assess the functional performance of primers and gun propellants. The approach was to evaluate the performance of both high-quality and deliberately defective primers to challenge the sensitivity of test methods. In addition, the ignition sensitivity of different manufacturing batches of gun propellants was evaluated. The results of the investigation revealed that improvements can be made in functional evaluations that can assist in identifying and reducing ignition and performance variations. The "best" functional evaluation of primers and propellant is achieved through a combination of both Army and NASA test methods. Incorporating the recommendations offered in this report may provide for considerable savings in reducing the number of cartridge firings, while significantly lowering the rejection rate of primer, propellant and cartridge lots. The most probable causes for ignition and combustion-related hangfires were the lack of calcium silicide in the primer mix, a low output performance of primers, and finally, poor ignition sensitivity of gun propellant. Cold temperatures further reduce propellant ignition sensitivity, as well as reducing burn rate and chamber pressures.

29 CFR 1926.154 - Temporary heating devices.

Code of Federal Regulations, 2010 CFR

2010-07-01

..., unless otherwise permitted by the manufacturer's markings. (d) Solid fuel salamanders. Solid fuel salamanders are prohibited in buildings and on scaffolds. (e) Oil-fired heaters. (1) Flammable liquid-fired...

29 CFR 1926.154 - Temporary heating devices.

Code of Federal Regulations, 2011 CFR

2011-07-01

..., unless otherwise permitted by the manufacturer's markings. (d) Solid fuel salamanders. Solid fuel salamanders are prohibited in buildings and on scaffolds. (e) Oil-fired heaters. (1) Flammable liquid-fired...

Testing of a Liquid Oxygen/Liquid Methane Reaction Control Thruster in a New Altitude Rocket Engine Test Facility

NASA Technical Reports Server (NTRS)

Meyer, Michael L.; Arrington, Lynn A.; Kleinhenz, Julie E.; Marshall, William M.

2012-01-01

A relocated rocket engine test facility, the Altitude Combustion Stand (ACS), was activated in 2009 at the NASA Glenn Research Center. This facility has the capability to test with a variety of propellants and up to a thrust level of 2000 lbf (8.9 kN) with precise measurement of propellant conditions, propellant flow rates, thrust and altitude conditions. These measurements enable accurate determination of a thruster and/or nozzle s altitude performance for both technology development and flight qualification purposes. In addition the facility was designed to enable efficient test operations to control costs for technology and advanced development projects. A liquid oxygen-liquid methane technology development test program was conducted in the ACS from the fall of 2009 to the fall of 2010. Three test phases were conducted investigating different operational modes and in addition, the project required the complexity of controlling propellant inlet temperatures over an extremely wide range. Despite the challenges of a unique propellant (liquid methane) and wide operating conditions, the facility performed well and delivered up to 24 hot fire tests in a single test day. The resulting data validated the feasibility of utilizing this propellant combination for future deep space applications.

Metallized solid rocket propellants based on AN/AP and PSAN/AP for access to space

NASA Astrophysics Data System (ADS)

Levi, S.; Signoriello, D.; Gabardi, A.; Molinari, M.; Galfetti, L.; Deluca, L. T.; Cianfanelli, S.; Klyakin, G. F.

2009-09-01

Solid rocket propellants based on dual mixes of inorganic crystalline oxidizers (ammonium nitrate (AN) and ammonium perchlorate (AP)) with binder and a mixture of micrometric-nanometric aluminum were investigated. Ammonium nitrate is a low-cost oxidizer, producing environment friendly combustion products but with lower specific impulse compared to AP. The better performance obtained with AP and the low quantity of toxic emissions obtained by using AN have suggested an interesting compromise based on a dual mixture of the two oxidizers. To improve the thermal response of raw AN, different types of phase stabilized AN (PSAN) and AN/AP co-crystals were investigated.

Thrust-isolating mounting. [characteristics of support for loads mounted in spacecraft

NASA Technical Reports Server (NTRS)

Wetzler, D. G. (Inventor)

1974-01-01

A supporting frame for a load, such as one or more telescopes, is isolated from all multi-gravitational forces, which will be developed within that load as that load is propelled into space, by using a shroud to fully and solidly hold that load until that load has been propelled into space. Thereafter, that shroud will be jettisoned; and then supports which are on, and which are movable with, that load will have surfaces thereon moved into supporting engagement with complementary surfaces on that supporting frame to enable that supporting frame and those supports to fully and solidly hold that load.

Azidated Ether-Butadiene-Ether Block Copolymers as Binders for Solid Propellants

NASA Astrophysics Data System (ADS)

Cappello, Miriam; Lamia, Pietro; Mura, Claudio; Polacco, Giovanni; Filippi, Sara

2016-07-01

Polymeric binders for solid propellants are usually based on hydroxyl-terminated polybutadiene (HTPB), which does not contribute to the overall energy output. Azidic polyethers represent an interesting alternative but may have poorer mechanical properties. Polybutadiene-polyether copolymers may combine the advantages of both. Four different ether-butadiene-ether triblock copolymers were prepared and azidated starting from halogenated and/or tosylated monomers using HTPB as initiator. The presence of the butadiene block complicates the azidation step and reduces the storage stability of the azidic polymer. Nevertheless, the procedure allows modifying the binder properties by varying the type and lengths of the energetic blocks.

Efficient solid rocket propulsion for access to space

NASA Astrophysics Data System (ADS)

Maggi, Filippo; Bandera, Alessio; Galfetti, Luciano; De Luca, Luigi T.; Jackson, Thomas L.

2010-06-01

Space launch activity is expected to grow in the next few years in order to follow the current trend of space exploitation for business purpose. Granting high specific thrust and volumetric specific impulse, and counting on decades of intense development, solid rocket propulsion is a good candidate for commercial access to space, even with common propellant formulations. Yet, some drawbacks such as low theoretical specific impulse, losses as well as safety issues, suggest more efficient propulsion systems, digging into the enhancement of consolidated techniques. Focusing the attention on delivered specific impulse, a consistent fraction of losses can be ascribed to the multiphase medium inside the nozzle which, in turn, is related to agglomeration; a reduction of agglomerate size is likely. The present paper proposes a model based on heterogeneity characterization capable of describing the agglomeration trend for a standard aluminized solid propellant formulation. Material microstructure is characterized through the use of two statistical descriptors (pair correlation function and near-contact particles) looking at the mean metal pocket size inside the bulk. Given the real formulation and density of a propellant, a packing code generates the material representative which is then statistically analyzed. Agglomerate predictions are successfully contrasted to experimental data at 5 bar for four different formulations.

Detonator-activated ball shutter

DOEpatents

McWilliams, R.A.; Holle, W.G. von.

1983-08-16

A detonator-activated ball shutter for closing an aperture in about 300[mu] seconds. The ball shutter containing an aperture through which light, etc., passes, is closed by firing a detonator which propels a projectile for rotating the ball shutter, thereby blocking passage through the aperture. 3 figs.

Space storable propellant performance program coaxial injector characterization

NASA Technical Reports Server (NTRS)

Burick, R. J.

1972-01-01

An experimental program was conducted to characterize the circular coaxial injector concept for application with the space-storable gas/liquid propellant combination FLOX(82.6% F2)/CH4(g) at high pressure. The primary goal of the program was to obtain high characteristic velocity efficiency in conjunction with acceptable injector/chamber compatibility. A series of subscale (single element) cold flow and hot fire experiments was employed to establish design criteria for a 3000-lbf (sea level) engine operating at 500 psia. The subscale experiments characterized both high performance core elements and peripheral elements with enhanced injector/chamber compatibility. The full-scale injector which evolved from the study demonstrated a performance level of 99 percent of the theoretical shifting characteristic exhaust velocity with low chamber heat flux levels. A 44-second-duration firing demonstrated the durability of the injector. Parametric data are presented that are applicable for the design of circular, coaxial injectors that operate with injection dynamics (fuel and oxidizer velocity, etc.) similar to those employed in the work reported.

Performance of a green propellant thruster with discharge plasma

NASA Astrophysics Data System (ADS)

Shindo, Takahiro; Wada, Asato; Maeda, Hiroshi; Watanabe, Hiroki; Takegahara, Haruki

2017-02-01

A discharge plasma was applied to initiate the combustion of a hydroxylammonium nitrate-based propellant as a substitute for the catalysts that are typically employed. The resulting thrust and thrust-to-power ratio during short interval firing tests as well as the chamber pressure with a single pulse discharge were evaluated. A 1.5-s firing test generated a maximum thrust of 322 mN along with a thrust-to-power ratio of 0.95 mN/W. During the single-pulse discharge trials, pulsed discharge capacitor energies of 5.4, 10.8, and 16.4 J were assessed, and the maximum chamber pressure was found to increase as the energy was raised. The maximum chamber pressures varied widely between experimental trials, and a 16.4-J energy value resulted in the highest chamber pressure of over 1 MPaG. The time spans between the pulsed discharge and the peak chamber pressure were in the range of 1-2 ms, representing a chamber pressure increase rate much higher than those obtained with standard catalysts.

Fuze for explosive magnetohydrodynamic generator

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

Webb, G.

