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.

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

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.

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).

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

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

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

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.

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.

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...

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.

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.

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 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.

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.

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.

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

NASA Astrophysics Data System (ADS)

Li, Qiang; Liu, Peijin; He, Guoqiang

2015-05-01

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

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

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

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.

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.

Main lines of scientific and technical research at the Soviet Jet Propulsion Research Institute (RNII), 1933 - 1942

NASA Technical Reports Server (NTRS)

Shchetinkov, Y. S.

1977-01-01

The rapid development of rocketry in the U.S.S.R. during the post-war years was due largely to pre-war activity; in particular, to investigations conducted in the Jet Propulsion Research Institute (RNII). The history of RNII commenced in 1933, resulting from the merger of two rocket research organizations. Previous research was continued in areas of solid-propellant rockets, jet-assisted take-off of aircraft, liquid propellant engines (generally with nitric acid as the oxidizer), liquid-propellant rockets (generally with oxgen as the oxidizer), ram jet engines, rockets with and without wings, and rocket planes. RNII research is described and summarized for the years 1933-1942.

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.

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.

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.

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.

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.

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.

Positioning Mechanism For Hoisting

NASA Technical Reports Server (NTRS)

Marlin, John D., III; Moore, Barry J.; Myers, Robert I.

1992-01-01

Mechanism positions large, heavy objects in container for lifting out by hoist, crane, or winch. Handles objects gently and ensures they are lifted cleanly away in vertical direction without bumping container. Developed for lifting offset pieces of solid-propellant core out of rocket motor through its propellant port. Similar specialized mechanisms can be developed to lift other specially shaped, specially contained heavy objects. Track in base of mechanism guides each trunnion and piece to which attached to middle as hydraulic rods extend. When mechanism lifted, tilted pieces swing inward and come to rest on energy-absorbing paddle.

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.

Demonstration of a sterilizable solid rocket motor system

NASA Technical Reports Server (NTRS)

Mastrolia, E. J.; Santerre, G. M.; Lambert, W. L.

1975-01-01

A solid propellant rocket motor containing 60.9 Kg (134-lb) of propellant was successfully static fired after being subjected to eight heat sterilization cycles (three 54-hour cycles plus five 40-hour cycles) at 125 C (257 F). The test motor, a modified SVM-3 chamber, incorporated a flexible grain retention system of EPR rubber to relieve thermal shrinkage stresses. The propellant used in the motor was ANB-3438, and 84 wt% solids system (18 wt% aluminum) containing 66 wt% stabilized ammonium perchlorate oxidizer and a saturated hydroxylterminated polybutadiene binder. Bonding of the propellant to the EPR insulation (GenGard V-4030) was provided by the use of SD-886, an epoxy urethane restriction.

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.

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.

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.

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.

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.

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.

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

Measurements of Particulates in Solid Propellant Rocket Motors

DTIC Science & Technology

1987-10-01

gradients created during a firing, however, could be a problem. Finally, a torch was placed in the motor to study temperature effects. The nitrogen...techniques available for studying particulate behavior in solid propellant rocket motors is holography. For the exposed scene a hologram provides both...is underway to study the effects of addition of aluminum and other metallic particles on the magnitude of the performance losses in propellant motors

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.

Technology for low cost solid rocket boosters.

NASA Technical Reports Server (NTRS)

Ciepluch, C.

1971-01-01

A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

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.

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.

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.

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 considerations for a pressure-driven multi-stage rocket

NASA Astrophysics Data System (ADS)

Sauerwein, Steven Craig

2002-01-01

The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

NASTRAN cyclic symmetry capability. [application to solid rocket propellant grains and space antennas

NASA Technical Reports Server (NTRS)

Macneal, R. H.; Harder, R. L.; Mason, J. B.

1973-01-01

A development for NASTRAN which facilitates the analysis of structures made up of identical segments symmetrically arranged with respect to an axis is described. The key operation in the method is the transformation of the degrees of freedom for the structure into uncoupled symmetrical components, thereby greatly reducing the number of equations which are solved simultaneously. A further reduction occurs if each segment has a plane of reflective symmetry. The only required assumption is that the problem be linear. The capability, as developed, will be available in level 16 of NASTRAN for static stress analysis, steady state heat transfer analysis, and vibration analysis. The paper includes a discussion of the theory, a brief description of the data supplied by the user, and the results obtained for two example problems. The first problem concerns the acoustic modes of a long prismatic cavity imbedded in the propellant grain of a solid rocket motor. The second problem involves the deformations of a large space antenna. The latter example is the first application of the NASTRAN Cyclic Symmetry capability to a really large problem.

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.

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.

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...

A Novel Data System for Verification of Internal Parameters of Motor Design

NASA Technical Reports Server (NTRS)

Smith, Doug; Saint Jean, Paul; Everton, Randy; Uresk, Bonnie

2003-01-01

Three major obstacles have limited the amount of information that can be obtained from inside an operating solid rocket motor. The first is a safety issue due to the presence of live propellant interacting with classical, electrical instrumentation. The second is a pressure vessel feed through risk arising from bringing a large number of wires through the rocket motor wall safely. The third is an attachment/protection issue associated with connecting gages to live propellant. Thiokol has developed a highly miniaturized, networked, electrically isolated data system that has safely delivered information from classical, electrical instrumentation (even on the burning propellant surface) to the outside world. This system requires only four wires to deliver 80 channels of data at 2300 samples/second/channel. The feed through leak path risk is massively reduced from the current situation where each gage requires at least three pressure vessel wire penetrations. The external electrical isolation of the system is better than that of the propellant itself. This paper describes the new system.

Standardization of the carbon-phenolic materials and processes. Vol. 1: Experimental studies

NASA Technical Reports Server (NTRS)

Hall, William B.

1988-01-01

Carbon-phenolic composite materials are used as ablative material in the solid rocket motor nozzle of the Space Shuttle. The nozzle is lined with carbon cloth-phenolic resin composites. The nominal effects of the completely consumed solid propellant on the carbon-phenolic material are given. The extreme heat and erosion of the burning propellant are controlled by the carbon-phenolic composite by ablation, the heat and mass transfer process in which a large amount of heat is absorbed by sacrificially removing material from the nozzle surface. Phenolic materials ablate with the initial formation of a char. The depth of the char is a function of the heat conduction coefficient of the composite. The char layer is a very poor heat conductor so it protects the underlying phenolic composite from the high heat of the burning propellant. The nozzle component ablative liners (carbon cloth-phenolic composites) are tape wrapped, hydroclave and/or autoclave cured, machined, and assembled. The tape consists of a prepreg broadcloth. The materials flow sheet for the nozzle ablative liners is shown. The prepreg is a three component system: phenolic resin, carbon cloth, and carbon filler. This is Volume 1 of two, Experimental Studies.

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.

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.

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

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.

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.

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.

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.

Fe(0) Nanomotors in Ton Quantities (10(20) Units) for Environmental Remediation.

PubMed

Teo, Wei Zhe; Zboril, Radek; Medrik, Ivo; Pumera, Martin

2016-03-24

Despite demonstrating potential for environmental remediation and biomedical applications, the practical environmental applications of autonomous self-propelled micro-/nanorobots have been limited by the inability to fabricate these devices in large (kilograms/tons) quantities. In view of the demand for large-scale environmental remediation by micro-/nanomotors, which are easily synthesized and powered by nontoxic fuel, we have developed bubble-propelled Fe(0) Janus nanomotors by a facile thermally induced solid-state procedure and investigated their potential as decontamination agents of pollutants. These Fe(0) Janus nanomotors, stabilized by an ultrathin iron oxide shell, were fuelled by their decomposition in citric acid, leading to the asymmetric bubble propulsion. The degradation of azo-dyes was dramatically increased in the presence of moving self-propelled Fe(0) nanomotors, which acted as reducing agents. Such enhanced pollutant decomposition triggered by biocompatible Fe(0) (nanoscale zero-valent iron motors), which can be handled in the air and fabricated in ton quantities for low cost, will revolutionize the way that environmental remediation is carried out. © 2016 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

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.

Solid-propellant motors for high-incremental-velocity low-acceleration maneuvers in space

NASA Technical Reports Server (NTRS)

Shafer, J. I.

1972-01-01

The applicability of solid-propellant rockets into a regime of high-performance long-burning tasks beyond the capability of existing motors is discussed. Successful static test firings have demonstrated the feasibility of: (1) utilizing fully case-bonded end-burning propellant charges without mechanical stress relief; (2) using an all-carbon radiative nozzle markedly lighter than the flight-weight ablative nozzle it replaces, and (3) producing low spacecraft acceleration rates during the thrust transient through a controlled-flow igniter that promotes operation below the previous combustion limit.

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).

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.

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.

Accuracy of real time radiography burning rate measurement

NASA Astrophysics Data System (ADS)

Olaniyi, Bisola

The design of a solid propellant rocket motor requires the determination of a propellant's burning-rate and its dependency upon environmental parameters. The requirement that the burning-rate be physically measured, establishes the need for methods and equipment to obtain such data. A literature review reveals that no measurement has provided the desired burning rate accuracy. In the current study, flash x-ray modeling and digitized film-density data were employed to predict motor-port area to length ratio. The pre-fired port-areas and base burning rate were within 2.5% and 1.2% of their known values, respectively. To verify the accuracy of the method, a continuous x-ray and a solid propellant rocket motor model (Plexiglas cylinder) were used. The solid propellant motor model was translated laterally through a real-time radiography system at different speeds simulating different burning rates. X-ray images were captured and the burning-rate was then determined. The measured burning rate was within 1.65% of the known values.

Analysis of solid propellant combustion in a closed vessel including secondary reaction

NASA Technical Reports Server (NTRS)

Benreuven, M.; Summerfield, M.

1980-01-01

A theory for combustion of solid propellants in a closed vessel is presented allowing for residual exothermic chemical reaction in the bulk of the gas in the vessel. Particular attention is given to propellants exhibiting thick gaseous flame zones such as nitrocellulose, double-base and nitramine propellants. For these, the reaction at high pressures is assumed to involve mainly the oxidation of residual hydrocarbons by NO. It is shown that the direct dynamic coupling between the exothermicity, the molecular weight reduction and the changing pressure can influence the dp/dt-p traces obtained, in a manner not directly related to mass burning rate of the solid. Energy and species conservation equations are derived for the bulk of the vessel in differential form; the system is solved numerically. The results show the effect of extended chemical reaction upon measurable combustion characteristics such as dp/dt-p and burn rate pressure exponent, demonstrating its potential importance in interpretation of closed vessel firing data, depending on the pace of the residual gas phase reactions.

Advanced Booster Composite Case/Polybenzimidazole Nitrile Butadiene Rubber Insulation Development

NASA Technical Reports Server (NTRS)

Gentz, Steve; Taylor, Robert; Nettles, Mindy

2015-01-01

The NASA Engineering and Safety Center (NESC) was requested to examine processing sensitivities (e.g., cure temperature control/variance, debonds, density variations) of polybenzimidazole nitrile butadiene rubber (PBI-NBR) insulation, case fiber, and resin systems and to evaluate nondestructive evaluation (NDE) and damage tolerance methods/models required to support human-rated composite motor cases. The proposed use of composite motor cases in Blocks IA and II was expected to increase performance capability through optimizing operating pressure and increasing propellant mass fraction. This assessment was to support the evaluation of risk reduction for large booster component development/fabrication, NDE of low mass-to-strength ratio material structures, and solid booster propellant formulation as requested in the Space Launch System NASA Research Announcement for Advanced Booster Engineering Demonstration and/or Risk Reduction. Composite case materials and high-energy propellants represent an enabling capability in the Agency's ability to provide affordable, high-performing advanced booster concepts. The NESC team was requested to provide an assessment of co- and multiple-cure processing of composite case and PBI-NBR insulation materials and evaluation of high-energy propellant formulations.

Fluid-dynamically coupled solid propellant combustion instability - cold flow simulation

NASA Astrophysics Data System (ADS)

Ben-Reuven, M.

1983-10-01

The near-wall processes in an injected, axisymmetric, viscous flow is examined. Solid propellant rocket instability, in which cold flow simulation is evaluated as a tool to elucidate possible instability driving mechanisms is studied. One such prominent mechanism seems to be visco-acoustic coupling. The formulation is presented in terms of a singular boundary layer problem, with detail (up to second order) given only to the near wall region. The injection Reynolds number is assumed large, and its inverse square root serves as an appropriate small perturbation quantity. The injected Mach number is also small, and taken of the same order as the aforesaid small quantity. The radial-dependence of the inner solutions up to second order is solved, in polynominal form. This leaves the (x,t) dependence to much simpler partial differential equations. Particular results demonstrate the existence of a first order pressure perturbation, which arises due to the dissipative near wall processes. This pressure and the associated viscous friction coefficient are shown to agree very well with experimental injected flow data.

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.

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.

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.

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.

Solid motor aft closure insulation erosion. [heat flux correlation for rate analysis

NASA Technical Reports Server (NTRS)

Stampfl, E.; Landsbaum, E. M.

1973-01-01

The erosion rate of aft closure insulation in a number of large solid propellant motors was empirically analyzed by correlating the average ablation rate with a number of variables that had previously been demonstrated to affect heat flux. The main correlating parameter was a heat flux based on the simplified Bartz heat transfer coefficient corrected for two-dimensional effects. A multiplying group contained terms related to port-to-throat ratio, local wall angle, grain geometry and nozzle cant angle. The resulting equation gave a good correlation and is a useful design tool.