1976-12-23

An apparatus is examined by which high explosive charges are propelled into and detonated at the center of an MHD-X generator. The high explosive charge units are engaged and propelled by a reciprocating ram device. Detonating in each instance is achieved by striking with a firing pin a detonator charge that is in register with a booster charge, the booster charge being in detonating communication with the high explosive charge. Various safety requirements are satisfied by a spring loaded slider operating in a channel transverse and adjacent to the booster charge. The slide retains the detonator charge out of registermore » with the booster charge until a safety pin that holds the slider in place is pulled by a lanyard attached between the reciprocating ram and the safety pin. Removal of the safety pin permits the detonator charge to slide into alignment with the booster charge. Firing pin actuation is initiated by the slider at the instant the detonator charge and the booster charge come into register.« less

Development of Life Prediction Capabilities for Liquid Propellant Rocket Engines. Task 4. Post-Fire Diagnostic System for the SSME System Architecture Study.

DTIC Science & Technology

1991-07-31

90 START MCC LN CAV PR 3 UNDERSHOOT ABOVE THRESHOLD YES MI A2-492 2/13/90 MAINSTAGE HPOT DS TMP CHANNEL A/B DIVERGENCE NO MI A2-492 2/13/90 MAINSTAGE ...System for the SSME System Architecture Study Y, , Contract NAS 3 -25883 JUL 31󈧣 CR-187112 Prepared for: National Aeronautics and Space...Liquid Propellant Rocket Engines Contract No. NAS 3 -25883 Eli Ki ,,, July 31, 1991 BY Dist Prepared By.: Mr. Mark Gage Aerojet Propulsion Division Box

Evacuation areas for transportation accidents involving propellant tank pressure bursts

NASA Technical Reports Server (NTRS)

Siewert, R. D.

1972-01-01

Evacuation areas are defined for those transportation accidents where volatile chemical propellant tanks are exposed to fire in the wreckage and eventually explode with consequent risks from fragments in surrounding populated areas. An evacuation area with a minimum radius of 600 m (2000 ft) is recommended to limit the statistical probability of fatality to one in 100 such accidents. The result was made possible by the derivation of a distribution function of distances reached by fragments from bursting chemical car tanks. Data concerning fragments was obtained from reports or tank car pressure bursts between 1958 and 1971.

KSC-06pd0096

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — NASA’s New Horizons spacecraft roars into the cloud-scattered sky trailing fire and smoke from the Atlas V rocket that propels it. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/Kim Shiflett

Development of sensing techniques for weaponry health monitoring

NASA Astrophysics Data System (ADS)

Edwards, Eugene; Ruffin, Paul B.; Walker, Ebonee A.; Brantley, Christina L.

2013-04-01

Due to the costliness of destructive evaluation methods for assessing the aging and shelf-life of missile and rocket components, the identification of nondestructive evaluation methods has become increasingly important to the Army. Verifying that there is a sufficient concentration of stabilizer is a dependable indicator that the missile's double-based solid propellant is viable. The research outlined in this paper summarizes the Army Aviation and Missile Research, Development, and Engineering Center's (AMRDEC's) comparative use of nanoporous membranes, carbon nanotubes, and optical spectroscopic configured sensing techniques for detecting degradation in rocket motor propellant. The first sensing technique utilizes a gas collecting chamber consisting of nanoporous structures that trap the smaller solid propellant particles for measurement by a gas analysis device. In collaboration with NASA-Ames, sensing methods are developed that utilize functionalized single-walled carbon nanotubes as the key sensing element. The optical spectroscopic sensing method is based on a unique light collecting optical fiber system designed to detect the concentration of the propellant stabilizer. Experimental setups, laboratory results, and overall effectiveness of each technique are presented in this paper. Expectations are for the three sensing mechanisms to provide nondestructive evaluation methods that will offer cost-savings and improved weaponry health monitoring.

Combustion mechanism of double-base propellant containing nitrogen heterocyclic nitroamines (II): The temperature distribution of the flame and its chemical structure

NASA Astrophysics Data System (ADS)

Yan, Qi-Long; Song, Zhen-Wei; Shi, Xiao-Bing; Yang, Zhi-Yuan; Zhang, Xiao-Hong

2009-03-01

In order to evaluate the actual pros and cons in the use of new nitroamines for solid rocket applications, the combustion properties of double-base propellants containing nitrogen heterocyclic nitroamines such as RDX, TNAD, HMX and DNP are investigated by means of high-speed photography technique, Non-contact wavelet-based measurement of flame temperature distribution. The chemical reactions in different combustion zone which control the burning characteristics of the double-base propellant containing nitrogen heterocyclic nitroamines were systematically investigated and descriptions of the detailed thermal decomposition mechanisms from solid phase to liquid phase or to gas phase are also included. It was indicated that the thermodynamic phase transition consisting of both evaporation and condensation of NC+NG, HMX, TNAD, RDX and DNP, are considered to provide a complete description of the mass transfer process in the combustion of these double-base propellants, and the combustion mechanisms of them are mainly involved with the oxidation mechanism of the NO 2, formaldehyde (CH 2O) and hydrogen cyanide (HCN). The entire oxidation reaction rate might be dependent on the pressure of the combustion chamber and temperature of the gas phase.

Snakes on a plane: modeling flexible active nematics

NASA Astrophysics Data System (ADS)

Selinger, Robin

Active soft matter systems of self-propelled rod-shaped particles exhibit ordered phases and collective behavior that are remarkably different from their passive analogs. In nature, many self-propelled rod-shaped particles, such as gliding bacteria and kinesin-driven microtubules, are flexible and can bend. We model these ``living liquid crystals'' to explore their phase behavior, dynamics, and pattern formation. We model particles as short polymers via molecular dynamics with a Langevin thermostat and various types of activity, substrate, and environments. For self-propelled polar particles gliding on a solid substrate, we map out the phase diagram as a function of particle density and flexibility. We compare simulated defect structures to those observed in colonies of gliding myxobacteria; compare spooling behavior to that observed in microtubule gliding assays; and analyze emergence of nematic and polar order. Next we explore pattern formation of self-propelled polar particles under flexible encapsulation, and on substrates with non-uniform Gaussian curvature. Lastly, we impose an activity mechanism that mimics extensile shear, study flexible particles both on solid substrates and coupled to a lipid membrane, and discuss comparisons to relevant experiments. Work performed in collaboration with Michael Varga (Kent State) and Luca Giomi (Universiteit Leiden.) Work supported by NSF DMR-1409658.

[Injury patterns and roentgen findings in gunshot wounds with rare flint ammunition].

PubMed

Pollak, S; Lindermann, A

1990-01-01

Smoothbore shotgun barrels can fire cartridges with common pellet loads as well as shotgun slugs and rubber bullets. Other than conventional shot, the cylindrical Brenneke-type rifled shotgun slugs sometimes cause perforating wounds. The shotgun ammunition for use in self-defence can have a single projectile or several rubber pellets. Where the propellant is black powder, short range shots will probably leave searing marks and intensive soot deposits. Fired at close range, rubber bullets can penetrate through the skin into the body, fired at greater distance they cause contusions. A case of homicide (repeated firing with a 12-ga. pump gun) is used to present and discuss the injury patterns and X-ray findings after impact of Brenneke-type slugs and rubber bullets as well as of "classical" shot pellets.

Development of life prediction capabilities for liquid propellant rocket engines. Post-fire diagnostic system for the SSME system architecture study

NASA Technical Reports Server (NTRS)

Gage, Mark; Dehoff, Ronald

1991-01-01

This system architecture task (1) analyzed the current process used to make an assessment of engine and component health after each test or flight firing of an SSME, (2) developed an approach and a specific set of objectives and requirements for automated diagnostics during post fire health assessment, and (3) listed and described the software applications required to implement this system. The diagnostic system described is a distributed system with a database management system to store diagnostic information and test data, a CAE package for visual data analysis and preparation of plots of hot-fire data, a set of procedural applications for routine anomaly detection, and an expert system for the advanced anomaly detection and evaluation.

Four-component numerical simulation model of radiative convective interactions in large-scale oxygen-hydrogen turbulent fire balls

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

Surzhikov, S.T.

1996-12-31

Two-dimensional radiative gas dynamics model for numerical simulation of oxygen-hydrogen fire ball which may be generated by an explosion of a launch vehicle with cryogenic (LO{sub 2}-LH{sub 2}) fuel components is presented. The following physical-chemical processes are taken into account in the numerical model: and effective chemical reaction between the gaseous components (O{sub 2}-H{sub 2}) of the propellant, turbulent mixing and diffusion of the components, and radiative heat transfer. The results of numerical investigations of the following problems are presented: The influence of radiative heat transfer on fire ball gas dynamics during the first 13 sec after explosion, the effectmore » of the fuel gaseous components afterburning on fire ball gas dynamics, and the effect of turbulence on fire ball gas dynamics (in a framework of algebraic model of turbulent mixing).« less

Investigation of the flow turning loss in unstable solid propellant rocket motors

NASA Astrophysics Data System (ADS)

Matta, Lawrence Mark

The goal of this study was to improve the understanding of the flow turning loss, which contributes to the damping of axial acoustic instabilities in solid propellant rocket motors. This understanding is needed to develop practical methods for designing motors that do not exhibit such instabilities. The flow turning loss results from the interaction of the flow of combustion products leaving the surface of the propellant with the acoustic field in an unstable motor. While state of the art solid rocket stability models generally account for the flow turning loss, its magnitude and characteristics have never been fully investigated. This thesis describes a combined theoretical, numerical, and experimental investigation of the flow turning loss and its dependence upon various motor design and operating parameters. First, a one dimensional acoustic stability equation that verifies the existence of the flow turning loss was derived for a chamber with constant mean pressure and temperature. The theoretical development was then extended to include the effects of mean temperature gradients to accommodate combustion systems in which mean temperature gradients and heat losses are significant. These analyses provided the background and expressions necessary to guide an experimental study. The relevant equations were then solved for the developed experimental setup to predict the behavior of the flow turning loss and the other terms of the developed acoustic stability equation. This was followed by and experimental study in which the flow turning region of an unstable solid propellant rocket motor was simulated. The setup was used, with and without combustion, to determine the dependence of the flow turning loss upon operating conditions. These studies showed that the flow turning loss strongly depends upon the gas velocity at the propellant surface and the location of the flow turning region relative to the standing acoustic wave. The flow turning loss measured in the experiment was found to be small relative to other mechanisms. This, however, was characteristic of the experimental setup and is not representative of actual rocket motors, in which the flow turning loss is often a significant part of the overall stability.

Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines

NASA Technical Reports Server (NTRS)

Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

1975-01-01

The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

Solid Hydrogen Experiments for Atomic Propellants: Particle Formation, Imaging, Observations, and Analyses

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2005-01-01

This report presents particle formation observations and detailed analyses of the images from experiments that were conducted on the formation of solid hydrogen particles in liquid helium. Hydrogen was frozen into particles in liquid helium, and observed with a video camera. The solid hydrogen particle sizes and the total mass of hydrogen particles were estimated. These newly analyzed data are from the test series held on February 28, 2001. Particle sizes from previous testing in 1999 and the testing in 2001 were similar. Though the 2001 testing created similar particles sizes, many new particle formation phenomena were observed: microparticles and delayed particle formation. These experiment image analyses are some of the first steps toward visually characterizing these particles, and they allow designers to understand what issues must be addressed in atomic propellant feed system designs for future aerospace vehicles.

Mars Flyer Rocket Propulsion Risk Assessment Kaiser Marquardt Testing

NASA Technical Reports Server (NTRS)

Marquardt, Kaiser

2001-01-01

This report describes the investigation of a 10-N, bipropellant thruster, operating at -40 C, with monomethylhydrazine (MMH) and 25% nitric oxide in nitrogen tetroxide (MON-25). The thruster testing was conducted as part of a risk reduction activity for the Mars Flyer, a proposed mission to fly a miniature airplane in the Martian atmosphere. Testing was conducted using an existing thruster, designed for MMH and MON-3 propellants. The nitric oxide content of MON-3 was increased to 25%, to lower its freezing point to -55 C. The thruster was conditioned, along with the propellants, to temperature prior to hot firing. Thruster operating parameters included oxidizer-to-fuel mixture ratios of 1.6 to 2.7 and inlet pressure ranging from 689 to 2070 kPa. The test matrix consisted of many 10-second firings and several 60-, 300-, 600-, and 1200-second firings, as well as pulse testing. The thruster successfully accumulated nearly 10,000 seconds of operation without failure, at temperatures ranging from -40 C to 22 C. At nominal inlet pressures, the ignition delay was comparable to MMH/MON-3 operation. The optimal performance for the 8.9-N thruster was determined to be at a mixture ratio of 1.93 with an average specific impulse of 298 sec.

Solid-liquid staged combustion space boosters

NASA Technical Reports Server (NTRS)

Culver, D. W.

1990-01-01

NASA has begun to evaluate solid-liquid hybrid propulsion for launch vehicle booster. A three-phase program was outlined to identify, acquire, and demonstrate technology needed to approximate solid and liquid propulsion state of the art. Aerojet has completed a Phase 1 study and recommends a solid-liquid staged combustion concept in which turbopump fed LO2 is burned with fuel-rich solid propellant effluent in aft-mounted thrust chambers.These reasonably sized thrust chambers are LO2 regeneratively cooled, supplemented with fuel-rich barrier cooling. Turbopumps are driven by the resulting GO2 coolant in an expander-bleed-burnoff cycle. Turbine exhaust pressurizes the LO2 tankage directly, and the excess is bled into supersonic nozzle splitlines, where it combusts with the fuel rich boundary layer. Thrust vector control is enhanced by supersonic nozzle movement on flexseal mounts. Every hybrid solid-liquid concept examined improves booster energy management and launch propellant safety compared to current solid boosters. Solid-liquid staged combustion improves hybrid performance by improving both combustion efficiency and combustion stability, especially important for large boosters. These improvements result from careful fluid management and use of smaller combustors. The study shows NASA safety, reliability, cost, and performance criteria are best met with this concept, wherein simple hardware relies on several separate emerging technologies, all of which have been demonstrated successfully.

On the combustion mechanisms of ZrH2 in double-base propellant.

PubMed

Yang, Yanjing; Zhao, Fengqi; Yuan, Zhifeng; Wang, Ying; An, Ting; Chen, Xueli; Xuan, Chunlei; Zhang, Jiankan

2017-12-13

Metal hydrides are regarded as a series of promising hydrogen-supplying fuel for solid rocket propellants. Their effects on the energetic and combustion performances of propellants are closely related to their reaction mechanisms. Here we report a first attempt to determine the reaction mechanism of ZrH 2 , a high-density metal hydride, in the combustion of a double-base propellant to evaluate its potential as a fuel. ZrH 2 is determined to possess good resistance to oxidation by nitrocellulose and nitroglycerine. Thus its combustion starts with dehydrogenation to generate H 2 and metallic Zr. Subsequently, the newly formed Zr and H 2 participate in the combustion and, especially, Zr melts and then combusts on the burning surface which favors the heat feedback to the propellant. This phenomenon is completely different from the combustion behavior of the traditional fuel Al, where the Al particles are ejected off the burning surface of the propellant to get into the luminous flame zone to burn. The findings in this work validate the potential of ZrH 2 as a hydrogen-supplying fuel for double-base propellants.

High-Fidelity Microstructural Characterization and Performance Modeling of Aluminized Composite Propellant

DOE PAGES

Kosiba, Graham D.; Wixom, Ryan R.; Oehlschlaeger, Matthew A.

2017-10-27

Image processing and stereological techniques were used to characterize the heterogeneity of composite propellant and inform a predictive burn rate model. Composite propellant samples made up of ammonium perchlorate (AP), hydroxyl-terminated polybutadiene (HTPB), and aluminum (Al) were faced with an ion mill and imaged with a scanning electron microscope (SEM) and x-ray tomography (micro-CT). Properties of both the bulk and individual components of the composite propellant were determined from a variety of image processing tools. An algebraic model, based on the improved Beckstead-Derr-Price model developed by Cohen and Strand, was used to predict the steady-state burning of the aluminized compositemore » propellant. In the presented model the presence of aluminum particles within the propellant was introduced. The thermal effects of aluminum particles are accounted for at the solid-gas propellant surface interface and aluminum combustion is considered in the gas phase using a single global reaction. In conclusion, properties derived from image processing were used directly as model inputs, leading to a sample-specific predictive combustion model.« less

High-Fidelity Microstructural Characterization and Performance Modeling of Aluminized Composite Propellant

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

Kosiba, Graham D.; Wixom, Ryan R.; Oehlschlaeger, Matthew A.

Image processing and stereological techniques were used to characterize the heterogeneity of composite propellant and inform a predictive burn rate model. Composite propellant samples made up of ammonium perchlorate (AP), hydroxyl-terminated polybutadiene (HTPB), and aluminum (Al) were faced with an ion mill and imaged with a scanning electron microscope (SEM) and x-ray tomography (micro-CT). Properties of both the bulk and individual components of the composite propellant were determined from a variety of image processing tools. An algebraic model, based on the improved Beckstead-Derr-Price model developed by Cohen and Strand, was used to predict the steady-state burning of the aluminized compositemore » propellant. In the presented model the presence of aluminum particles within the propellant was introduced. The thermal effects of aluminum particles are accounted for at the solid-gas propellant surface interface and aluminum combustion is considered in the gas phase using a single global reaction. In conclusion, properties derived from image processing were used directly as model inputs, leading to a sample-specific predictive combustion model.« less

Transient processes in the combustion of nitramine propellants

NASA Technical Reports Server (NTRS)

Cohen, N. S.; Strand, L. D.

1978-01-01

A transient combustion model of nitramine propellants is combined with an isentropic compression shock formation model to determine the role of nitramine propellant combustion in DDT, excluding effects associated with propellant structural properties or mechanical behavior. The model is derived to represent the closed pipe experiment that is widely used to characterize explosives, except that the combustible material is a monolithic charge rather than compressed powder. Computations reveal that the transient combustion process cannot by itself produce DDT by this model. Compressibility of the solid at high pressure is the key factor limiting pressure buildups created by the combustion. On the other hand, combustion mechanisms which promote pressure buildups are identified and related to propellant formulation variables. Additional combustion instability data for nitramine propellants are presented. Although measured combustion response continues to be low, more data are required to distinguish HMX and active binder component contributions. A design for a closed vessel apparatus for experimental studies of high pressure combustion is discussed.

40 CFR 266.202 - Definition of solid waste.

Code of Federal Regulations, 2013 CFR

2013-07-01

... MANAGEMENT FACILITIES Military Munitions § 266.202 Definition of solid waste. (a) A military munition is not... personnel or explosives and munitions emergency response specialists (including training in proper destruction of unused propellant or other munitions); or (ii) Use in research, development, testing, and...