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.

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.

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.

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.

Instrumentation of sampling aircraft for measurement of launch vehicle effluents

NASA Technical Reports Server (NTRS)

Wornom, D. E.; Woods, D. C.; Thomas, M. E.; Tyson, R. W.

1977-01-01

An aircraft was selected and instrumented to measure effluents emitted from large solid propellant rockets during launch activities. The considerations involved in aircraft selection, sampling probes, and instrumentation are discussed with respect to obtaining valid airborne measurements. Discussions of the data acquisition system used, the instrument power system, and operational sampling procedures are included. Representative measurements obtained from an actual rocket launch monitoring activity are also presented.

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.

An Extension of Holographic Moiré to Micromechanics

NASA Astrophysics Data System (ADS)

Sciammarella, C. A.; Sciammarella, F. M.

The electronic Holographic Moiré is an ideal tool for micromechanics studies. It does not require a modification of the surface by the introduction of a reference grating. This is of particular advantage when dealing with materials such as solid propellant grains whose chemical nature and surface finish makes the application of a reference grating very difficult. Traditional electronic Holographic Moiré presents some difficult problems when large magnifications are needed and large rigid body motion takes place. This paper presents developments that solves these problems and extends the application of the technique to micromechanics.

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.

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.

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.

Space Shuttle Projects

NASA Image and Video Library

1989-01-20

This photograph shows a static firing test of the Solid Rocket Qualification Motor-8 (QM-8) at the Morton Thiokol Test Site in Wasatch, Utah. The twin solid rocket 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.

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.

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.

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.

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).

An Investigation of Particulate Behavior in Solid Rocket Motors

DTIC Science & Technology

1981-06-01

that in the latter only a relatively few Al203 particles (of circular cross-section) are present. The other residue appears to be from the inhibitor ...cast in the propellant (Figure 16). The presence of large amounts of inhibitor residue obviously affected the scattered-light intensity profile and the...calculations. Therefore, the quantity of inhibitor used in future experi- ments should be minimized. D. DISCUSSION OF RESULTS The volume-surface mean

Atomic hydrogen as a launch vehicle propellant

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1990-01-01

An analysis of several atomic hydrogen launch vehicles was conducted. A discussion of the facilities and the technologies that would be needed for these vehicles is also presented. The Gross Liftoff Weights (GLOW) for two systems were estimated; their specific impulses (I sub sp) were 750 and 1500 lb (sub f)/s/lb(sub m). The atomic hydrogen launch vehicles were also compared to the currently planned Advanced Launch System design concepts. Very significant GLOW reductions of 52 to 58 percent are possible over the Advanced Launch System designs. Applying atomic hydrogen propellants to upper stages was also considered. Very high I(sub sp) (greater than 750 1b(sub f)/s/lb(sub m) is needed to enable a mass savings over advanced oxygen/hydrogen propulsion. Associated with the potential benefits of high I(sub sp) atomic hydrogen are several challenging problems. Very high magnetic fields are required to maintain the atomic hydrogen in a solid kilogauss (3 Tesla). Also the storage temperature of the propellant is 4 K. This very low temperature will require a large refrigeration facility for the launch vehicle. The design considerations for a very high recombination rate for the propellant are also discussed. A recombination rate of 210 cm/s is predicted for atomic hydrogen. This high recombination rate can produce very high acceleration for the launch vehicle. Unique insulation or segmentation to inhibit the propellant may be needed to reduce its recombination rate.

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.

AP reclamation and reuse in RSRM propellant

NASA Technical Reports Server (NTRS)

Miks, Kathryn F.; Harris, Stacey A.

1995-01-01

A solid propellant ingredient reclamation pilot plant has been evaluated at the Strategic Operations of Thiokol Corporation, located in Brigham City, Utah. The plant produces AP wet cake (95 percent AP, 5 percent water) for recycling at AP vendors. AP has been obtained from two standard propellant binder systems (PBAN and HTPB). Analytical work conducted at Thiokol indicates that the vendor-recrystallized AP meets Space Shuttle propellant specification requirements. Thiokol has processed 1-, 5-, and 600-gallon propellant mixes with the recrystallized AP. Processing, cast, cure, ballistic, mechanical, and safety properties have been evaluated. Phillips Laboratory static-test-fired 70-pound and 800-pound BATES motors. The data indicate that propellant processed with reclaimed AP has nominal properties.

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.

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

On the importance of reduced scale Ariane 5 P230 solid rocket motor models in the comprehension and prevention of thrust oscillations

NASA Astrophysics Data System (ADS)

Hijlkema, J.; Prévost, M.; Casalis, G.

2011-09-01

Down-scaled solid propellant motors are a valuable tool in the study of thrust oscillations and the underlying, vortex-shedding-induced, pressure instabilities. These fluctuations, observed in large segmented solid rocket motors such as the Ariane 5 P230, impose a serious constraint on both structure and payload. This paper contains a survey of the numerous configurations tested at ONERA over the last 20 years. Presented are the phenomena searched to reproduce and the successes and failures of the different approaches tried. The results of over 130 experiments have contributed to numerous studies aimed at understanding the complicated physics behind this thorny problem, in order to pave the way to pressure instability reduction measures. Slowly but surely our understanding of what makes large segmented solid boosters exhibit this type of instabilities will lead to realistic modifications that will allow for a reduction of pressure oscillations. A "quieter" launcher will be an important advantage in an ever more competitive market, giving a easier ride to payload and designers alike.

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.

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.

Firing test of propellant-cracked solid motor under X-ray TV

NASA Astrophysics Data System (ADS)

Fujiwara, Tsutomu; Tanemura, Toshiharu; Itoh, Katsuya; Kakuta, Yoshiaki; Shimizu, Morio; Takahashi, Michio

This paper presents the effects of a big crack on the combustion behaviors of the scaled-down Japanese H-I upper stage motors of the National Space Development Agency (NASDA). The big crack was generated by cooling down the propellant grain below -100 C; the crack was identified and measured with the X-ray computer tomography (CT) system designed for medical use. It was found that the crack spread widely from inner bore to liner and fore-and-aft of the motor. The firing test of the propellant-cracked solid motor was performed under X-ray TV observation, and the motor exploded just after the ignition because of the abrupt chamber pressure increase due to flame propagation into the crack.

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

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

NASA Technical Reports Server (NTRS)

Dushkin, L. S.

1977-01-01

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

Summary of Air Force Research Laboratory Support for the NASA Green Propellant Infusion Mission

DTIC Science & Technology

2015-07-01

system to transfer propellant from a bulk propellant tank into a spacecraft tank. It also called for the transfer of propellant from a large transport...launch pressurized propellant tanks on a spacecraft or satellite, a fracture mechanics analysis is required to verify the safe design life of the...a bulk propellant tank into a spacecraft tank. It also called for the transfer of propellant from a large transport container into a specialized

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.

Conceptual Launch Vehicles Using Metallic Hydrogen Propellant

NASA Astrophysics Data System (ADS)

Cole, John W.; Silvera, Isaac F.; Foote, John P.

2008-01-01

Solid molecular hydrogen is predicted to transform into an atomic solid with metallic properties under pressures >4.5 Mbar. Atomic metallic hydrogen is predicted to be metastable, limited by some critical temperature and pressure, and to store very large amounts of energy. Experiments may soon determine the critical temperature, critical pressure, and specific energy availability. It is useful to consider the feasibility of using metastable atomic hydrogen as a rocket propellant. If one assumes that metallic hydrogen is stable at usable temperatures and pressures, and that it can be affordably produced, handled, and stored, then it may be a useful rocket propellant. Assuming further that the available specific energy can be determined from the recombination of the atoms into molecules (216 MJ/kg), then conceptual engines and launch vehicle concepts can be developed. Under these assumptions, metallic hydrogen would be a revolutionary new rocket fuel with a theoretical specific impulse of 1700 s at a chamber pressure of 100 atm. A practical problem that arises is that rocket chamber temperatures may be too high for the use of this pure fuel. This paper examines an engine concept that uses liquid hydrogen or water as a diluent coolant for the metallic hydrogen to reduce the chamber temperature to usable values. Several launch vehicles are then conceptually developed. Results indicate that if metallic hydrogen is experimentally found to have the properties assumed in this analysis, then there are significant benefits. These benefits become more attractive as the chamber temperatures increase.

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.

Experimental and Numerical Study of Ammonium Perchlorate Counterflow Diffusion Flames

NASA Technical Reports Server (NTRS)

Smooke, M. D.; Yetter, R. A.; Parr, T. P.; Hanson-Parr, D. M.; Tanoff, M. A.

1999-01-01

Many solid rocket propellants are based on a composite mixture of ammonium perchlorate (AP) oxidizer and polymeric binder fuels. In these propellants, complex three-dimensional diffusion flame structures between the AP and binder decomposition products, dependent upon the length scales of the heterogeneous mixture, drive the combustion via heat transfer back to the surface. Changing the AP crystal size changes the burn rate of such propellants. Large AP crystals are governed by the cooler AP self-deflagration flame and burn slowly, while small AP crystals are governed more by the hot diffusion flame with the binder and burn faster. This allows control of composite propellant ballistic properties via particle size variation. Previous measurements on these diffusion flames in the planar two-dimensional sandwich configuration yielded insight into controlling flame structure, but there are several drawbacks that make comparison with modeling difficult. First, the flames are two-dimensional and this makes modeling much more complex computationally than with one-dimensional problems, such as RDX self- and laser-supported deflagration. In addition, little is known about the nature, concentration, and evolution rates of the gaseous chemical species produced by the various binders as they decompose. This makes comparison with models quite difficult. Alternatively, counterflow flames provide an excellent geometric configuration within which AP/binder diffusion flames can be studied both experimentally and computationally.

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.

Effect of silicone oil on solid propellant combustion in small motors. [for rockets

NASA Technical Reports Server (NTRS)

Ramohalli, K.

1980-01-01

The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.

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.

Boron epoxy rocket motor case program

NASA Technical Reports Server (NTRS)

Stang, D. A.

1971-01-01

Three 28-inch-diameter solid rocket motor cases were fabricated using 1/8 inch wide boron/epoxy tape. The cases had unequal end closures (4-1/8-inch-diameter forward flanges and 13-inch-diameter aft flanges) and metal attachment skirts. The flanges and skirts were titanium 6Al-4V alloy. The original design for the first case was patterned after the requirements of the Applications Technology Satellite apogee kick motor. The second and third cases were designed and fabricated to approximate the requirements of a small Applications Technology Satellite apogee kick motor. The program demonstrated the feasibility of designing and fabricating large-scale filament-wound solid propellant rocket motor cases with boron/epoxy tape.

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

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.

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

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.

Solid Rocket Launch Vehicle Explosion Environments

NASA Technical Reports Server (NTRS)

Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.

2014-01-01

Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

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.

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.

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.

Atomic hydrogen as a launch vehicle propellant

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1990-01-01

An analysis of several atomic hydrogen launch vehicles was conducted. A discussion of the facilities and the technologies that would be needed for these vehicles is also presented. The Gross Liftoff Weights (GLOW) for two systems were estimated; their specific impulses (I sub sp) were 750 and 1500 lb(sub f)/s/lb(sub m). The atomic hydrogen launch vehicles were also compared to the currently planned Advanced Launch System design concepts. Very significant GLOW reductions of 52 to 58 percent are possible over the Advanced Launch System designs. Applying atomic hydrogen propellants to upper stages was also considered. Very high I(sub sp) (greater than 750 lb(sub f)/s/lb(sub m)) is needed to enable a mass savings over advanced oxygen/hydrogen propulsion. Associated with the potential benefits of high I(sub sp) atomic hydrogen are several challenging problems. Very high magnetic fields are required to maintain the atomic hydrogen in a solid hydrogen matrix. The magnetic field strength was estimated to be 30 kilogauss (3 Tesla). Also the storage temperature of the propellant is 4 K. This very low temperature will require a large refrigeration facility for the launch vehicle. The design considerations for a very high recombination rate for the propellant are also discussed. A recombination rate of 210 cm/s is predicted for atomic hydrogen. This high recombination rate can produce very high acceleration for the launch vehicle. Unique insulation or segmentation to inhibit the propellant may be needed to reduce its recombination rate.

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).

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.

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.

Dark Zones of Solid Propellant Flames: Critical Assessment and Quantitative Modeling of Experimental Datasets With Analysis of Chemical Pathways and Sensitivities

DTIC Science & Technology

2011-01-01

with a collection of information if it does not display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 1...Research Associate at ARL with WRA, and largely completed more recently while at Dept. of Chem., SUNY, Cortland, NY. Currently unaffiliated. †Former...promised to provide an extensive, definitive review critically assessing our current understanding of DZ structure and chemistry, and providing a documented

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

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

Altered combustion characteristics of metallized energetics due to stable secondary material inclusion

NASA Astrophysics Data System (ADS)

Terry, Brandon C.