40 CFR 266.202 - Definition of solid waste.

Code of Federal Regulations, 2012 CFR

2012-07-01

... MANAGEMENT FACILITIES Military Munitions § 266.202 Definition of solid waste. (a) A military munition is not... personnel or explosives and munitions emergency response specialists (including training in proper destruction of unused propellant or other munitions); or (ii) Use in research, development, testing, and...

40 CFR 266.202 - Definition of solid waste.

Code of Federal Regulations, 2014 CFR

2014-07-01

... MANAGEMENT FACILITIES Military Munitions § 266.202 Definition of solid waste. (a) A military munition is not... personnel or explosives and munitions emergency response specialists (including training in proper destruction of unused propellant or other munitions); or (ii) Use in research, development, testing, and...

Tradespace Exploration of Distributed Propulsors for Advanced On-Demand Mobility Concepts

NASA Technical Reports Server (NTRS)

Borer, Nicholas K.; Moore, Mark D.; Turnbull, Andrew R.

2014-01-01

Combustion-based sources of shaft power tend to significantly penalize distributed propulsion concepts, but electric motors represent an opportunity to advance the use of integrated distributed propulsion on an aircraft. This enables use of propellers in nontraditional, non-thrust-centric applications, including wing lift augmentation, through propeller slipstream acceleration from distributed leading edge propellers, as well as wingtip cruise propulsors. Developing propellers for these applications challenges long-held constraints within propeller design, such as the notion of optimizing for maximum propulsive efficiency, or the use of constant-speed propellers for high-performance aircraft. This paper explores the design space of fixed-pitch propellers for use as (1) lift augmentation when distributed about a wing's leading edge, and (2) as fixed-pitch cruise propellers with significant thrust at reduced tip speeds for takeoff. A methodology is developed for evaluating the high-level trades for these types of propellers and is applied to the exploration of a NASA Distributed Electric Propulsion concept. The results show that the leading edge propellers have very high solidity and pitch well outside of the empirical database, and that the cruise propellers can be operated over a wide RPM range to ensure that thrust can still be produced at takeoff without the need for a pitch change mechanism. To minimize noise exposure to observers on the ground, both the leading edge and cruise propellers are designed for low tip-speed operation during takeoff, climb, and approach.

Hybrid rocket propellants from lunar material

NASA Astrophysics Data System (ADS)

Sparks, Douglas R.

This paper examines the use of lunar material for hybrid rocket propellants. Liquid oxygen is identified as the primary oxidizer and metals such as aluminum, magnesium, calcium, titanium and silicon are compared as possible fuels. Due to the reduced transportation costs, the use of lunar materials for both oxidizer and fuel will dramatically reduce the cost of a sustained space program. The advantage of hybrid rocket systems over liquid and solid rockets is discussed. It is pointed out that this type of hybrid rocket propellant could also be obtained from asteroidal and planetary soils, thereby facilitating the exploration and industrialization of the inner solar system.

Combustion Mechanisms of Solids

DTIC Science & Technology

1992-02-24

ELEMENT NO. NO NO ACCESSION NO Arlington, VA 22217-5000 11 TITLE (include Security Classification) COMBUSTION MECHANISMS OF SOLIDS 12. PERSONAL AUTHOR(S...FIELD GROUP I SUB-GROUP COMBUSTION , SOLID PROPELLANT 19 ABSTRACT (Continue on reverse if necessary and identify by block number) This report...ingredients tested (AP, AN, PBAN, NMMO and BAMO-THF). Ingredient combustion behavior was studied by the edge burning sandwich method using sandwiches

Specific Impulses Losses in Solid Propellant Rockets

DTIC Science & Technology

1974-12-17

binder -- polyvinyl, polyurethane, or polybutadiene) markedly increases performance. Aluminum is the most widely used metal since its energy properties...temperature is also used. -5- The specific impulse values calculated for a typical propellant with 16.4% aluminum are as follows: (p0 70 atm. p - 1 atm...Direct Measurement of Combuction Efficiency of Aluminum Analysis of the condensed phase enables the proportion of unburnt aluminum to be determined

Peculiar Traits of Coarse AP

DTIC Science & Technology

2014-01-01

propellant. Since coarse AP in particles larger than about 150 microns are used in great majority for AP oxidized solid propellants, the nature of...Microscopic amounts of liquid containing water were contained in the reactive centers. The maximum size for reactive centers was reasoned to be...bond in the original chlorate ion. Oxygen atom swapping between chlorate and perchlorate ions would provide chlorate migration without use of forces

Problem Definition Study: Lead Beta-Resorcylate

DTIC Science & Technology

1979-02-01

unless so desig- nated by other authorized documents. -3- »SUMMARY Lead ß-resorcylate is used as a burning rate modifier in solid propel- lant... sediment and biota 3. Acute mammalian toxicity study 4. Chronic mammalian toxicity study 5. Determine the effectiveness of proposed treatment...burning rate moderator in solvent and solventless double base propellents. This salt enters the environment in the wastewater generated during the

Self-Propelled Hovercraft Based on Cold Leidenfrost Phenomenon

PubMed Central

Shi, Meng; Ji, Xing; Feng, Shangsheng; Yang, Qingzhen; Lu, Tian Jian; Xu, Feng

2016-01-01

The Leidenfrost phenomenon of liquid droplets levitating and dancing when placed upon a hot plate due to propulsion of evaporative vapor has been extended to many self-propelled circumstances. However, such self-propelled Leidenfrost devices commonly need a high temperature for evaporation and a structured solid substrate for directional movements. Here we observed a “cold Leidenfrost phenomenon” when placing a dry ice device on the surface of room temperature water, based on which we developed a controllable self-propelled dry ice hovercraft. Due to the sublimated vapor, the hovercraft could float on water and move in a programmable manner through designed structures. As demonstrations, we showed that the hovercraft could be used as a cargo ship or a petroleum contamination collector without consuming external power. This phenomenon enables a novel way to utilize programmable self-propelled devices on top of room temperature water, holding great potential for applications in energy, chemical engineering and biology. PMID:27338595

Self-Propelled Hovercraft Based on Cold Leidenfrost Phenomenon.

PubMed

Shi, Meng; Ji, Xing; Feng, Shangsheng; Yang, Qingzhen; Lu, Tian Jian; Xu, Feng

2016-06-24

The Leidenfrost phenomenon of liquid droplets levitating and dancing when placed upon a hot plate due to propulsion of evaporative vapor has been extended to many self-propelled circumstances. However, such self-propelled Leidenfrost devices commonly need a high temperature for evaporation and a structured solid substrate for directional movements. Here we observed a "cold Leidenfrost phenomenon" when placing a dry ice device on the surface of room temperature water, based on which we developed a controllable self-propelled dry ice hovercraft. Due to the sublimated vapor, the hovercraft could float on water and move in a programmable manner through designed structures. As demonstrations, we showed that the hovercraft could be used as a cargo ship or a petroleum contamination collector without consuming external power. This phenomenon enables a novel way to utilize programmable self-propelled devices on top of room temperature water, holding great potential for applications in energy, chemical engineering and biology.

The Effect of Propellant Variables on Slag in Subscale Spin Motors. Part 1; Design and Qualification of a Slag Discrimination Motor

NASA Technical Reports Server (NTRS)

Perkins, F. M.; Beus, R. W.; May, D. H.

1995-01-01

The formation, collection, and expulsion of aluminum oxide slag is known to affect the performance of many solid rocket motor systems. Slag expulsion, in particular, is believed to be capable of causing pressure and thrust perturbations. Propellant combustion studies, performed and documented by many investigators, have shown that variations in propellant raw materials and processing affect the nature of alumina droplets at the burning propellant surface, and hence, may affect the quantity of slag retained in the motor chamber, available for expulsion. Thiokol has completed an experimental and analytical evaluation to determine the effects of several material and process variables on Space SHuttle propellant and its propensity to 'slag'. This paper describes the test article, a small scale spin motor with special nozzle, designed and qualified as a slag discriminating tool for use in the evaluation.

Modeling and simulation of the debonding process of composite solid propellants

NASA Astrophysics Data System (ADS)

Feng, Tao; Xu, Jin-sheng; Han, Long; Chen, Xiong

2017-07-01

In order to study the damage evolution law of composite solid propellants, the molecular dynamics particle filled algorithm was used to establish the mesoscopic structure model of HTPB(Hydroxyl-terminated polybutadiene) propellants. The cohesive element method was employed for the adhesion interface between AP(Ammonium perchlorate) particle and HTPB matrix and the bilinear cohesive zone model was used to describe the mechanical response of the interface elements. The inversion analysis method based on Hooke-Jeeves optimization algorithm was employed to identify the parameters of cohesive zone model(CZM) of the particle/binder interface. Then, the optimized parameters were applied to the commercial finite element software ABAQUS to simulate the damage evolution process for AP particle and HTPB matrix, including the initiation, development, gathering and macroscopic crack. Finally, the stress-strain simulation curve was compared with the experiment curves. The result shows that the bilinear cohesive zone model can accurately describe the debonding and fracture process between the AP particles and HTPB matrix under the uniaxial tension loading.

Characterization of Air Emissions from Open Burning and Open Detonation of Gun Propellants and Ammunition

EPA Science Inventory

Emissions from open burning (OB) and open detonation (OD) of military ordnance and static fires (SF) of rocket motors were sampled in fall, 2013 at the Dundurn Depot (Saskatchewan, Canada). Emission sampling was conducted with an aerostat-lofted instrument package termed the “Fl...

Early Rockets

NASA Image and Video Library

1950-01-01

Test firing of a Redstone Missile at Redstone Test Stand in the early 1950's. The Redstone was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the von Braun Team under the management of the U.S. Army. The Redstone was the first major rocket development program in the United States.

46 CFR 28.160 - Portable fire extinguishers.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 2,500 square feet (269.1 sq. meters) or fraction thereof suitable for hazards involved. Paint... square feet (269.1 sq. meters) or fraction thereof located in the vicinity of exits, either inside or... spaces; Internal combustion propelling machinery B-II 1 for each 1,000 brake horsepower or fraction...

Post Accident Procedures for Chemicals and Propellants.

DTIC Science & Technology

1982-09-01

benzoyl derivatives, saccharin, medicines, dyes and perfumes ; as a source of toluene diiscyanates (polyurethane resins); in explosives (TNT...menasi ene piC hh, etri Sin io. pr-- -1 ra inqapparatus Fire Ctiiifnltnelnestenlndlesn 617 9gnili Tsnsp-eeteee 914*F q.I Ni..th ’It iha. A.. dii ~e iha

Full field study of strain distribution near the crack tip in the fracture of solid propellants via large strain digital image correlation and optical microscopy

NASA Astrophysics Data System (ADS)

Gonzalez, Javier

A full field method for visualizing deformation around the crack tip in a fracture process with large strains is developed. A digital image correlation program (DIC) is used to incrementally compute strains and displacements between two consecutive images of a deformation process. Values of strain and displacements for consecutive deformations are added, this way solving convergence problems in the DIC algorithm when large deformations are investigated. The method developed is used to investigate the strain distribution within 1 mm of the crack tip in a particulate composite solid (propellant) using microscopic visualization of the deformation process.

Aerodynamic characteristics of the Grumman H-33 orbiter mated to a three segment solid propellant booster

NASA Technical Reports Server (NTRS)

Sims, F.; Olive, R.

1971-01-01

Experimental aerodynamic investigations were conducted on a .003366-scale model of the Grumman space shuttle configuration mounted to a three (3) segmented solid propellant booster. These tests were conducted in the MSFC 14-inch trisonic wind tunnel over a Mach number range of 0.6 to 4.96. The purpose of the test was to determine the aerodynamic characteristics of this configuration. Aerodynamic data was taken over a nominal angle of attack and angle of sideslip of -10 degrees to 10 degrees at zero degrees beta and alpha respectively. In addition, data was obtained for the H-33 orbiter alone to supplement data from TWT 502 and TWT 503.

Irradiated ignition over solid materials in reduce pressure environment: Fire safety issue in man-made enclosure system

NASA Astrophysics Data System (ADS)

Nakamura, N.; Aoki, A.

Effects of ambient pressure and oxygen yield on irradiated ignition characteristics over solid combustibles have been studied experimentally Aim of the present study is to elucidate the flammability and chance of fire in depressurized enclosure system and give ideas for the fire safety and fire fighting strategies in such environment Thin cellulosic paper is considered as the solid combustible since cellulose is one of major organic compounds and flammables in the nature Applied atmosphere consists of inert gas either CO2 or N2 and oxygen and various mixture ratios are of concerned Total ambient pressure level is varied from 0 1MPa standard atmospheric pressure to 0 02MPa Ignition is initiated by external thermal flux exposed into the solid surface as a model of unexpected thermal input to initiate the localized fire Thermal degradation of the solid induces combustible gaseous products e g CO H2 or other low class of HCs and the gas mixes with ambient oxygen to form the combustible mixture over the solid Heat transfer from the hot irradiated surface into the mixture accelerates the local exothermic reaction in the gas phase and finally thermal runaway ignition is achieved Ignition event is recorded by high-speed digital video camera to analyze the ignition characteristics Flammable map in partial pressure of oxygen Pox and total ambient pressure Pt plane is made to reveal the fire hazard in depressurized environment Results show that wider flammable range is obtained depending on the imposed ambient

Unsteady combustion of solid propellants

NASA Astrophysics Data System (ADS)

Chung, T. J.; Kim, P. K.

The oscillatory motions of all field variables (pressure, temperature, velocity, density, and fuel fractions) in the flame zone of solid propellant rocket motors are calculated using the finite element method. The Arrhenius law with a single step forward chemical reaction is used. Effects of radiative heat transfer, impressed arbitrary acoustic wave incidence, and idealized mean flow velocities are also investigated. Boundary conditions are derived at the solid-gas interfaces and at the flame edges which are implemented via Lagrange multipliers. Perturbation expansions of all governing conservation equations up to and including the second order are carried out so that nonlinear oscillations may be accommodated. All excited frequencies are calculated by means of eigenvalue analyses, and the combustion response functions corresponding to these frequencies are determined. It is shown that the use of isoparametric finite elements, Gaussian quadrature integration, and the Lagrange multiplier boundary matrix scheme offers a convenient approach to two-dimensional calculations.

Gun Propellant Residues Dispersed from Static Artillery Firings of LG1 Mark 2 and C3 105-mm Howitzers

DTIC Science & Technology

2008-10-01

nitroglycerine (NG). In this study, aluminium witness plates were placed in front of the muzzle of the gun to collect residues propelled in the environment...témoins en aluminium ont été placées en face de la bouche du canon pour récolter les résidus de tirs propulsés dans l’environnement. Les plaques ont été...the Mark II and C3 105-mm howitzers, during an artillery exercise performed from May 9 to May 12, 2005 at CFB Gagetown in New-Brunswick. Aluminium

Historical perspective - Viking Mars Lander propulsion

NASA Technical Reports Server (NTRS)

Morrisey, Donald C.

1989-01-01

This paper discusses the Viking 1 and 2 missions to Mars in 1975-1976 and describes the design evolution of the Viking Terminal Descent Rocket Engines responsible for decelerating the Viking Mars Landers during the final portion of their descent from orbit. The Viking Terminal Descent Rocket Engines have twice the thrust of the largest monopropellant hydrazine engine developed previously but weigh considerably less. The engine has 18 nozzles, the capability of 10:1 throttling, is totally sealed until fired, employs no organic unsealed materials, is 100 percent germ free, utilized hydrazine STM-20 as the propellant, and starts at a temperature more than 45 F below the propellant's freezing point.

Calculated maximum Hl ground-level concentrations downwind from launch pad aborts of the space shuttle and Titan 3 C vehicles at Kennedy Space Center

NASA Technical Reports Server (NTRS)

Dumbauld, R. K.; Bjorklund, J. R.

1972-01-01

A quantitative assessment is described of the potential environmental hazard posed by the atmospheric release of HCl resulting from the burning of solid propellant during two hypothetical on-pad aborts of the Titan 3 C and space shuttle vehicles at Kennedy Space Center. In one pad-abort situation, it is assumed that the cases of the two solid-propellant engines are ruptured and the burning propellant falls to the ground in the immediate vicinity of the launch pad where it continues to burn for 5 minutes. In the other pad-abort situation considered, one of the two solid engines on each vehicle is assumed to ignite and burn at the normal rate while the vehicle remains on the launch pad. Calculations of maximum HCl ground-level concentration for the above on-pad abort situations were made using the computerized NASA/MSFC multilayer diffusion models in conjunction with appropriate meteorological and source inputs. Three meteorological regimes are considered-fall, spring, and afternoon sea-breeze. Source inputs for the hazard calculations were developed. The principal result of the calculations is that maximum ground-level HCl concentrations at distances greater than 1 kilometer from the launch pad are less than 3 parts per million in all cases considered.

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.

Breadboard RL10-2B low-thrust operating mode (second iteration) test report

NASA Technical Reports Server (NTRS)

Kanic, Paul G.; Kaldor, Raymond B.; Watkins, Pia M.

1988-01-01

Cryogenic rocket engines requiring a cooling process to thermally condition the engine to operating temperature can be made more efficient if cooling propellants can be burned. Tank head idle and pumped idle modes can be used to burn propellants employed for cooling, thereby providing useful thrust. Such idle modes required the use of a heat exchanger to vaporize oxygen prior to injection into the combustion chamber. During December 1988, Pratt and Whitney conducted a series of engine hot firing demonstrating the operation of two new, previously untested oxidizer heat exchanger designs. The program was a second iteration of previous low thrust testing conducted in 1984, during which a first-generation heat exchanger design was used. Although operation was demonstrated at tank head idle and pumped idle, the engine experienced instability when propellants could not be supplied to the heat exchanger at design conditions.

Iridium-coated rhenium thrusters by CVD

NASA Technical Reports Server (NTRS)

Harding, J. T.; Kazaroff, J. M.; Appel, M. A.

1989-01-01

Operation of spacecraft thrusters at increased temperature reduces propellant requirements. Inasmuch as propellant comprises the bulk of a satellite's mass, even a small percentage reduction makes possible a significant enhancement of the mission in terms of increased payload. Because of its excellent high temperature strength, rhenium is often the structural material of choice. It can be fabricated into free-standing shapes by chemical vapor deposition (CVD) onto an expendable mandrel. What rhenium lacks is oxidation resistance, but this can be provided by a coating of iridium, also by CVD. This paper describes the process used by Ultramet to fabricate 22-N (5-lbf) and, more recently, 445-N (100-lbf) Ir/Re thrusters; characterizes the CVD-deposited materials; and summarizes the materials effects of firing these thrusters. Optimal propellant mixture ratios can be employed because the materials withstand an oxidizing environment up to the melting temperature of iridium, 2400 C (4350 F).