Though metals and metalloids have been widely considered as reactive fuels, the ability to tune their ignition and combustion characteristics remains challenging. One means to accomplish this may be through low-level inclusion of secondary materials into the metallized fuel. While there are several potential methods to stably introduce secondary inclusion materials, this work focuses on the use of mechanical activation (MA) and metal alloys. Recent work has shown that low-level inclusion of fluoropolymers into aluminum particles can have a substantial effect on their combustion characteristics. The reflected shock ignition of mechanically activated aluminum/polytetrafluoroethylene (MA Al/PTFE) is compared to a physical mixture (PM) of Al/PTFE, neat spherical aluminum, and flake aluminum. It was found that the powders with higher specific surface areas ignited faster than the spherical particles of the same size, and had ignition delay times comparable to agglomerates of aluminum particles that were two orders of magnitude smaller in size. Flake aluminum powder had the same ignition delay as MA Al/PTFE, indicating that any initial aluminum/fluoropolymer reactions did not yield an earlier onset of aluminum oxidation. However, MA Al/PTFE did have a shorter total burn time. The PM of Al/PTFE powder had a shorter ignition delay than neat spherical aluminum due to the rapid decomposition of PTFE into reactive fluorocarbon compounds, but the subsequent fluorocarbon reactions also created a secondary luminosity profile that significantly increased the total burn time of the system. The explosive shock ignition of aluminum and aluminum-silicon eutectic alloy compacts was evaluated with and without polymer inclusions. A statistical analysis was completed, investigating the effects of: detonation train orientation (into or not into a hard surface); the high explosive driver; whether the metal/polymer system is mechanically activated; particle size; particle morphology (spherical or flake); metal type (Al or Al-Si); and whether the inclusion material is interacting or non-interacting with the parent metal. It was found that mechanically activated particles with an interacting inclusion material (polytetrafluoroethylene) and smaller particle sizes yielded increased blast wave strength, and more complete metal combustion. It was also found that orientation of the detonation train has a substantial effect on the completeness of combustion. While aluminum alloys are generally employed for their structural and mechanical properties, the low-level inclusion of secondary metals and metalloids may make such materials advantageous in propellant formulations and have not been fully considered. The aluminum-silicon (Al-Si) eutectic alloy was evaluated as a potential solid composite propellant fuel. Equilibrium calculations showed that Al-Si based propellants had comparable theoretical performance to equivalent aluminum based propellants, though at a typical specific impulse (ISP) reduction of roughly 2.5 seconds for most mixture ratios of interest. Interacting (polytetrafluoroethylene, PTFE) and non-interacting inclusion materials were mechanically activated (MA) with Al-Si (70/30 wt.% Al-Si/PTFE and 90/10 wt.% Al-Si/LDPE), which were shown to increase the powder reactivity. Neat and MA Al-Si powders were used in 15/71/14 wt.% (fuel additive)/(ammonium perchlorate)/binder propellant formulations. Environmentally cleaner solid composite propellants have been widely investigated as a means to reduce hydrochloric acid (HCl) formation. Past efforts to scavenge 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., salt). While this technique can potentially reduce HCl formation, it also results in reduced theoretical specific impulse. Thermochemical calculations show that using aluminum-lithium (Al-Li) binary alloy can reduce HCl formation to less than 5% and increase the theoretical ISP by roughly 7 seconds compared to neat aluminum. Two solid propellants were made using 80/20 Al-Li alloy and neat aluminum as fuel additives. It was observed that the propellant combustion with neat aluminum formed large molten droplets at the surface, which is a well-known problem with aluminized propellants. In contrast, the Al-Li propellant formed an Al-Li melt-layer on the propellant surface during combustion. Droplets that were ejected from the melt-layer would typically undergo dispersive boiling or a shattering microexplosion, due to the large disparity in volatility (i.e., boiling points) between the aluminum and the lithium in the molten alloy. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments. Additionally, no HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption. (Abstract shortened by UMI.).

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.

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.

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.

Agglomerates, smoke oxide particles, and carbon inclusions in condensed combustion products of an aluminized GAP-based propellant

NASA Astrophysics Data System (ADS)

Ao, Wen; Liu, Peijin; Yang, Wenjing

2016-12-01

In solid propellants, aluminum is widely used to improve the performance, however the condensed combustion products especially the large agglomerates generated from aluminum combustion significantly affect the combustion and internal flow inside the solid rocket motor. To clarify the properties of the condensed combustion products of aluminized propellants, a constant-pressure quench vessel was adopted to collect the combustion products. The morphology and chemical compositions of the collected products, were then studied by using scanning electron microscopy coupled with energy dispersive (SEM-EDS) method. Various structures have been observed in the condensed combustion products. Apart from the typical agglomerates or smoke oxide particles observed before, new structures including the smoke oxide clusters, irregular agglomerates and carbon-inclusions are discovered and investigated. Smoke oxide particles have the highest amount in the products. The highly dispersed oxide particle is spherical with very smooth surface and is on the order of 1-2 μm, but due to the high temperature and long residence time, these small particles will aggregate into smoke oxide clusters which are much larger than the initial particles. Three types of spherical agglomerates have been found. As the ambient gas temperature is much higher than the boiling point of Al2O3, the condensation layer inside which the aluminum drop is burning would evaporate quickly, which result in the fact that few "hollow agglomerates" has been found compared to "cap agglomerates" and "solid agglomerates". Irregular agglomerates usually larger than spherical agglomerates. The formation of irregular agglomerates likely happens by three stages: deformation of spherical aluminum drops; combination of particles with various shape; finally production of irregular agglomerates. EDS results show the ratio of O to Al on the surface of agglomerates is lower in comparison to smoke oxide particles. C and O account for most element compositions for all the carbon inclusions. The rough, spherical, strip shape and flake shape carbon-inclusions are believed to be derived from the degradation products of the binder or oxidizer, while the fiber silk is possibly the combustion product of fiber inside the heat insulation layer of the propellants. Images of products at different pressures reveal high pressure reduces the degree of agglomeration. The chemical compositions, size range and content of all the observed structures are given in this paper. Results of our study are expected to provide better insight in the working process of solid rocket motor.

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.

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.

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.

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.

Large-Eddy Simulation of Propeller Crashback

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2013-11-01

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield negative thrust. The crashback condition is dominated by the interaction of free stream flow with strong reverse flow. Crashback causes highly unsteady loads and flow separation on blade surface. This study uses Large-Eddy Simulation to predict the highly unsteady flow field in propeller crashback. Results are shown for a stand-alone open propeller, hull-attached open propeller and a ducted propeller. The simulations are compared to experiment, and used to discuss the essential physics behind the unsteady loads. This work is supported by the Office of Naval Research.

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.

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.

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.

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. .

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.

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.

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.

Combustion Processes in Solid Propellant Cracks

DTIC Science & Technology

1981-06-01

Ignition at the Closed End of an Inert Ctack . . ......................... 38 12. Block Diagram of Remotely-Controlled Ignition and Photography System ...41 13. Block Diagram of Data Acquisition System ... ........ .. 42 14. Measured Pressure-Time Traces for Crack...ignition system has been designed and fabricated. 5. Experimental firings with single-pore propellant grain have been conducted to study the effects of

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

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.

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

Study of a High-Energy Upper Stage for Future Shuttle Missions

NASA Technical Reports Server (NTRS)

Dressler, Gordon A.; Matuszak, Leo W.; Stephenson, David D.

2003-01-01

Space Shuttle Orbiters are likely to remain in service to 2020 or beyond for servicing the International Space Station and for launching very high value spacecraft. There is a need for a new STS-deployable upper stage that can boost certain Orbiter payloads to higher energy orbits, up to and including Earth-escape trajectories. The inventory of solid rocket motor Inertial Upper Stages has been depleted, and it is unlikely that a LOX/LH2-fueled upper stage can fly on Shuttle due to safety concerns. This paper summarizes the results of a study that investigated a low cost, low risk approach to quickly developing a new large upper stage optimized to fly on the existing Shuttle fleet. Two design reference missions (DRMs) were specified: the James Webb Space Telescope (JWST) and the Space Interferometry Mission (SIM). Two categories of upper stage propellants were examined in detail: a storable liquid propellant and a storable gel propellant. Stage subsystems 'other than propulsion were based largely on heritage hardware to minimize cost, risk and development schedule span. The paper presents the ground rules and guidelines for conducting the study, the preliminary conceptual designs margins, assessments of technology readiness/risk, potential synergy with other programs, and preliminary estimates of development and production costs and schedule spans. Although the Orbiter Columbia was baselined for the study, discussion is provided to show how the results apply to the remaining STS Orbiter fleet.

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.

Active space debris removal—A preliminary mission analysis and design

NASA Astrophysics Data System (ADS)

Castronuovo, Marco M.

2011-11-01

The active removal of five to ten large objects per year from the low Earth orbit (LEO) region is the only way to prevent the debris collisions from cascading. Among the three orbital regions near the Earth where most catastrophic collisions are predicted to occur, the one corresponding to a sun-synchronous condition is considered the most relevant. Forty-one large rocket bodies orbiting in this belt have been identified as the priority targets for removal. As part of a more comprehensive system engineering solution, a space mission dedicated to the de-orbiting of five rocket bodies per year from this orbital regime has been designed. The selected concept of operations envisages the launch of a satellite carrying a number of de-orbiting devices, such as solid propellant kits. The satellite performs a rendezvous with an identified object and mates with it by means of a robotic arm. A de-orbiting device is attached to the object by means of a second robotic arm, the object is released and the device is activated. The spacecraft travels then to the next target. The present paper shows that an active debris removal mission capable of de-orbiting 35 large objects in 7 years is technically feasible, and the resulting propellant mass budget is compatible with many existing platforms.

Fault Diagnostics and Prognostics for Large Segmented SRMs

NASA Technical Reports Server (NTRS)

Luchinsky, Dmitry; Osipov, Viatcheslav V.; Smelyanskiy, Vadim N.; Timucin, Dogan A.; Uckun, Serdar; Hayashida, Ben; Watson, Michael; McMillin, Joshua; Shook, David; Johnson, Mont;

Dynamic Simulation of VEGA SRM Bench Firing By Using Propellant Complex Characterization

NASA Astrophysics Data System (ADS)

Di Trapani, C. D.; Mastrella, E.; Bartoccini, D.; Squeo, E. A.; Mastroddi, F.; Coppotelli, G.; Linari, M.

2012-07-01

During the VEGA launcher development, from the 2004 up to now, 8 firing tests have been performed at Salto di Quirra (Sardinia, Italy) and Kourou (Guyana, Fr) with the objective to characterize and qualify of the Zefiros and P80 Solid Rocket Motors (SRM). In fact the VEGA launcher configuration foreseen 3 solid stages based on P80, Z23 and Z9 Solid Rocket Motors respectively. One of the primary objectives of the firing test is to correctly characterize the dynamic response of the SRM in order to apply such a characterization to the predictions and simulations of the VEGA launch dynamic environment. Considering that the solid propellant is around 90% of the SRM mass, it is very important to dynamically characterize it, and to increase the confidence in the simulation of the dynamic levels transmitted to the LV upper part from the SRMs. The activity is articulated in three parts: • consolidation of an experimental method for the dynamic characterization of the complex dynamic elasticity modulus of elasticity of visco-elastic materials applicable to the SRM propellant operative conditions • introduction of the complex dynamic elasticity modulus in a numerical FEM benchmark based on MSC NASTRAN solver • analysis of the effect of the introduction of the complex dynamic elasticity modulus in the Zefiros FEM focusing on experimental firing test data reproduction with numerical approach.

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.

Space Shuttle Solid Rocket Motor (SRM) development and qualification

NASA Technical Reports Server (NTRS)

Lund, R. K.; Brinton, B. C.

1980-01-01

The configuration of reusable solid propellant motors for the space shuttle vehicle is delineated and traces their design evolution. Also presented are the summary results of the first two of the three qualification motor firings designated QM-1 and QM-2.

Numerical investigation of performance of vane-type propellant management device by VOF methods

NASA Astrophysics Data System (ADS)

Liu, J. T.; Zhou, C.; Wu, Y. L.; Zhuang, B. T.; Li, Y.

2015-01-01

The orbital propellant management performance of the vane-type tank is so important for the propellant system and it determines the lifetime of the satellite. The propellant in the tank can be extruded by helium gas. To study the two phase distribution in the vane-type surface tension tank and the capability of the vane-type propellant management device (PMD), a large volume vane-type surface tension tank is analysed using 3-D unsteady numerical simulations. VOF methods are used to analyse the location of the interface of the two phase. Performances of the propellant acquisition vanes and propellant refillable reservoir in the tank are investigated. The flow conductivity of the propellant acquisition vanes and the liquid storage capacity of propellant refillable reservoir can be affected by the value of the gravity and the volume of the propellant in the tank. To avoid the large resistance causing by surface tension in an outflow of a small hole, the design of the vanes in a propellant refillable reservoir should have suitable space.

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.

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...

PHYSICAL CHARACTERISTICS AND NON-KEPLERIAN ORBITAL MOTION OF 'PROPELLER' MOONS EMBEDDED IN SATURN'S RINGS

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

Tiscareno, Matthew S.; Burns, Joseph A.; Hedman, Matthew M.

2010-08-01

We report the discovery of several large 'propeller' moons in the outer part of Saturn's A ring, objects large enough to be followed over the 5 year duration of the Cassini mission. These are the first objects ever discovered that can be tracked as individual moons, but do not orbit in empty space. We infer sizes up to 1-2 km for the unseen moonlets at the center of the propeller-shaped structures, though many structural and photometric properties of propeller structures remain unclear. Finally, we demonstrate that some propellers undergo sustained non-Keplerian orbit motion.