Iridium-coated rhenium thrusters by CVD

NASA Technical Reports Server (NTRS)

Harding, John T.; Kazaroff, John M.; Appel, Marshall A.

1988-01-01

Operation of spacecraft thrusters at increased temperature reduces propellant requirements. Inasmuch as propellant comprises the bulk of a satellite's mass, even a small percentage reduction makes possible a significant enhancement of the mission in terms of increased payload. Because of its excellent high temperature strength, rhenium is often the structural material of choice. It can be fabricated into free-standing shapes by chemical vapor deposition (CVD) onto an expendable mandrel. What rhenium lacks is oxidation resistance, but this can be provided by a coating of iridium, also by CVD. This paper describes the process used by Ultramet to fabricate 22-N (5-lbf) and, more recently, 445-N (100-lbf) Ir/Re thrusters; characterizes the CVD-deposited materials; and summarizes the materials effects of firing these thrusters. Optimal propellant mixture ratios can be employed because the materials withstand an oxidizing environment up to the meltimg temperature of iridium, 2400 C (4350 F).

Study of solid rocket motors for a space shuttle booster. Volume 4: Mass properties report

NASA Technical Reports Server (NTRS)

Vonderesch, A. H.

1972-01-01

Mass properties data for the 156 inch diameter, parallel burn, solid propellant rocket engine for the space shuttle booster are presented. Design ground rules and assumptions applicable to generation of the mass properties data are described, together with pertinent data sources.

Combustion Mechanisms of Solids

DTIC Science & Technology

1986-08-01

PYROLYSIS OF POLYMERIC SOLID PROPELLANT BINDERS Approved Edward W. Price, Chairman Gary A. randro Jech e I.dJagoda Robert K. SVigman Date Approved by...committee, Drs. I. A. Jagoda, F. L. Cook, G. A. Flandro , R. K. Sigman, I am indebted for their many hours of discussion and their suggestions. Special

Design and performance evaluations of a LO2/methane reaction control engine

NASA Astrophysics Data System (ADS)

Johnson, Aaron

Liquid oxygen (LOX) and liquid methane (LCH4) are a propellant combination viewed as a potential enabling technology for spacecraft propulsion. Reasons why LOX/LCH4 is being used as an alternative propellant source include: it is less toxic than other propellants, it has the possibility to be harvested on extraterrestrial soil, LCH4 has a higher energy density than liquid hydrogen (LH2; commonly used on vehicle main engines), and LOX/LCH4 has comparable performance to other well-known propellant combinations. Through the continued partnership between the National Aeronautics and Space Administration (NASA) and the University of Texas at El Paso (UTEP) a LOX/LCH4 reaction control engine (RCE) was developed and researched. The RCE was developed for the purpose of being integrated into two UTEP LOX/LCH4 vehicles, Janus and Daedalus, and was designed based on previous engines tested both at NASA and the center for space exploration and technology research (cSETR) lab. This report details the design process and manufacturing of the engine, cold flow studies evaluating injector design, and preliminary hot fire tests to give insight into engine performance.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

Testing of Wrought Iridium/Chemical Vapor Deposition Rhenium Rocket

NASA Technical Reports Server (NTRS)

Reed, Brian D.; Schneider, Steven J.

1996-01-01

A 22-N class, iridium/rhenium (Ir/Re) rocket chamber, composed of a thick (418 miocrometer) wrought iridium (Ir) liner and a rhenium substrate deposited via chemical vapor deposition, was tested over an extended period on gaseous oxygen/gaseous hydrogen (GO2/GH2) propellants. The test conditions were designed to produce species concentrations similar to those expected in an Earth-storable propellant combustion environment. Temperatures attained in testing were significantly higher than those expected with Earth-storable propellants, both because of the inherently higher combustion temperature of GO2/GH2 propellants and because the exterior surface of the rocket was not treated with a high-emissivity coating that would be applied to flight class rockets. Thus the test conditions were thought to represent a more severe case than for typical operational applications. The chamber successfully completed testing (over 11 hr accumulated in 44 firings), and post-test inspections showed little degradation of the Ir liner. The results indicate that use of a thick, wrought Ir liner is a viable alternative to the Ir coatings currently used for Ir/Re rockets.

Study of Spray Disintegration in Accelerating Flow Fields

NASA Technical Reports Server (NTRS)

Nurick, W. H.

1972-01-01

An analytical and experimental investigation was conducted to perform "proof of principlem experiments to establish the effects of propellant combustion gas velocity on propella'nt atomization characteristics. The propellants were gaseous oxygen (GOX) and Shell Wax 270. The fuel was thus the same fluid used in earlier primary cold-flow atomization studies using the frozen wax method. Experiments were conducted over a range in L* (30 to 160 inches) at two contraction ratios (2 and 6). Characteristic exhaust velocity (c*) efficiencies varied from SO to 90 percent. The hot fire experimental performance characteristics at a contraction ratio of 6.0 in conjunction with analytical predictions from the drovlet heat-up version of the Distributed Energy Release (DER) combustion computer proDam showed that the apparent initial dropsize compared well with cold-flow predictions (if adjusted for the gas velocity effects). The results also compared very well with the trend in perfomnce as predicted with the model. significant propellant wall impingement at the contraction ratio of 2.0 precluded complete evaluation of the effect of gross changes in combustion gas velocity on spray dropsize.

Development and Testing of a Novel Green Propellant Piston Tank

NASA Technical Reports Server (NTRS)

Diaz, C. E.; Cavender, D. P.; Higdon, K.; Abrams, J.; Duchek, M. E.; Mader, H.

2017-01-01

Analytical Mechanics Associates (AMA), in cooperation with NASA Marshall Space Flight Center's (MSFC's) Spacecraft Propulsion Systems Branch, developed and tested a novel propellant tank design that employs an internal piston pressurized with an inert gas to expel propellant to thrusters. During the course of this activity, AMA designed, oversaw fabrication, and delivered to MSFC for testing, a piston propellant tank sized for 3U or larger CubeSats. MSFC conducted liquid expulsion testing using ethylene glycol as a referee fluid to map the tank's performance at different pressures and piston positions. Following the expulsion test campaign, the tank is planned to be integrated into a propulsion system test bed for hot fire tests with a 100mN monopropellant thruster to evaluate the tank's influence on thruster performance when operated in a flight like manner. Described in this paper is a comprehensive summary of how the tanks were designed, built, and tested. The fundamental knowledge gained through the fabrication and testing of these tanks gives evidence that the piston tank design may be scalable to meet the requirements and constraints of other small satellites.

40 CFR 63.865 - Performance test requirements and test methods.

Code of Federal Regulations, 2011 CFR

2011-07-01

... the chemical recovery system at the kraft or soda pulp mill, kg/Mg (lb/ton) of black liquor solids... the performance test, megagrams per day (Mg/d) (tons per day (ton/d)) of black liquor solids fired. ER1ref, SDT = reference emission rate of 0.10 kg/Mg (0.20 lb/ton) of black liquor solids fired for...

40 CFR 63.865 - Performance test requirements and test methods.

Code of Federal Regulations, 2010 CFR

2010-07-01

... the chemical recovery system at the kraft or soda pulp mill, kg/Mg (lb/ton) of black liquor solids... the performance test, megagrams per day (Mg/d) (tons per day (ton/d)) of black liquor solids fired. ER1ref, SDT = reference emission rate of 0.10 kg/Mg (0.20 lb/ton) of black liquor solids fired for...

RSRM-13 (360Q013) ballistics mass properties flight designation STS-41

NASA Technical Reports Server (NTRS)

Laubacher, Brian A.; Richards, M. C.

1990-01-01

The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-13 motors which were assigned to the STS-41 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information are also presented. The presented ballistic performance was based on the Operational Flight Instrumentation. The adjustments made to the raw data on this flight include biasing the data to correct ambient pressure before liftoff. The performance from each motor as well as matched pair performance values were well within the CEI Specification requirements.

Detection of iron atoms by emission spectroscopy and laser-induced fluorescence in solid propellant flames.

PubMed

Vilmart, G; Dorval, N; Orain, M; Lambert, D; Devillers, R; Fabignon, Y; Attal-Tretout, B; Bresson, A

2018-05-10

Planar laser-induced fluorescence on atomic iron is investigated in this paper, and a measurement strategy is proposed to monitor the fluorescence of iron atoms with good sensitivity. A model is proposed to fit the experimental fluorescence spectra, and good agreement is found between simulated and experimental spectra. Emission and laser-induced fluorescence measurements are performed in the flames of ammonium perchlorate composite propellants containing iron-based catalysts. A fluorescence signal from iron atoms after excitation at 248 nm is observed for the first time in propellant flames. Images of the spatial distribution of iron atoms are recorded in the flame in which turbulent structures are generated. Iron fluorescence is detected up to 1.0 MPa, which opens the way to application in propellant combustion.

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

A new blade element method for calculating the performance of high and intermediate solidity axial flow fans

NASA Technical Reports Server (NTRS)

Borst, H. V.