A Method for Incorporating Changing Structural Characteristics Due to Propellant Mass Usage in a Launch Vehicle Ascent Simulation

NASA Technical Reports Server (NTRS)

McGhee, D. S.

2004-01-01

Launch vehicles consume large quantities of propellant quickly, causing the mass properties and structural dynamics of the vehicle to change dramatically. Currently, structural load assessments account for this change with a large collection of structural models representing various propellant fill levels. This creates a large database of models complicating the delivery of reduced models and requiring extensive work for model changes. Presented here is a method to account for these mass changes in a more efficient manner. The method allows for the subtraction of propellant mass as the propellant is used in the simulation. This subtraction is done in the modal domain of the vehicle generalized model. Additional computation required is primarily for constructing the used propellant mass matrix from an initial propellant model and further matrix multiplications and subtractions. An additional eigenvalue solution is required to uncouple the new equations of motion; however, this is a much simplier calculation starting from a system that is already substantially uncoupled. The method was successfully tested in a simulation of Saturn V loads. Results from the method are compared to results from separate structural models for several propellant levels, showing excellent agreement. Further development to encompass more complicated propellant models, including slosh dynamics, is possible.

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.

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.

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.

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.

Acoustic emission strand burning technique for motor burning rate prediction

NASA Technical Reports Server (NTRS)

Christensen, W. N.

1978-01-01

An acoustic emission (AE) method is being used to measure the burning rate of solid propellant strands. This method has a precision of 0.5% and excellent burning rate correlation with both subscale and large rocket motors. The AE procedure burns the sample under water and measures the burning rate from the acoustic output. The acoustic signal provides a continuous readout during testing, which allows complete data analysis rather than the start-stop clockwires used by the conventional method. The AE method helps eliminate such problems as inhibiting the sample, pressure increase and temperature rise, during testing.

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.

Plasma propulsion for space applications

NASA Astrophysics Data System (ADS)

Fruchtman, Amnon

2000-04-01

The various mechanisms for plasma acceleration employed in electric propulsion of space vehicles will be described. Special attention will be given to the Hall thruster. Electric propulsion utilizes electric and magnetic fields to accelerate a propellant to a much higher velocity than chemical propulsion does, and, as a result, the required propellant mass is reduced. Because of limitations on electric power density, electric thrusters will be low thrust engines compared with chemical rockets. The large jet velocity and small thrust of electric thrusters make them most suitable for space applications such as station keeping of GEO communication satellites, low orbit drag compensation, orbit raising and interplanetary missions. The acceleration in the thruster is either thermal, electrostatic or electromagnetic. The arcjet is an electrothermal device in which the propellant is heated by an electric arc and accelerated while passing through a supersonic nozzle to a relatively low velocity. In the Pulsed Plasma Thruster a solid propellant is accelerated by a magnetic field pressure in a way that is similar in principle to pulsed acceleration of plasmas in other, very different devices, such as the railgun or the plasma opening switch. Magnetoplasmadynamic thrusters also employ magnetic field pressure for the acceleration but with a reasonable efficiency at high power only. In an ion thruster ions are extracted from a plasma through a double grid structure. Ion thrusters provide a high jet velocity but the thrust density is low due to space-charge limitations. The Hall thruster, which in recent years has enjoyed impressive progress, employs a quasi-neutral plasma, and therefore is not subject to a space-charge limit on the current. An applied radial magnetic field impedes the mobility of the electrons so that the applied potential drops across a large region inside the plasma. Methods for separately controlling the profiles of the electric and the magnetic fields will be described. The role of the sonic transition in plasma accelerators will be discussed. It will be shown that large potential drops can be localized to regions of an abrupt sonic transition in a Hall plasma. A configuration with segmented side electrodes can be used to further control the electric field profile and to increase the efficiency.

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.

Atmospheric scavenging of solid rocket exhaust effluents

NASA Technical Reports Server (NTRS)

Fenton, D. L.; Purcell, R. Y.

1978-01-01

Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

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

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.

Large-payload earth-orbit transportation with electric propulsion

NASA Technical Reports Server (NTRS)

Stearns, J. W.

1976-01-01

Economical unmanned earth orbit transportation for large payloads is evaluated. The high exhaust velocity achievable with electric propulsion is attractive because it minimizes the propellant that must be carried to low earth orbit. Propellant transport is a principal cost item. Electric propulsion subsystems utilizing advanced ion thrusters are compared to magnetoplasmadynamic (MPD) thrust subsystems. For very large payloads, a large lift vehicle is needed to low earth orbit, and argon propellant is required for electric propulsion. Under these circumstances, the MPD thruster is shown to be desirable over the ion thruster for earth orbit transportation.

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.

Earth-to-orbit propellant transportation overview

NASA Technical Reports Server (NTRS)

Fester, D.

1984-01-01

The transportation of large quantities of cryogenic propellants which are needed to support Space Station/OTV operation is discussed. Two ways to send propellants into space are: transporting them in dedicated tankers or scavenging unused STS propellant. Scavenging propellant, both with and without an aft cargo carrier system is examined. An average of two to four flights per year can be saved by scavenging and manifesting propellant as payload. Addition of an aft cargo carrier permits loading closer to maximum, reduces the required number of flights, and reduces the propellant available for scavenging. Sufficient propellant remains, however, for OTV needs.

Drag and Propulsive Characteristics of Air-Cooled Engine-Nacelle Installations for Large Airplane

NASA Technical Reports Server (NTRS)

Silverstein, Abe; Wilson, Herbert A , Jr

1942-01-01

An investigation was conducted in the NACA full-scale wind tunnel to determine the drag and the propulsive efficiency of nacelle-propeller arrangements for a large range of nacelle sizes. In contrast with usual tests with a single nacelle, these tests were conducted with nacelle-propeller installations on a large model of a four-engine airplane. Data are presented on the first part of the investigation, covering seven nacelle arrangements with nacelle diameters from 0.53 to 1.5 times the wing thickness. These ratios are similar to those occurring on airplanes weighing from about 20 to 100 tons. The results show the drag, the propulsive efficiency, and the over-all efficiency of the various nacelle arrangements as functions of the nacelle size, the propeller position, and the airplane lift coefficient. The effect of the nacelles on the aerodynamic characteristics of the model is shown for both propeller-removed and propeller-operating conditions.

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.

Summary and recent results from the NASA advanced High Speed Propeller Research Program

NASA Technical Reports Server (NTRS)

Mitchell, G. A.; Mikkelson, D. C.

1982-01-01

Advanced high-speed propellers offer large performance improvements for aircraft that cruise in the Mach 0.7 to 0.8 speed regime. The current status of the NASA research program on high-speed propeller aerodynamics, acoustics, and aeroelastics is described. Recent wind tunnel results for five 8- to 10-blade advanced models are compared with analytical predictions. Test results show that blade sweep was important in achieving net efficiencies near 80 percent at Mach 0.8 and reducing near-field cruise noise by dB. Lifting line and lifting surface aerodynamic analysis codes are under development and some initial lifting line results are compared with propeller force and probe data. Some initial laser velocimeter measurements of the flow field velocities of an 8-bladed 45 deg swept propeller are shown. Experimental aeroelastic results indicate that cascade effects and blade sweep strongly affect propeller aeroelastic characteristics. Comparisons of propeller near-field noise data with linear acoustic theory indicate that the theory adequate predicts near-field noise for subsonic tip speeds but overpredicts the noise for supersonic tip speeds. Potential large gains in propeller efficiency of 7 to 11 percent at Mach 0.8 may be possible with advanced counter-rotation propellers.

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.

Effects of solid-propellant temperature gradients on the internal ballistics of the Space Shuttle

NASA Technical Reports Server (NTRS)

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

1978-01-01

The internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRMs). A simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch. The internal ballistic effects of the gradients are established by use of a mathematical model which permits the propellant burning rate to vary circumferentially. Comparative results are presented for uniform and axisymmetric temperature distributions and the anticipated gradients based on an earlier two-dimensional analysis of the center SRM segment. The thrust imbalance potential of the booster stage is also assessed based on the difference in the thermal loading of the individual SRMs of the motor pair which may be encountered in both summer and winter environments at the launch site. Results indicate that grain temperature gradients could cause the thrust imbalance to be approximately 10% higher in the Space Shuttle than the imbalance caused by SRM manufacturing and propellant physical property variability alone.

Hybrid boosters for future launch vehicles

NASA Astrophysics Data System (ADS)

Dargies, E.; Lo, R. E.

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

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.

Large Scale Production of Densified Hydrogen Using Integrated Refrigeration and Storage

NASA Technical Reports Server (NTRS)

Notardonato, William U.; Swanger, Adam Michael; Jumper, Kevin M.; Fesmire, James E.; Tomsik, Thomas M.; Johnson, Wesley L.

2017-01-01

Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage (IRAS) technology at NASA Kennedy Space Center led to the production of large quantities of solid densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. System energy balances and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing (up to 1 K), and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. Twenty silicon diode temperature sensors were recorded over approximately one month for testing at two different fill levels (33 67). The phenomenon, observed at both two fill levels, is described and presented detailed and explained herein., and The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Dependency of the apparent contact angle on nonisothermal conditions

NASA Astrophysics Data System (ADS)

Krahl, Rolf; Gerstmann, Jens; Behruzi, Philipp; Bänsch, Eberhard; Dreyer, Michael E.

2008-04-01

The dynamic behavior of liquids in partly filled containers is influenced to a large extend by the angle between the gas-liquid phase boundary and the solid container wall at the contact line. This contact angle in turn is influenced by nonisothermal conditions. In the case of a cold liquid meniscus spreading over a hot solid wall, the contact angle apparently becomes significantly larger. In this paper we want to establish a quantitative equation for this enlargement, both from experimental and numerical data. Our findings can be used to build a subgrid model for computations, where the resolution is not sufficient to resolve the boundary layers. This might be the case for large containers which are exposed to low accelerations and where the contact angle boundary condition determines the position of the free surface. These types of computation are performed, for example, to solve propellant management problems in launcher and satellite tanks. In this application, the knowledge of the position of the free surface is very important for the withdrawal of liquid and the calculation of heat and mass transfer.

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Gazzaniga, John A.

1989-01-01

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Resonant Laser Ignition Study of HAN-HEHN Propellant Mixture (Preprint)

DTIC Science & Technology

2008-07-17

results to larger samples can be predicted by the adaptation of modeling 4 formalism previously reported for solid propellant laser ignition (15-17...The inclusion of a chemical heat release term in the form of an Arrhenius expression within a heat conduction model can also give valuable...the face of the pressure transducer. In this case the BaF2 cell entrance window failed quietly at 30 µs following the initial shock sequence. The

Propellant Nonlinear Constitutive Theory Extension: Preliminary Results.

DTIC Science & Technology

1983-08-01

Farris, R. J., Hermann , L. R., Hutchinson, J. R., and Schapery, R. A., "Development of a Solid Rocket Propellant Nonlinear Viscoelastic Constitu- tive...Publication 331, Dec. 1980. pp. 127- 133. 27. Mullins, L., "Softening of Rubber by Deformation," Rubber Chem. Technol., 1969, Vol. 31, pp. 333-362. 28. Oberth ...June 1973. 30. Hermann , L. R., and Peterson, F. E., "A Numerical Procedure for Viscoelastic Stress Analysis," Proc. 7th Mtg. of ICRPG Mech. Beh

Boris Novozhilov: Life and contribution to the physics of combustion

NASA Astrophysics Data System (ADS)

Novozhilov, Vasily

2018-04-01

Professor Boris Novozhilov (1930-2017) passed away on February 19th, 2017 in Moscow. The present paper provides brief account of his life and contributions to the physics of combustion. From extensive scientific legacy left by Boris, several major achievements are discussed here: Zeldovich-Novozhilov (ZN) theory of unsteady solid propellant combustion, contributions to thermal explosion theory, the theory of spin combustion, discovery of propellant combustion transition to chaotic regimes through Feigenbaum period bifurcation scenario.

Holographic Investigation of Solid Propellant Particulates.

DTIC Science & Technology

1981-12-01

4~ .A*4 ~.Zwe SOUMVV Ch.&4 0IVC&TIN 0 e*9 066so. 4 evt’ o R..e High speed, high resolution motion pictures were taken to compare the cinematic data...propellant. High speed, high resolution motion pictures were taken to compare the cinematic data with that available from the holograms. TABLE OF...of the motor ignition system, the repeatability of the pressure-time trace, and the timing of the cinematic /holographic obser- vation. The current

Propellant production from the Martian atmosphere

NASA Technical Reports Server (NTRS)

Bowles, J. V.; Tauber, M. E.; Anagnost, A. J.; Whittaker, T.

1992-01-01

Results are presented from a calculation of the specific impulses that can be generated through the combustion of cryogenic CO and O2 over a range of fuel/oxidizer ratios, chamber pressures, nozzle expansion ratios, freestream pressures representative of Mars, and the limiting conditions of equilibrium and frozen nozzle flow. For an expansion ratio of 80 and 100-atm. chamber pressure, a specific impulse of 298 sec was obtained; this is comparable to the best solid rocket propellants.

Propeller flaps in eyelid reconstruction.