1978-01-01

A method is presented to design and predict the performance of axial flow rotors operating in a duct. The same method is suitable for the design of ducted fans and open propellers. The unified method is based on the blade element approach and the vortex theory for determining the three dimensional effects, so that two dimensional airfoil data can be used for determining the resultant force on each blade element. Resolution of this force in the thrust and torque planes and integration allows the total performance of the rotor, fan or propeller to be predicted. Three different methods of analysis, one based on a momentum flow theory; another on the vortex theory of propellers; and a third based on the theory of ducted fans, agree and reduce cascade airfoil data to single line as a function of the loading and induced angle of attack at values of constant inflow angle. The theory applies for any solidity from .01 to over 1 and any blade section camber. The effects of the duct and blade number can be determined so that the procedure applies over the entire range from two blade open propellers, to ducted helicopter tail rotors, to axial flow compressors with or without guide vanes, and to wind tunnel drive fans.

Use of potassium ferrocyanide as habit modifier in the size reduction and phase modification of ammonium nitrate crystals in slurries.

PubMed

Vargeese, Anuj A; Joshi, Satyawati S; Krishnamurthy, V N

2010-08-15

Ammonium nitrate (AN) is an inorganic crystalline compound used as a solid propellant oxidizer and as a nitrogenous fertilizer. The practical use of AN as solid propellant oxidizer is restricted due to the near room temperature polymorphic phase transition and hygroscopicity. A good deal of effort has been expended for last many years to stabilize the polymorphic transitions of AN, so as to minimize the storage difficulties of AN based fertilizers and to achieve more environmentally benign propellant systems. Also, particles with aspect ratio nearer to one are a vital requirement in fertilizer and propellant industries. In the present study AN is crystallized in presence of trace amount of potassium ferrocyanide (K(4)Fe(CN)(6)) crystal habit modifier and kept for different time intervals. And the effect of K(4)Fe(CN)(6) on the habit and phase modification of AN was studied. Phase modified ammonium nitrate (PMAN) with a particle aspect ratio nearer to one was obtained by this method and the reasons for this modifications are discussed. The morphology changes were studied by SEM, the phase modifications were studied by DSC and the structural properties were studied by powder XRD. Copyright 2010 Elsevier B.V. All rights reserved.

Solid AFFF Technology Investigation

DTIC Science & Technology

2010-12-01

Aqueous film forming foam, Solid AFFF, MIL - F - 24385 , Foam fire tests, Firefighting handlines, Shipboard fire protection 18. Distribution Statement...28 ft2) fire test in MIL - F -24385F was used as a screening method to determine the viability of the concept. It was determined during this program...military specification for AFFF, MIL - F -24385F, was used as a screening method to determine the viability of the concept. This test has been shown

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kubiak, Jonathan M.; Arnett, Lori A.

2016-01-01

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

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.

Credit PSR. This interior view of the building equipment room ...

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

Credit PSR. This interior view of the building equipment room displays heat exchangers and fan units with insulated piping for hot and cold water at left. Environmental controls and fire fighting system controls appear at right - Jet Propulsion Laboratory Edwards Facility, Propellant Curing Building, Edwards Air Force Base, Boron, Kern County, CA

It's Magic: An Educator's Vision of the Future

ERIC Educational Resources Information Center

Timmerman, Annemarie

2007-01-01

This article presents the author's vision of the future of education and focuses on instructional revolution during the 21st century when school leaders and the democracy were struggling with redesigning the educational system. The author states that the instructional revolution was itself propelled by the rapid-fire advancement in information and…

Space Storable Rocket Technology (SSRT) basic program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Mueller, T.; Casillas, A. R.; Huang, D.

1992-01-01

The Space Storable Rocket Technology Program (SSRT) was conducted to establish a technology for a new class of high performance and long life bipropellant engines using space storable propellants. The results are described. Task 1 evaluated several characteristics for a number of fuels to determine the best space storable fuel for use with LO2. The results indicated that LO2-N2H4 is the best propellant combination and provides the maximum mission/system capability maximum payload into GEO of satellites. Task 2 developed two models, performance and thermal. The performance model indicated the performance goal of specific impulse greater than or = 340 seconds (sigma = 204) could be achieved. The thermal model was developed and anchored to hot fire test data. Task 3 consisted of design, fabrication, and testing of a 200 lbf thrust test engine operating at a chamber pressure of 200 psia using LO2-N2H4. A total of 76 hot fire tests were conducted demonstrating performance greater than 340 (sigma = 204) which is a 25 second specific impulse improvement over the existing highest performance flight apogee type engines.

Experimental Study on an Unsteady Pressure Gain Combustion Hypergolic Rocket Engine Concept

NASA Astrophysics Data System (ADS)

Kan, Brandon K.

An experimental study is conducted to investigate pulsed combustion in a lab-scale bipropellant rocket engine using hypergolic propellants. The propellant combination is high concentration hydrogen peroxide and a catalyst-laced triglyme fuel. A total of 50 short duration firings have been conducted; the vast majority in an open-chamber configuration. High amplitude pulsations were evident in nearly all cases and have been assessed with high frequency pressure measurements. Both pintle and unlike impinging quadlet injector types have been evaluated although the bulk of the testing was with the latter configuration. Several firings were conducted with a transparent chamber in an attempt to gain understanding using a high-speed camera in the visible spectrum. Peak chamber pressures in excess of 5000 psi have been recorded with surface mounted high frequency gages with pulsation frequencies exceeding 600 Hz. A characterization of time-averaged performance is made for the unsteady system, where time-resolved thrust and pressure measurements were attempted. While prior literature describes this system as a pulse detonation rocket engine, the combustion appears to be more "constant volume" in nature.

Potential low cost, safe, high efficiency propellant for future space program

NASA Astrophysics Data System (ADS)

Zhou, D.

2005-03-01

Mixtures of nanometer or micrometer sized carbon powder suspended in hydrogen and methane/hydrogen mixtures are proposed as candidates for low cost, high efficiency propellants for future space programs. While liquid hydrogen has low weight and high heat of combustion per unit mass, because of the low mass density the heat of combustion per unit volume is low, and the liquid hydrogen storage container must be large. The proposed propellants can produce higher gross heat combustion with small volume with trade off of some weight increase. Liquid hydrogen can serve as the fluid component of the propellant in the mixtures and thus used by current rocket engine designs. For example, for the same volume a mixture of 5% methane and 95% hydrogen, can lead to an increase in the gross heat of combustion by about 10% and an increase in the Isp (specific impulse) by 21% compared to a pure liquid hydrogen propellant. At liquid hydrogen temperatures of 20.3 K, methane will be in solid state, and must be formed as fine granules (or slush) to satisfy the requirement of liquid propellant engines.

Modeling of combustion processes of stick propellants via combined Eulerian-Lagrangian approach

NASA Technical Reports Server (NTRS)

Kuo, K. K.; Hsieh, K. C.; Athavale, M. M.

1988-01-01

This research is motivated by the improved ballistic performance of large-caliber guns using stick propellant charges. A comprehensive theoretical model for predicting the flame spreading, combustion, and grain deformation phenomena of long, unslotted stick propellants is presented. The formulation is based upon a combined Eulerian-Lagrangian approach to simulate special characteristics of the two phase combustion process in a cartridge loaded with a bundle of sticks. The model considers five separate regions consisting of the internal perforation, the solid phase, the external interstitial gas phase, and two lumped parameter regions at either end of the stick bundle. For the external gas phase region, a set of transient one-dimensional fluid-dynamic equations using the Eulerian approach is obtained; governing equations for the stick propellants are formulated using the Lagrangian approach. The motion of a representative stick is derived by considering the forces acting on the entire propellant stick. The instantaneous temperature and stress fields in the stick propellant are modeled by considering the transient axisymmetric heat conduction equation and dynamic structural analysis.

Toxic trace elements in solid airborne particles and ecological risk assessment in the vicinity of local boiler house plants

NASA Astrophysics Data System (ADS)

Talovskaya, Anna V.; Osipova, Nina A.; Yazikov, Egor G.; Shakhova, Tatyana S.

2017-11-01

The article deals with assessment of anthropogenic pollution in vicinity of local boilers using the data on microelement composition of solid airborne particles deposited in snow. The anthropogenic feature of elevated accumulation levels of solid airborne particles deposited in snow in the vicinity of coal-fired boiler house is revealed in elevated concentrations (3-25 higher than background) of Cd, Sb, Mo, Pb, Sr, Ba, Ni, Mo, Zn and Co. In the vicinity oil-fired boiler house the specific elements as parts of solid airborne particles deposited in snow are V, Ni and Sb, as their content exceeds the background from 3 to 8 times. It is determined that the maximum shares in non-carcinogenic human health risk from chronic inhalation of trace elements to the human body in the vicinity of coal-fired boiler house belong to Al, Mn, Cu, Ba, Co, Pb, whereas in the vicinity of oil-fired boiler house - Al, Mn, Cu, Ni, V.

KSC-06pd0087

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — Viewed from a vantage point on the nearby river bank, NASA’s New Horizons spacecraft roars into the cloud-scattered sky trailing fire from the Atlas V rocket that propels it. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/George Shelton

KSC-06pd0088

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — Viewed from a nearby vantage point, NASA’s New Horizons spacecraft roars into the cloud-scattered sky trailing fire and smoke from the Atlas V rocket that propels it. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/George Shelton

KSC-06pd0091

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — Viewed from the NASA News Center, NASA’s New Horizons spacecraft roars into the cloud-scattered sky trailing fire and smoke from the Atlas V rocket that propels it. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/Fletch Hildreth

KSC-06pd0089

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — Photographers and spectators watch NASA’s New Horizons spacecraft, trailing fire and smoke from the Atlas V rocket that propels it, as it roars into the cloud-scattered sky. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/George Shelton

KSC-06pd0092

NASA Image and Video Library

2006-01-19

KENNEDY SPACE CENTER, FLA. — Viewed from the NASA News Center, NASA’s New Horizons spacecraft roars into the cloud-scattered sky trailing fire and smoke from the Atlas V rocket that propels it. Liftoff was on time at 2 p.m. EST from Complex 41 on Cape Canaveral Air Force Station in Florida. This was the third launch attempt in as many days after scrubs due to weather concerns. The compact, 1,050-pound piano-sized probe will get a boost from a kick-stage solid propellant motor for its journey to Pluto. New Horizons will be the fastest spacecraft ever launched, reaching lunar orbit distance in just nine hours and passing Jupiter 13 months later. The New Horizons science payload, developed under direction of Southwest Research Institute, includes imaging infrared and ultraviolet spectrometers, a multi-color camera, a long-range telescopic camera, two particle spectrometers, a space-dust detector and a radio science experiment. The dust counter was designed and built by students at the University of Colorado, Boulder. The launch at this time allows New Horizons to fly past Jupiter in early 2007 and use the planet’s gravity as a slingshot toward Pluto. The Jupiter flyby trims the trip to Pluto by as many as five years and provides opportunities to test the spacecraft’s instruments and flyby capabilities on the Jupiter system. New Horizons could reach the Pluto system as early as mid-2015, conducting a five-month-long study possible only from the close-up vantage of a spacecraft. Photo credit: NASA/Fletch Hildreth

The Direction of Fluid Dynamics for Liquid Propulsion at NASA Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Griffin, Lisa W.

2012-01-01

Marshall Space Flight Center (MSFC) is the National Aeronautics and Space Administration (NASA)-designated center for the development of space launch systems. MSFC is particularly known for propulsion system development. Many engineering skills and technical disciplines are needed to accomplish this mission. This presentation will focus on the work of the Fluid Dynamics Branch (ER42). ER42 resides in the Propulsion Systems Department at MSFC. The branch is responsible for all aspects of the discipline of fluid dynamics applied to propulsion or propulsion-induced loads and environments. This work begins with design trades and parametric studies, and continues through development, risk assessment, anomaly investigation and resolution, and failure investigations. Applications include the propellant delivery system including the main propulsion system (MPS) and turbomachinery; combustion devices for liquid engines and solid rocket motors; coupled systems; and launch environments. An advantage of the branch is that it is neither analysis nor test centric, but discipline centric. Fluid dynamics assessments are made by analysis, from lumped parameter modeling through unsteady computational fluid dynamics (CFD); testing, which can be cold flow or hot fire; or a combination of analysis and testing. Integration of all discipline methods into one branch enables efficient and accurate support to the projects. To accomplish this work, the branch currently employs approximately fifty engineers divided into four teams -- Propellant Delivery CFD, Combustion Driven Flows CFD, Unsteady and Experimental Flows, and Acoustics and Stability. This discussion will highlight some of the work performed in the branch and the direction in which the branch is headed.

ADAPTATION OF A TECHNIQUE FOR PREDICTING LARGE SOLID ROCKET MOTOR SPECIFIC IMPULSE FROM DATA OBTAINED IN MICROMOTORS.

DTIC Science & Technology

Laboratory. The purpose of this technique is to predict specific impulse in large solid rocket motors based on data obtained in micromotors . As little as 2...concerning performance of a propellant in a large solid motor. Predictions, based on data obtained in micromotors , were within 0.6% of the delivered impulse in 6-pound motors and 70-pound BATES motors. (Author)

Crack Propagation and Branching in Burning Solid Propellants and Ignition of Nitramine-Based Composite Propellants

DTIC Science & Technology

1986-01-01

TAPE ), %V.. ,RECORDER], WAVEFORM ,DIGITIZER AND RECORDER RPLOTER CPUTER OSCILLOSCOPE Fig. 19 S&*tic Diagram of Dat Acqisition Sytem...signals from pressure transducers are amplified by charge atplifiers and then recoiJed on - high-speed magnetic tape recorder and a 2 MHz transient...R. A. Schapery, Dec. 1985 A 92 11. Erdogan , F., "Fracture Mechanics Notes," Department of Mechanical Engineering and Mechanics, Lehigh University

Alternate propellant program, phase 1

NASA Technical Reports Server (NTRS)

Anderson, F. A.; West, W. R.

1979-01-01

Candidate propellant systems for the shuttle booster solid rocket motor (SRM), which would eliminate, or greatly reduce, the amount of HCl produced in the exhaust of the shuttle SRM were investigated. Ammonium nitrate was selected for consideration as the main oxidizer, with ammonium perchlorate and the nitramine, cyclo-tetramethylene-tetranitramine as secondary oxidizers. The amount of ammonium perchlorate used was limited to an amount which would produce an exhaust containing no more than 3% HCl.

High performance ammonium nitrate propellant

NASA Technical Reports Server (NTRS)

Anderson, F. A. (Inventor)

1979-01-01

A high performance propellant having greatly reduced hydrogen chloride emission is presented. It is comprised of: (1) a minor amount of hydrocarbon binder (10-15%), (2) at least 85% solids including ammonium nitrate as the primary oxidizer (about 40% to 70%), (3) a significant amount (5-25%) powdered metal fuel, such as aluminum, (4) a small amount (5-25%) of ammonium perchlorate as a supplementary oxidizer, and (5) optionally a small amount (0-20%) of a nitramine.

Study of solid rocket motors for a space shuttle booster. Volume 2, book 3: Cost estimating data

NASA Technical Reports Server (NTRS)

Vanderesch, A. H.

1972-01-01

Cost estimating data for the 156 inch diameter, parallel burn solid rocket propellant engine selected for the space shuttle booster are presented. The costing aspects on the baseline motor are initially considered. From the baseline, sufficient data is obtained to provide cost estimates of alternate approaches.

Release Of Gaseous NH(3) From NH(4)CIO(4) By HTPB-Bonding Agents

NASA Technical Reports Server (NTRS)

Mccomb, James C.

1993-01-01

Report describes experimental study of rate of generation of ammonia and total amount of ammonia generated by chemical reactions between bonding agents and grains of ammonium perchlorate in solid rocket propellants. Also provides insight into mechanisms of chemical reactions between several types of organic amines with solid ammonium perchlorate.

Space Shuttle Projects

NASA Image and Video Library

1976-01-01

This image illustrates the solid rocket motor (SRM)/solid rocket booster (SRB) configuration. The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the SRM's were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment. The boosters are designed to survive water impact at almost 60 miles per hour, maintain flotation with minimal damage, and preclude corrosion of the hardware exposed to the harsh seawater environment. Under the project management of the Marshall Space Flight Center, the SRB's are assembled and refurbished by the United Space Boosters. The SRM's are provided by the Morton Thiokol Corporation.

Testing of electroformed deposited iridium/powder metallurgy rhenium rockets

NASA Technical Reports Server (NTRS)

Reed, Brian D.; Dickerson, Robert

1996-01-01

High-temperature, oxidation-resistant chamber materials offer the thermal margin for high performance and extended lifetimes for radiation-cooled rockets. Rhenium (Re) coated with iridium (Ir) allow hours of operation at 2200 C on Earth-storable propellants. One process for manufacturing Ir/Re rocket chambers is the fabrication of Re substrates by powder metallurgy (PM) and the application of Ir coatings by using electroformed deposition (ED). ED Ir coatings, however, have been found to be porous and poorly adherent. The integrity of ED Ir coatings could be improved by densification after the electroforming process. This report summarizes the testing of two 22-N, ED Ir/PM Re rocket chambers that were subjected to post-deposition treatments in an effort to densify the Ir coating. One chamber was vacuum annealed, while the other chamber was subjected to hot isostatic pressure (HIP). The chambers were tested on gaseous oxygen/gaseous hydrogen propellants, at mixture ratios that simulated the oxidizing environments of Earth-storable propellants. ne annealed ED Ir/PM Re chamber was tested for a total of 24 firings and 4.58 hr at a mixture ratio of 4.2. After only 9 firings, the annealed ED Ir coating began to blister and spall upstream of the throat. The blistering and spalling were similar to what had been experienced with unannealed, as-deposited ED Ir coatings. The HIP ED Ir/PM Re chamber was tested for a total of 91 firings and 11.45 hr at mixture ratios of 3.2 and 4.2. The HIP ED Ir coating remained adherent to the Re substrate throughout testing; there were no visible signs of coating degradation. Metallography revealed, however, thinning of the HIP Ir coating and occasional pores in the Re layer upstream of the throat. Pinholes in the Ir coating may have provided a path for oxidation of the Re substrate at these locations. The HIP ED Ir coating proved to be more effective than vacuum annealed and as-deposited ED Ir. Further densification is still required to match the integrity of chemically vapor deposited Ir coatings. Despite this, the successful long duration testing of the HIP ED Ir chamber, in an oxidizing environment comparable to Earth-storable propellants, demonstrated the viability of this Ir/Re rocket fabrication process.