PubMed

Rajak, Saul N; Huilgol, Shyamala C; Murakami, Masahiro; Selva, Dinesh

2018-03-14

Propeller flaps are island flaps that reach the recipient site through an axial rotation. The flap has a subcutaneous pedicle on which it pivots, thereby resembling a helicopter propeller. We present our series of propeller flaps for the reconstruction of large eyelid defects. This is a retrospective review of the clinical case notes of eight patients that underwent tumour excision with reconstruction with a cutaneous propeller flap supplied by a non-perforator orbicularis pedicle between July and December 2016. Propeller flaps were used in the reconstruction of five lower lid defects (size range 19 × 5 mm to 25 × 8 mm), one medial canthus defect (13 mm diameter), one complete upper lid defect (42 × 19 mm diameter) and one lid sparing extenteration defect. The flaps were recruited from nasolabial, lateral canthal, temple or medial upper cheek skin. Post-operatively one case had 'trapdooring' which required flap revision at 4 months and one had persistent oedema that settled without intervention. The reconstruction of large eyelid defects is challenging in part because of the paucity of locally available skin. Propeller flaps are a paradigm shift in periocular reconstruction in which the subcutaneous pedicle enables the recruitment of large and highly mobile skin flaps from a wide area of regional tissue.

Ozone depletion caused by NO and H2O emissions from hydrazine-fueled rockets

NASA Astrophysics Data System (ADS)

Ross, M. N.; Danilin, M. Y.; Weisenstein, D. K.; Ko, M. K. W.

2004-11-01

Rockets using unsymmetrical dimethyl hydrazine (N(CH3)2NH2) and dinitrogen tetroxide (N2O4) propellants account for about one third of all stratospheric rocket engine emissions, comparable to the solid-fueled rocket emissions. We use plume and global atmosphere models to provide the first estimate of the local and global ozone depletion caused by NO and H2O emissions from the Proton rocket, the largest hydrazine-fueled launcher in use. NO and H2O emission indices are assumed to be 20 and 350 g/kg (propellant), respectively. Predicted maximum ozone loss in the plume of the Proton rocket is 21% at 44 km altitude. Plume ozone loss at 20 km equals 8% just after launch and steadily declines to 2% by model sunset. Predicted steady state global ozone loss from ten Proton launches annually is 1.2 × 10-4%, with nearly all of the loss due to the NO component of the emission. Normalized by stratospheric propellant consumption, the global ozone depletion efficiency of the Proton is approximately 66-90 times less than that of solid-fueled rockets. In situ Proton plume measurements are required to validate assumed emission indices and to assess the role of rocket emissions not considered in these calculations. Such future studies would help to establish a formalism to evaluate the relative ozone depletion caused by different rocket engines using different propellants.

Induction simulation of gas core nuclear engine

NASA Technical Reports Server (NTRS)

Poole, J. W.; Vogel, C. E.

1973-01-01

The design, construction and operation of an induction heated plasma device known as a combined principles simulator is discussed. This device incorporates the major design features of the gas core nuclear rocket engine such as solid feed, propellant seeding, propellant injection through the walls, and a transpiration cooled, choked flow nozzle. Both argon and nitrogen were used as propellant simulating material, and sodium was used for fuel simulating material. In addition, a number of experiments were conducted utilizing depleted uranium as the fuel. The test program revealed that satisfactory operation of this device can be accomplished over a range of operating conditions and provided additional data to confirm the validity of the gas core concept.

NASA Researcher Adjusts a Travelling Magnetic Wave Plasma Engine

NASA Image and Video Library

1964-02-21

Raymond Palmer, of the Electromagnetic Propulsion Division’s Plasma Flow Section, adjusts the traveling magnetic wave plasma engine being operated in the Electric Power Conversion at the National Aeronautics and Space Administration (NASA) Lewis Research Center. During the 1960s Lewis researchers were exploring several different methods of creating electric propulsion systems, including the traveling magnetic wave plasma engine. The device operated similarly to alternating-current motors, except that a gas, not a solid, was used to conduct the electricity. A magnetic wave induced a current as it passed through the plasma. The current and magnetic field pushed the plasma in one direction. Palmer and colleague Robert Jones explored a variety of engine configurations in the Electric Propulsion Research Building. The engine is seen here mounted externally on the facility’s 5-foot diameter and 16-foot long vacuum tank. The four magnetic coils are seen on the left end of the engine. The researchers conducted two-minute test runs with varying configurations and used of both argon and xenon as the propellant. The Electric Propulsion Research Building was built in 1942 as the Engine Propeller Research Building, often called the Prop House. It contained four test cells to study large reciprocating engines with their propellers. After World War II, the facility was modified to study turbojet engines. By the 1960s, the facility was modified again for electric propulsion research and given its current name.

Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

NASA Technical Reports Server (NTRS)

Martin, Heath Thomas

2013-01-01

Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

Study of solid rocket motor for a space shuttle booster. Appendix A: SRM water entry loads

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the water entry loads imposed on the reusable solid propellant rocket engine of the space shuttle following parachute descent is presented. The cases discussed are vertical motion, horizontal motion, and motion after penetration. Mathematical models, diagrams, and charts are included to support the theoretical considerations.

Study of solid rocket motor for space shuttle booster. Volume 4: Cost

NASA Technical Reports Server (NTRS)

1972-01-01

The cost data for solid propellant rocket engines for use with the space shuttle are presented. The data are based on the selected 156 inch parallel and series burn configurations. Summary cost data are provided for the production of the 120 inch and 260 inch configurations. Graphs depicting parametric cost estimating relationships are included.

SUB-PPB QUANTITATION AND CONFIRMATION OF PERCHLORATE IN DRINKING WATERS CONTAINING HIGH TOTAL DISSOLVED SOLIDS USING ION CHROMATOGRAPHY WITH MASS SPECTROMETRIC DETECTION

EPA Science Inventory

Perchlorate (ClO4 -) is a drinking water contaminant originating from the dissolution of the salts of ammonium, potassium, magnesium, or sodium in water. It is used primarily as an oxidant in solid propellant for rockets, missiles, pyrotechnics, as a component in air bag infla...

Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning

NASA Technical Reports Server (NTRS)

1972-01-01

The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

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.

20th JANNAF Propulsion Systems Hazards Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Cocchiaro, James E. (Editor); Eggleston, Debra S. (Editor); Gannaway, Mary T. (Editor); Inzar, Jeanette M. (Editor)

2002-01-01

This volume, the first of two volumes, is a collection of 24 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 20th Propulsion Systems Hazards Subcommittee (PSHS), 38th Combustion Subcommittee (CS), 26th Airbreathing Propulsion Subcommittee (APS), and 21 Modeling and Simulation Subcommittee meeting. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics covered include: insensitive munitions and hazard classification testing of solid rocket motors and other munitions; vulnerability of gun propellants to impact stimuli; thermal decomposition and cookoff properties of energetic materials; burn-to-violent reaction phenomena in energetic materials; and shock-to-detonation properties of solid propellants and energetic materials.

Economics of the solid rocket booster for space shuttle

NASA Technical Reports Server (NTRS)

Rice, W. C.

1979-01-01

The paper examines economics of the solid rocket booster for the Space Shuttle. Costs have been held down by adapting existing technology to the 146 in. SRB selected, with NASA reducing the cost of expendables and reusing the expensive nonexpendable hardware. Drop tests of Titan III motor cases and nozzles proved that boosters can survive water impact at vertical velocities of 100 ft/sec so that SRB components can be reused. The cost of expendables was minimized by selecting proven propellants, insulation, and nozzle ablatives of known costs; the propellant has the lowest available cost formulation, and low cost ablatives, such as pitch carbon fibers, will be used when available. Thus, the use of proven technology and low cost expendables will make the SRB an economical booster for the Space Shuttle.

Pulse Phase Dynamic Thermal Tomography Investigation on the Defects of the Solid-Propellant Missile Engine Cladding Layer

NASA Astrophysics Data System (ADS)

Peng, Wei; Wang, Fei; Liu, Jun-yan; Xiao, Peng; Wang, Yang; Dai, Jing-min

2018-04-01

Pulse phase dynamic thermal tomography (PP-DTT) was introduced as a nondestructive inspection technique to detect the defects of the solid-propellant missile engine cladding layer. One-dimensional thermal wave mathematical model stimulated by pulse signal was developed and employed to investigate the thermal wave transmission characteristics. The pulse phase algorithm was used to extract the thermal wave characteristic of thermal radiation. Depth calibration curve was obtained by fuzzy c-means algorithm. Moreover, PP-DTT, a depth-resolved photothermal imaging modality, was employed to enable three-dimensional (3D) visualization of cladding layer defects. The comparison experiment between PP-DTT and classical dynamic thermal tomography was investigated. The results showed that PP-DTT can reconstruct the 3D topography of defects in a high quality.

Solid rocket booster performance evaluation model. Volume 3: Sample case. [propellant combustion simulation/internal ballistics

NASA Technical Reports Server (NTRS)

1974-01-01

The solid rocket booster performance evaluation model (SRB-11) is used to predict internal ballistics in a sample motor. This motor contains a five segmented grain. The first segment has a 14 pointed star configuration with a web which wraps partially around the forward dome. The other segments are circular in cross-section and are tapered along the interior burning surface. Two of the segments are inhibited on the forward face. The nozzle is not assumed to be submerged. The performance prediction is broken into two simulation parts: the delivered end item specific impulse and the propellant properties which are required as inputs for the internal ballistics module are determined; and the internal ballistics for the entire burn duration of the motor are simulated.

The PROPEL Electrodynamic Tether Mission and Connecting to the Ionosphere

NASA Technical Reports Server (NTRS)

Gilchrist, Brian; Bilen, Sven; Hoyt, Rob; Stone,Nobie; Vaughn, Jason; Fuhrhop, Keith; Krause, Linda; Khazanov, George; Johnson, Les

2012-01-01

The exponential increase of launch system size.and cost.with delta-V makes missions that require large total impulse cost prohibitive. Led by NASA's Marshall Space Flight Center, a team from government, industry, and academia has developed a flight demonstration mission concept of an integrated electrodynamic (ED) tethered satellite system called PROPEL: "Propulsion using Electrodynamics". The PROPEL Mission is focused on demonstrating a versatile configuration of an ED tether to overcome the limitations of the rocket equation, enable new classes of missions currently unaffordable or infeasible, and significantly advance the Technology Readiness Level (TRL) to an operational level. We are also focused on establishing a far deeper understanding of critical processes and technologies to be able to scale and improve tether systems in the future. Here, we provide an overview of the proposed PROPEL mission. One of the critical processes for efficient ED tether operation is the ability to inject current to and collect current from the ionosphere. Because the PROPEL mission is planned to have both boost and deboost capability using a single tether, the tether current must be capable of flowing in both directions and at levels well over 1 A. Given the greater mobility of electrons over that of ions, this generally requires that both ends of the ED tether system can both collect and emit electrons. For example, hollow cathode plasma contactors (HCPCs) generally are viewed as state-of-the-art and high TRL devices; however, for ED tether applications important questions remain of how efficiently they can operate as both electron collectors and emitters. Other technologies will be highlighted that are being investigated as possible alternatives to the HCPC such as Solex that generates a plasma cloud from a solid material (Teflon) and electron emission (only) technologies such as cold-cathode electron field emission or photo-electron beam generation (PEBG) techniques.

Three-Dimensional Simulation of Base Bleed Unit with AP/HTPB Propellant in Fast Cook-off Conditions

NASA Astrophysics Data System (ADS)

Li, Wen-feng; Yu, Yong-gang; Ye, Rui; Yang, Hou-wen

2017-07-01

In this work, a three-dimensional unsteady heat transfer model of base bleed unit with trilobite ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) composite solid propellant is presented to analyze the cook-off characteristics. According to the two-step chemical reaction of AP/HTPB propellant, a small-scale cook-off test is established. A comparison of the experimental and calculated results is made to verify the rationality of the computation model. On this basis, a cook-off numerical simulation of the base bleed unit at the heating rates of 0.33, 0.58 and 0.83 K/s is presented to investigate the ignition and initiation characteristics. The results show that the ignitions occur on the head face of the AP/HTPB propellant and near the internal gas chamber in these conditions. As the heating rate increases, the runaway time decreases and the ignition temperature rises.

Erosive burning research. [for solid-propellant rocket engines

NASA Technical Reports Server (NTRS)

Strand, L.; Yang, L. C.; Nguyen, M. H.; Cohen, N. S.

1986-01-01

A status report is given on the results for the completed tests in a series of motor firings being carried out to measure the effects of the parameters that are considered to most strongly influence the scaling to larger rocket motor sizes of the transition to/or threshold conditions for erosive burning rate augmentation. Propellant burning rates at locations along the axis of the test motors are measured with a newly developed plasma capacitance gauge technique. The measured results are compared with erosive-burning predictions from a supporting ballistics analysis. The completed motor firings have successfully demonstrated response to the designed test variables. The trends with varying propellant burning rate, chamber pressure, and mass flow rate are consistent with existing results, but no pronounced effect of surface roughness has been observed. Rather, the influence of propellant oxidizer particle size on erosive burning is through its effect on the base, no-corssflow burning rate.

JANNAF 37th Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

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

2000-01-01

This volume, the first of two volumes is a compilation of 59 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 37th Combustion Subcommittee (CS) meeting held jointly with the 25th Airbreathing Propulsion Subcommittee (APS), 19th Propulsion Systems Hazards Subcommittee (PSHS), and 1st Modeling and Simulation Subcommittee (MSS) meetings. The meeting was held 13-17 November 2000 at the Naval Postgraduate School and Hyatt Regency Hotel, Monterey, California. Topics covered at the CS meeting include: a keynote address on the Future Combat Systems, and review of a new JANNAF Modeling and Simulation Subcommittee, and technical papers on gun propellant burning rate, gun tube erosion, advanced gun propulsion concepts, ETC guns, novel gun propellants; liquid, hybrid and novel propellant combustion; solid propellant combustion kinetics, GAP, ADN and RDX combustion, sandwich combustion, metal combustion, combustion instability, and motor combustion instability.

Preliminary System Analysis of In Situ Resource Utilization for Mars Human Exploration

NASA Technical Reports Server (NTRS)

Rapp, Donald; Andringa, Jason; Easter, Robert; Smith, Jeffrey H .; Wilson, Thomas; Clark, D. Larry; Payne, Kevin

2005-01-01

We carried out a system analysis of processes for utilization of Mars resources to support human exploration of Mars by production of propellants from indigenous resources. Seven ISRU processes were analyzed to determine mass. power and propellant storage volume requirements. The major elements of each process include C02 acquisition, chemical conversion, and storage of propellants. Based on a figure of merit (the ratio of the mass of propellants that must be brought from Earth in a non-ISRU mission to the mass of the ISRU system. tanks and feedstocks that must be brought from Earth for a ISRU mission) the most attractive process (by far); is one where indigenous Mars water is accessible and this is processed via Sabatier/Electrolysis to methane and oxygen. These processes are technically relatively mature. Other processes with positive leverage involve reverse water gas shift and solid oxide electrolysis.

Safety and Performance Advantages of Nitrous Oxide Fuel Blends (NOFBX) Propellants for Manned and Unmanned Spaceflight Applications

NASA Astrophysics Data System (ADS)

Taylor, R.

2012-01-01

Hydrazine, N2H4, is the current workhorse monopropellant in the spacecraft industry. Although widely used since the 1960's, hydrazine is highly toxic and its specific impulse (ISP) performance of ~230s is far lower than bipropellants and solid motors. NOFBX™ monopropellants were originally developed under NASA's Mars Advanced Technology program (2004-2007) for deep space Mars missions. This work focused on characterizing various Nitrous Oxide Fuel Blend (NOFB) monopropellants which exhibited many favorable attributes to include: (1) Mono-propulsion, (2) Isp > 320s, (3) Non-toxic constituents, (4) Non-toxic effluents, (5) Low Cost, (6) High Density Specific Impulse, (7) Non-cryogenic, (8) Wide Storable Temperature Range, (9) Deeply throttlable [between 5 - 100lbs], (10) Self Pressurizing, (11) Wide Range of materials compatibility, along with many, many other benefits. All rocket propellants carry with them a history or stigma associated with either the development or implementation of that propellant and NOFBX™ is no exception. This paper examines the benefits of NOFBX™ propellants while addressing or dispelling a number of critiques N2O based propellants acquired through the decades of rocket propellant testing.

Waste Management Options for Long-Duration Space Missions: When to Reject, Reuse, or Recycle

NASA Technical Reports Server (NTRS)

Linne, Diane L.; Palaszewski, Bryan A.; Gokoglu, Suleyman; Gallo, Christopher A.; Balasubramaniam, Ramaswamy; Hegde, Uday G.

2014-01-01

The amount of waste generated on long-duration space missions away from Earth orbit creates the daunting challenge of how to manage the waste through reuse, rejection, or recycle. The option to merely dispose of the solid waste through an airlock to space was studied for both Earth-moon libration point missions and crewed Mars missions. Although the unique dynamic characteristics of an orbit around L2 might allow some discarded waste to intersect the lunar surface before re-impacting the spacecraft, the large amount of waste needed to be managed and potential hazards associated with volatiles recondensing on the spacecraft surfaces make this option problematic. A second option evaluated is to process the waste into useful gases to be either vented to space or used in various propulsion systems. These propellants could then be used to provide the yearly station-keeping needs at an L2 orbit, or if processed into oxygen and methane propellants, could be used to augment science exploration by enabling lunar mini landers to the far side of the moon.

Retro Rocket Motor Self-Penetrating Scheme for Heat Shield Exhaust Ports

NASA Technical Reports Server (NTRS)

Marrese-Reading, Colleen; St.Vaughn, Josh; Zell, Peter; Hamm, Ken; Corliss, Jim; Gayle, Steve; Pain, Rob; Rooney, Dan; Ramos, Amadi; Lewis, Doug;

Experimental Studies of Liquefaction and Densification of Liquid Oxygen

NASA Technical Reports Server (NTRS)

Partridge, Jonathan Koert

2010-01-01

The propellant combination that offers optimum performance is very reactive with a low average molecular weight of the resulting combustion products. Propellant combinations such as oxygen and hydrogen meet the above criteria, however, the propellants in gaseous form require large propellant tanks due to the low density of gas. Thus, rocketry employs cryogenic refrigeration to provide a more dense propellant stored as a liquid. In addition to propellant liquefaction, cryogenic refrigeration can also conserve propellant and provide propellant subcooling and propellant densification. Previous studies analyzed vapor conditioning of a cryogenic propellant, with the vapor conditioning by either a heat exchanger position in the vapor or by using the vapor in a refrigeration cycle as the working fluid. This study analyzes the effects of refrigeration heat exchanger located in the liquid of the common propellant oxidizer, liquid oxygen. This study predicted and determined the mass condensation rate and heat transfer coefficient for liquid oxygen.

Spontaneous mode-selection in the self-propelled motion of a solid/liquid composite driven by interfacial instability

NASA Astrophysics Data System (ADS)

Takabatake, Fumi; Magome, Nobuyuki; Ichikawa, Masatoshi; Yoshikawa, Kenichi

2011-03-01

Spontaneous motion of a solid/liquid composite induced by a chemical Marangoni effect, where an oil droplet attached to a solid soap is placed on a water phase, was investigated. The composite exhibits various characteristic motions, such as revolution (orbital motion) and translational motion. The results showed that the mode of this spontaneous motion switches with a change in the size of the solid scrap. The essential features of this mode-switching were reproduced by ordinary differential equations by considering nonlinear friction with proper symmetry.

Development of the Astrobee F sounding rocket system.

NASA Technical Reports Server (NTRS)

Jenkins, R. B.; Taylor, J. P.; Honecker, H. J., Jr.

1973-01-01

The development of the Astrobee F sounding rocket vehicle through the first flight test at NASA-Wallops Station is described. Design and development of a 15 in. diameter, dual thrust, solid propellant motor demonstrating several new technology features provided the basis for the flight vehicle. The 'F' motor test program described demonstrated the following advanced propulsion technology: tandem dual grain configuration, low burning rate HTPB case-bonded propellant, and molded plastic nozzle. The resultant motor integrated into a flight vehicle was successfully flown with extensive diagnostic instrumentation.-

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

NASA Astrophysics Data System (ADS)

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

1989-06-01

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

Experimental and Analytical Study of Erosive Burning of Solid Propellants

DTIC Science & Technology

1981-06-01

Identity by block number) Experirnert~iI - d analytical, *bdeling studies of the erosive burning ,..solfd propel l;n!t; w’r(, ,conducted at Atlantic Research...is approved for public ’release IAVV AFR 190-12 (Tb). Distribuiiou is unlitited. A. D . HLOSE Z Tecuhtgal Ina’o ’nation Offo icer - 3. Conduct...roughness. 8. Extend the erosive burning model from flat-plate geometry to axisymmetric flow. 9. Validate the 2- D model of erosive burning by experimental

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.

Theory, Solution Methods, and Implementation of the HERMES Model

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

Reaugh, John E.; White, Bradley W.; Curtis, John P.

The HERMES (high explosive response to mechanical stimulus) model was developed over the past decade to enable computer simulation of the mechanical and subsequent energetic response of explosives and propellants to mechanical insults such as impacts, perforations, drops, and falls. The model is embedded in computer simulation programs that solve the non-linear, large deformation equations of compressible solid and fluid flow in space and time. It is implemented as a user-defined model, which returns the updated stress tensor and composition that result from the simulation supplied strain tensor change. Although it is multi-phase, in that gas and solid species aremore » present, it is single-velocity, in that the gas does not flow through the porous solid. More than 70 time-dependent variables are made available for additional analyses and plotting. The model encompasses a broad range of possible responses: mechanical damage with no energetic response, and a continuous spectrum of degrees of violence including delayed and prompt detonation. This paper describes the basic workings of the model.« less

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.

Fuel-Cell Power Source Based on Onboard Rocket Propellants

NASA Technical Reports Server (NTRS)

Ganapathi, Gani; Narayan, Sri

2010-01-01

The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

Supplier's Status for Critical Solid Propellants, Explosive, and Pyrotechnic Ingredients

NASA Technical Reports Server (NTRS)

Sims, B. L.; Painter, C. R.; Nauflett, G. W.; Cramer, R. J.; Mulder, E. J.

2000-01-01

In the early 1970's a program was initiated at the Naval Surface Warfare Center/Indian Head Division (NSWC/IHDIV) to address the well-known problems associated with availability and suppliers of critical ingredients. These critical ingredients are necessary for preparation of solid propellants and explosives manufactured by the Navy. The objective of the program was to identify primary and secondary (or back-up) vendor information for these critical ingredients, and to develop suitable alternative materials if an ingredient is unavailable. In 1992 NSWC/IHDIV funded Chemical Propulsion Information Agency (CPIA) under a Technical Area Task (TAT) to expedite the task of creating a database listing critical ingredients used to manufacture Navy propellant and explosives based on known formulation quantities. Under this task CPIA provided employees that were 100 percent dedicated to the task of obtaining critical ingredient suppliers information, selecting the software and designing the interface between the computer program and the database users. TAT objectives included creating the Explosive Ingredients Source Database (EISD) for Propellant, Explosive and Pyrotechnic (PEP) critical elements. The goal was to create a readily accessible database, to provide users a quick-view summary of critical ingredient supplier's information and create a centralized archive that CPIA would update and distribute. EISD funding ended in 1996. At that time, the database entries included 53 formulations and 108 critical used to manufacture Navy propellant and explosives. CPIA turned the database tasking back over to NSWC/IHDIV to maintain and distribute at their discretion. Due to significant interest in propellant/explosives critical ingredients suppliers' status, the Propellant Development and Characterization Subcommittee (PDCS) approached the JANNAF Executive committee (EC) for authorization to continue the critical ingredient database work. In 1999, JANNAF EC approved the PDCS panel task. This paper is designed to emphasize the necessity of maintaining a JANNAF community supported database, which monitors PEP critical ingredient suppliers' status. The final product of this task is a user friendly, searchable database that provides a quick-view summary of critical ingredient supplier's information. This database must be designed to serve the needs of JANNAF and the propellant and energetic commercial manufacturing community as well. This paper provides a summary of the type of information to archive each critical ingredient.

A Design Tool for Matching UAV Propeller and Power Plant Performance

NASA Astrophysics Data System (ADS)

Mangio, Arion L.

A large body of knowledge is available for matching propellers to engines for large propeller driven aircraft. Small UAV's and model airplanes operate at much lower Reynolds numbers and use fixed pitch propellers so the information for large aircraft is not directly applicable. A design tool is needed that takes into account Reynolds number effects, allows for gear reduction, and the selection of a propeller optimized for the airframe. The tool developed in this thesis does this using propeller performance data generated from vortex theory or wind tunnel experiments and combines that data with an engine power curve. The thrust, steady state power, RPM, and tip Mach number vs. velocity curves are generated. The Reynolds number vs. non dimensional radial station at an operating point is also found. The tool is then used to design a geared power plant for the SAE Aero Design competition. To measure the power plant performance, a purpose built engine test stand was built. The characteristics of the engine test stand are also presented. The engine test stand was then used to characterize the geared power plant. The power plant uses a 26x16 propeller, 100/13 gear ratio, and an LRP 0.30 cubic inch engine turning at 28,000 RPM and producing 2.2 HP. Lastly, the measured power plant performance is presented. An important result is that 17 lbf of static thrust is produced.

Solid rocket motor internal insulation

NASA Technical Reports Server (NTRS)

Twichell, S. E. (Editor); Keller, R. B., Jr.

1976-01-01

Internal insulation in a solid rocket motor is defined as a layer of heat barrier material placed between the internal surface of the case propellant. The primary purpose is to prevent the case from reaching temperatures that endanger its structural integrity. Secondary functions of the insulation are listed and guidelines for avoiding critical problems in the development of internal insulation for rocket motors are presented.

Space Shuttle Projects

NASA Image and Video Library

1977-01-01

This illustration is a cutaway of the solid rocket booster (SRB) sections with callouts. 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. 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.

Mixing and combustion enhancement of Turbocharged Solid Propellant Ramjet

NASA Astrophysics Data System (ADS)

Liu, Shichang; Li, Jiang; Zhu, Gen; Wang, Wei; Liu, Yang

2018-02-01

Turbocharged Solid Propellant Ramjet is a new concept engine that combines the advantages of both solid rocket ramjet and Air Turbo Rocket, with a wide operation envelope and high performance. There are three streams of the air, turbine-driving gas and augment gas to mix and combust in the afterburner, and the coaxial intake mode of the afterburner is disadvantageous to the mixing and combustion. Therefore, it is necessary to carry out mixing and combustion enhancement research. In this study, the numerical model of Turbocharged Solid Propellant Ramjet three-dimensional combustion flow field is established, and the numerical simulation of the mixing and combustion enhancement scheme is conducted from the aspects of head region intake mode to injection method in afterburner. The results show that by driving the compressed air to deflect inward and the turbine-driving gas to maintain strong rotation, radial and tangential momentum exchange of the two streams can be enhanced, thereby improving the efficiency of mixing and combustion in the afterburner. The method of injecting augment gas in the transverse direction and making sure the injection location is as close as possible to the head region is beneficial to improve the combustion efficiency. The outer combustion flow field of the afterburner is an oxidizer-rich environment, while the inner is a fuel-rich environment. To improve the efficiency of mixing and combustion, it is necessary to control the injection velocity of the augment gas to keep it in the oxygen-rich zone of the outer region. The numerical simulation for different flight conditions shows that the optimal mixing and combustion enhancement scheme can obtain high combustion efficiency and have excellent applicability in a wide working range.

Preparation and Structure Study of Water-Blown Polyurethane/RDX Gun Propellant Foams

NASA Astrophysics Data System (ADS)

Yang, Weitao; Yang, Jianxing; Zhao, Yuhua; Zhang, Yucheng

2018-01-01

Water-blown polyurethane/RDX foamed propellants were prepared using polyols and isocyanate as reactive binder system, hexogen (RDX) as energetic component, triethanolamine (TEA)/Ditin butyl dilaurate (T-12) as composite catalysts, and H2O as blowing agent. The influences of catalyst ratio, blowing agent amount, and solid filler content on the inner porous structure were studied. The results show that the balance of gel rate and cream rate that could be adjusted by catalyst ratio is a major influencing factor on porous structure of foamed propellants. When the ratio of TEA/T-12 was adjusted to 1/0.7, the morphology of the foamed propellant exhibited spherical and closed porous structure. Besides, when the water amount was increased from 0.1% to 0.5%, the pore size increased from 0.43 to 0.64 mm. The contents of RDX particles affected the cell nucleation and thus, the cell geometry. When the blowing agent amount was constant, the increased content of RDX filler led to a decreased pore size. The closed bomb test results showed that foamed propellants burned progressively in an in-depth combustion mode.

Performance of an iodine-fueled radio-frequency ion-thruster

NASA Astrophysics Data System (ADS)

Holste, Kristof; Gärtner, Waldemar; Zschätzsch, Daniel; Scharmann, Steffen; Köhler, Peter; Dietz, Patrick; Klar, Peter J.

2018-01-01

Two sets of performance data of the same radio-frequency ion-thruster (RIT) have been recorded using iodine and xenon, respectively, as propellant. To characterize the thruster's performance, we have recorded the radio-frequency DC-power, required for yielding preset values of the extracted ion-beam currents, as a function of mass flow. For that purpose, an iodine mass flow system had to be developed, calibrated, and integrated into a newly-built test facility for studying corrosive propellants. The performance mappings for iodine and xenon differ significantly despite comparable operation conditions. At low mass flows, iodine exhibits the better performance. The situation changes at higher mass flows where the performance of iodine is significantly poorer than that of xenon. The reason is very likely related to the molecular nature of iodine. Our results show that iodine as propellant is compatible with RIT technology. Furthermore, it is a viable alternative as propellant for dedicated space missions. In particular, when taking into account additional benefits such as possible storage as a solid and its low price the use of iodine as propellant in ion thrusters is competitive.

Synthesis of ethylene diamine-based ferrocene terminated dendrimers and their application as burning rate catalysts.

PubMed

Zain-Ul-Abdin; Wang, Li; Yu, Haojie; Saleem, Muhammad; Akram, Muhammad; Khalid, Hamad; Abbasi, Nasir M; Yang, Xianpeng

2017-02-01

Ferrocene-based derivatives are widely used as ferrocene-based burning rate catalysts (BRCs) for ammonium perchlorate (AP)-based propellant. However, in long storage, small ferrocene-based derivatives migrate to the surface of the propellant, which results in changes in the designed burning parameters and finally causes unstable combustion. To retard the migration of ferrocene-based BRCs in the propellant and to increase the combustion of the solid propellant, zero to third generation ethylene diamine-based ferrocene terminated dendrimers (0G, 1G, 2G and 3G) were synthesized. The synthesis of these dendrimers was confirmed by 1 H NMR and FT-IR spectroscopy. The electrochemical behavior of 0G, 1G, 2G and 3G was investigated by cyclic voltammetry (CV) and the burning rate catalytic activity of 0G, 1G, 2G and 3G on thermal disintegration of AP was examined by thermogravimetry (TG) and differential thermogravimetry (DTG) techniques. Anti-migration studies show that 1G, 2G and 3G exhibit improved anti-migration behavior in the AP-based propellant. Copyright © 2016 Elsevier Inc. All rights reserved.

An improved model for the combustion of AP composite propellants

NASA Technical Reports Server (NTRS)

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

1981-01-01

This paper presents several improvements to the BDP model of steady-state burning of AP composite solid propellants. The Price-Boggs-Derr model of AP monopropellant burning is incorporated to represent the AP. A separate energy equation is written for the binder to permit a different surface temperature from the AP; this includes an analysis of the sharing of primary diffusion flame energy, and correction of a BDP model inconsistency in treating the binder regression rate. A method for assembling component contributions to calculate the burning rates of multimodal propellants is also presented. Results are shown in the form of representative burning rate curves, comparisons with data, and calculated internal details of interest. Ideas for future work are discussed in an Appendix.

An Investigation of Experimental Techniques for Obtaining Particulate Behavior in Metallized Solid Propellant Combustion.

DTIC Science & Technology

1984-02-01

Additive Particle Size Propellant Binder Oxidizer Metal Mean Metal Diameter, Designation % Weight % Weight* % Weight Microns WGS-5A HTPB 12 AP 83 AL 5 75...88 4 WGS-6A HTPB 12 AP 83 AL 5 45-62 WGS-7A HTPB 12 AP 83 AL 5 23-37 WGS-7 HTPB 12 AP 83 AL 5 6-7 WGS-9 HTPB 12 AP 78 AL 10 23-27 WGS-1O HTPB 12 AP 73...AL 15 23-27 WGS-ZrC HTPB 14 AP 84 ZrC 2 23, irregularly shaped WGS-G HTPB 14 AP 84 G 2 50x20x7, flakes * 65% 180 jim/35% 26 tim These propellants

The combustion of pure and composite propellants - The expansion and application of laminar flame theory to heterogeneous solid propellants.

NASA Technical Reports Server (NTRS)

Hertzberg, M.

1971-01-01

Development of a combustion theory based on the laminarized solutions to the energy and flow conservation equations, which is more realistic in recognizing the nature of the heating-rate problem and in obtaining a practical solution to estimating its magnitude. A new experimental approach is used for studying the combustion behavior of pure monopropellants and composite propellants which uses a laser beam to supply additional heat feedback to a burning surface. New experimental data are presented for the laser-induced combustion rate and ignition delay of pure ammonium perchlorate. The pure monopropellant theory is generalized to include such nonadiabatic effects, and the new experimental data are in good agreement with the nonadiabatic theory.-

Preliminary measurement of the noise from the 2/9 scale model of the Large-scale Advanced Propfan (LAP) propeller, SR-7A

NASA Technical Reports Server (NTRS)

Dittmar, J. H.

1985-01-01

Noise data on the Large-scale Advanced Propfan (LAP) propeller model SR-7A were taken into the NASA Lewis 8- by 6-Foot Wind Tunnel. The maximum blade passing tone decreases from the peak level when going to higher helical tip Mach numbers. This noise reduction points to the use of higher propeller speeds as a possible method to reduce airplane cabin noise while maintaining high flight speed and efficiency. Comparison of the SR-7A blade passing noise with the noise of the similarly designed SR-3 propeller shows good agreement as expected. The SR-7A propeller is slightly noisier than the SR-3 model in the plane of rotation at the cruise condition. Projections of the tunnel model data are made to the full-scale LAP propeller mounted on the test bed aircraft and compared with design predictions. The prediction method is conservative in the sense that it overpredicts the projected model data.

Development of a computerized analysis for solid propellant combustion instability with turbulence

NASA Technical Reports Server (NTRS)

Chung, T. J.; Park, O. Y.

1988-01-01

A multi-dimensional numerical model has been developed for the unsteady state oscillatory combustion of solid propellants subject to acoustic pressure disturbances. Including the gas phase unsteady effects, the assumption of uniform pressure across the flame zone, which has been conventionally used, is relaxed so that a higher frequency response in the long flame of a double-base propellant can be calculated. The formulation is based on a premixed, laminar flame with a one-step overall chemical reaction and the Arrhenius law of decomposition with no condensed phase reaction. In a given geometry, the Galerkin finite element solution shows the strong resonance and damping effect at the lower frequencies, similar to the result of Denison and Baum. Extended studies deal with the higher frequency region where the pressure varies in the flame thickness. The nonlinear system behavior is investigated by carrying out the second order expansion in wave amplitude when the acoustic pressure oscillations are finite in amplitude. Offset in the burning rate shows a negative sign in the whole frequency region considered, and it verifies the experimental results of Price. Finally, the velocity coupling in the two-dimensional model is discussed.

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.

Lessons Learned in Solid Rocket Combustion Instability

DTIC Science & Technology

2006-11-14

Gary Flandro . Also I wish to thank James Crump and H.B. Mathes who provided guidance during my first ten years at China Lake. VIII. References 1 L...It is also a form of analysis to examine the acoustic boundary with flow normal to the surface. It is sometimes known as the " Flandro boundary layer...Tso, "Flow Turning Losses in Solid Rocket Motors," AFAL-TR-87-095, March 1988. 30 G.A. Flandro , "Solid Propellant Acoustic Admittance Corrections

Launch Vehicle Performance for Bipropellant Propulsion Using Atomic Propellants With Oxygen

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2000-01-01

Atomic propellants for bipropellant launch vehicles using atomic boron, carbon, and hydrogen were analyzed. The gross liftoff weights (GLOW) and dry masses of the vehicles were estimated, and the 'best' design points for atomic propellants were identified. Engine performance was estimated for a wide range of oxidizer to fuel (O/F) ratios, atom loadings in the solid hydrogen particles, and amounts of helium carrier fluid. Rocket vehicle GLOW was minimized by operating at an O/F ratio of 1.0 to 3.0 for the atomic boron and carbon cases. For the atomic hydrogen cases, a minimum GLOW occurred when using the fuel as a monopropellant (O/F = 0.0). The atomic vehicle dry masses are also presented, and these data exhibit minimum values at the same or similar O/F ratios as those for the vehicle GLOW. A technology assessment of atomic propellants has shown that atomic boron and carbon rocket analyses are considered to be much more near term options than the atomic hydrogen rockets. The technology for storing atomic boron and carbon has shown significant progress, while atomic hydrogen is not able to be stored at the high densities needed for effective propulsion. The GLOW and dry mass data can be used to estimate the cost of future vehicles and their atomic propellant production facilities. The lower the propellant's mass, the lower the overall investment for the specially manufactured atomic propellants.

The Effect of Reduction Gearing on Propeller-body Interference as Shown by Full-Scale Wind-Tunnel Tests

NASA Technical Reports Server (NTRS)

Weick, Fred E

1931-01-01

This report presents the results of full-scale tests made on a 10-foot 5-inch propeller on a geared J-5 engine and also on a similar 8-foot 11-inch propeller on a direct-drive J-5 engine. Each propeller was tested at two different pitch settings, and with a large and a small fuselage. The investigation was made in such a manner that the propeller-body interference factors were isolated, and it was found that, considering this interference only, the geared propellers had an appreciable advantage in propulsive efficiency, partially due to the larger diameter of the propellers with respect to the bodies, and partially because the geared propellers were located farther ahead of the engines and bodies.

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

NMR imaging and hydrodynamic analysis of neutrally buoyant non-Newtonian slurry flows

NASA Astrophysics Data System (ADS)

Bouillard, J. X.; Sinton, S. W.

The flow of solids loaded suspension in cylindrical pipes has been the object of intense experimental and theoretical investigations in recent years. These types of flows are of great interest in chemical engineering because of their important use in many industrial manufacturing processes. Such flows are for example encountered in the manufacture of solid-rocket propellants, advanced ceramics, reinforced polymer composites, in heterogeneous catalytic reactors, and in the pipeline transport of liquid-solids suspensions. In most cases, the suspension microstructure and the degree of solids dispersion greatly affect the final performance of the manufactured product. For example, solid propellant pellets need to be extremely-well dispersed in gel matrices for use as rocket engine solid fuels. The homogeneity of pellet dispersion is critical to allow good uniformity of the burn rate, which in turn affects the final mechanical performance of the engine. Today's manufacturing of such fuels uses continuous flow processes rather than batch processes. Unfortunately, the hydrodynamics of such flow processes is poorly understood and is difficult to assess because it requires the simultaneous measurements of liquid/solids phase velocities and volume fractions. Due to the recent development in pulsed Fourier Transform NMR imaging, NMR imaging is now becoming a powerful technique for the non intrusive investigation of multi-phase flows. This paper reports and exposes a state-of-the-art experimental and theoretical methodology that can be used to study such flows. The hydrodynamic model developed for this study is a two-phase flow shear thinning model with standard constitutive fluid/solids interphase drag and solids compaction stresses. this model shows good agreement with experimental data and the limitations of this model are discussed.

First Stage Solid Propellant Multi Debris Thermal Analysis

NASA Technical Reports Server (NTRS)

Toleman, Benjamin M.

2011-01-01

The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

First Stage Solid Propellant Multiply Debris Thermal Analysis

NASA Technical Reports Server (NTRS)

Toleman, Benjamin M.

2011-01-01

Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

A method of calculating the performance of controllable propellers with sample computations

NASA Technical Reports Server (NTRS)

Hartman, Edwin P

1934-01-01

This paper contains a series of calculations showing how the performance of controllable propellers may be derived from data on fixed-pitch propellers given in N.A.C.A. Technical Report No. 350, or from similar data. Sample calculations are given which compare the performance of airplanes with fixed-pitch and with controllable propellers. The gain in performance with controllable propellers is shown to be largely due to the increased power available, rather than to an increase in efficiency. Controllable propellers are of particular advantage when used with geared and with supercharged engines. A controllable propeller reduces the take-off run, increases the rate of climb and the ceiling, but does not increase the high speed, except when operating above the design altitude of the previously used fixed-pitch propeller or when that propeller was designed for other than high speed.

Thermophysical Property Testing Using Transient Techniques.

DTIC Science & Technology

1984-06-29

WORDS (Continue on reverse side if necessary and identify by block number) Specific heat HMX carbon/carbon Diffusivity RDX solid propellants Conductivity...energetic materials (AP, " HMX , RDX and HTPB) used in solid rocket fuel to carbon/carbon materials used as rocket nozzles. Studies on AP included single...32 4.1b HMX and RDX ............................35 a 4.2 Carbon/Carbon Materials ...................... 36 5.0 SUMMARY

Study of solid rocket motor for space shuttle booster, volume 2, book 3, appendix A

NASA Technical Reports Server (NTRS)

1972-01-01

A systems requirements analysis for the solid propellant rocket engine to be used with the space shuttle was conducted. The systems analysis was developed to define the physical and functional requirements for the systems and subsystems. The operations analysis was performed to identify the requirements of the various launch operations, mission operations, ground operations, and logistic and flight support concepts.

Hybrid propulsion technology program. Volume 1: Conceptional design 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). As part of the design study critical technology issues were identified and a technology acquisition and demonstration plan was formulated.

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.

A investigation on unixial and quasi-biaxial tensile mechanical properties of aging HTPB propellant under dynamic loading at low temperature

NASA Astrophysics Data System (ADS)

Duan, Leiguang; Wang, Guang; Zhang, Guoxing; Sun, Xinya; Shang, Hehao

2018-06-01

In order to study the uniaxial and quasi-biaxial mechanical properties of aging solid propellants under low temperature and high strain rate, stress-strain curves and tensile fracture surfaces of HTPB propellant were obtained in a wide range of temperature (-30,25 °C) and strain rates (0.4,4.0 and 14.29 s-1), respectively, by means of uniaxial and biaxial tensile tests and electron microscopy scanning on the fracture cross section. The results indicate that the quasi-biaxial tensile mechanical properties of aging HTPB propellant is same as the uniaxial tensile mechanical properties influenced distinctly by temperature and strain rate. With decreasing temperature and increasing strain rate, the mechanical properties gradually strengthen. The damage for HTPB propellant changes from "dehumidification" to grain fracture. The initial elastic modulus E and maximum tensile stress σ of the uniaxial and biaxial tensile increase gradually with decreasing temperature and increasing strain rate, and well present linear-log function relation with strain rate. The ratio of quasi-biaxial and uniaxial stretching under different loading conditions was obtained so that the researchers could predict the quasi-biaxial tensile mechanical properties of the propellant based on the uniaxial test data.

Coal-Fired Rocket Engine

NASA Technical Reports Server (NTRS)

Anderson, Floyd A.

1987-01-01

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

LES of propeller crashback

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2014-11-01

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield a negative thrust. In crashback, the freestream interacts with the strong reverse flow from the propeller leading to massive flow separation and highly unsteady loads. We have used Large-Eddy Simulation (LES) in recent years to accurately simulate the flowfield in crashback around a stand-alone open propeller, hull-attached (posterior alone) open propeller and a ducted propeller with stator blades. This talk will discuss our work towards LES of crashback inclusive of the entire hull. The results will be compared to available experimental data, and the flow physics will be discussed. This work is supported by the Office of Naval Research.

Noise reduction for model counterrotation propeller at cruise by reducing aft-propeller diameter

NASA Technical Reports Server (NTRS)

Dittmar, James H.; Stang, David B.

1987-01-01

The forward propeller of a model counterrotation propeller was tested with its original aft propeller and with a reduced diameter aft propeller. Noise reductions with the reduced diameter aft propeller were measured at simulated cruise conditions. Reductions were as large as 7.5 dB for the aft-propeller passing tone and 15 dB in the harmonics at specific angles. The interaction tones, mostly the first, were reduced probably because the reduced-diameter aft-propeller blades no longer interacted with the forward propeller tip vortex. The total noise (sum of primary and interaction noise) at each harmonic was significantly reduced. The chief noise reduction at each harmonic came from reduced aft-propeller-alone noise, with the interaction tones contributing little to the totals at cruise. Total cruise noise reductions were as much as 3 dB at given angles for the blade passing tone and 10 dB for some of the harmonics. These reductions would measurably improve the fuselage interior noise levels and represent a definite cruise noise benefit from using a reduced diameter aft propeller.

Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rahman, Shamim A.

2010-01-01

Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.

Pulsed-Laser, High Speed Photography of Rocket Propellant Surface Deflagration.

DTIC Science & Technology

1986-05-01

Investigator was Dr Roger J. Becker. AFRPL Project Manager was Mr Gary L. Vogt. This technical report has been reviewed and is approved for publication...8217;YMlB)OI (/P’I I la . i tJ .o C ’ Gary L. Vogt (805) 277-5258 AFPLIDYCR DD FORM 1473,83 APR EDITION OF 1 JAN 73 IS OBSOLETE. Unclass i fied" SECURl iY...84-1236. 4. G. A. Flandro , "A Simple Conceptual Model for the Nonlinear Transient Combustion of a Solid Rocket Propellant," AIAA Paper No. 82-1222

Condensed-Phase Processes during Solid Propellant Combustion. Part 2: Chemical and Microscopic Examination of Conductively Quenched Samples of RDX, XM39, JA2, M30, and HMX-Binder Compositions

DTIC Science & Technology

1992-05-01

combustion of most of the propellants, with the possible exception of JA2; scanning electron microcope examination shows the existence of a liquid layer but... compounds are similar (Fifer et Sl. 1985; Hoffsommer, Glover, and Elban 1985), the relative Intensities In Table 2 should provide rough, order-of...top of the liquid layer. In addition, the HPLC chromatograms contained a number of very weak, unknown peaks apparently corresponding to compounds

AFRPL Graphite Performance Prediction Program. Improved Capability for the Design and Ablation Performance Prediction of Advanced Air Force Solid Propellant Rocket Nozzles

DTIC Science & Technology

1976-12-01

corrosive attack by both acids and alkali and, in addition, is provided with a special Dynel veil for protection against fluoride attack. 3.1.4...throat region, namely , the entrance, center, and exit. In addition, at each station, the diameters were determined at two angular positions 90° apart. The...characterization test matrix. 3.2.1.1 Rocket Motor Environments Rocket motor environments were based on three advanced MX propellants, namely , * XLDB * HTPB * PEG

SRM Internal Flow Tests and Computational Fluid Dynamic Analysis. Volume 3; Titan, ASRM, and Subscale Motor Analyses

NASA Technical Reports Server (NTRS)

1995-01-01

A computational fluid dynamics (CFD) analysis has been performed on the aft slot region of the Titan 4 Solid Rocket Motor Upgrade (SRMU). This analysis was performed in conjunction with MSFC structural modeling of the propellant grain to determine if the flow field induced stresses would adversely alter the propellant geometry to the extent of causing motor failure. The results of the coupled CFD/stress analysis have shown that there is a continual increase of flow field resistance at the aft slot due to the aft segment propellant grain being progressively moved radially toward the centerline of the motor port. This 'bootstrapping' effect between grain radial movement and internal flow resistance is conducive to causing a rapid motor failure.

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 role of solid phase heterogeneity on the low-pressure L-star instability of nonmetallized AP/PBAN propellants is explored. Four particle size distributions are employed in propellants that are otherwise identical. Over one hundred test firings were conducted in the 21/2 in. diameter L-star burner. Pressure time histories in the chamber and color movies of two firings constitute the raw data. An economical firing program was used which enables the interesting range of L-star values to be covered during a single firing (at a set mean pressure), through the variations in the depleting propellant volume. Time-independent combustion, Helmholtz mode, chuff mode, and the pressure-burst phenomena are revealed as the principal signatures. Of these, the Helmholtz mode is found to be the most ordered form of instability.

Low-Cost Propellant Launch to LEO from a Tethered Balloon - 'Propulsion Depots' Not 'Propellant Depots'

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.; Schneider, Evan G.; Vaughan, David A.; Hall, Jeffrey L.; Yu, Chi Yau

2011-01-01

As we have previously reported, it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond.

Mobile propeller dynamometer validation

NASA Astrophysics Data System (ADS)

Morris, Mason Wade

With growing interest in UAVs and OSU's interest in propeller performance and manufacturing, evaluating UAV propeller and propulsion system performance has become essential. In attempts to evaluate these propellers a mobile propeller dynamometer has been designed, built, and tested. The mobile dyno has been designed to be cost effective through the ability to load it into the back of a test vehicle to create simulated forward flight characteristics. This allows much larger propellers to be dynamically tested without the use of large and expensive wind tunnels. While evaluating the accuracy of the dyno, several improvements had to be made to get accurate results. The decisions made to design and improve the mobile propeller dyno will be discussed along with attempts to validate the dyno by comparing its results against known sources. Another large part of assuring the accuracy of the mobile dyno is determining if the test vehicle will influence the flow going into the propellers being tested. The flow into the propeller needs to be as smooth and uniform as possible. This is determined by characterizing the boundary layer and accelerated flow over the vehicle. This evaluation was accomplished with extensive vehicle aerodynamic measurements with the use of full-scale tests using a pitot-rake and the actual test vehicle. Additional tests were conducted in Oklahoma State University's low speed wind tunnel with a 1/8-scale model using qualitative flow visualization with smoke. Continuing research on the mobile dyno will be discussed, along with other potential uses for the dyno.

Inert Reassessment Document for Diethanolamine - CAS No. 111-42-2

EPA Pesticide Factsheets

Diethyl phthalate used is as a plasticizer in a wide variety of consumer products, including plastic packaging film, automotive components, toys, cosmetic formulations, toiletries, medical tubing, solid rocket propellants, and as a ingredient in aspirin

Combustion of metal agglomerates in a solid rocket core flow

NASA Astrophysics Data System (ADS)

Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

2013-12-01

The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

Calculation of the Rate of Combustion of a Metallized Composite Solid Propellant with Allowance for the Size Distribution of Agglomerates

NASA Astrophysics Data System (ADS)

Poryazov, V. A.; Krainov, A. Yu.

2016-05-01

A physicomathematical model of combustion of a metallized composite solid propellant based on ammonium perchlorate has been presented. The model takes account of the thermal effect of decomposition of a condensed phase (c phase), convection, diffusion, the exothermal chemical reaction in a gas phase, the heating and combustion of aluminum particles in the gas flow, and the velocity lag of the particles behind the gas. The influence of the granulometric composition of aluminum particles escaping from the combustion surface on the linear rate of combustion has been investigated. It has been shown that information not only on the kinetics of chemical reactions in the gas phase, but also on the granulometric composition of aluminum particles escaping from the surface of the c phase into the gas, is of importance for determination of the linear rate of combustion.

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

2002-01-01

Propellant densification has been identified as a critical technology in the development of single-stage-to-orbit reusable launch vehicles. Technology to create supercooled high-density liquid oxygen (LO2) and liquid hydrogen (LH2) is a key means to lowering launch vehicle costs. The densification of cryogenic propellants through subcooling allows 8 to 10 percent more propellant mass to be stored in a given unit volume, thereby improving the launch vehicle's overall performance. This allows for higher propellant mass fractions than would be possible with conventional normal boiling point cryogenic propellants, considering the normal boiling point of LO2 and LH2.

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.

An analysis of the orbital distribution of solid rocket motor slag

NASA Astrophysics Data System (ADS)

Horstman, Matthew F.; Mulrooney, Mark

2009-01-01

The contribution by solid rocket motors (SRMs) to the orbital debris environment is potentially significant and insufficiently studied. Design and combustion processes can lead to the emission of enough by-products to warrant assessment of their contribution to orbital debris. These particles are formed during SRM tail-off, or burn termination, by the rapid solidification of molten Al2O3 slag accumulated during the burn. The propensity of SRMs to generate particles larger than 100μm raises concerns regarding the debris environment. Sizes as large as 1 cm have been witnessed in ground tests, and comparable sizes have been estimated via observations of sub-orbital tail-off events. Utilizing previous research we have developed more sophisticated size distributions and modeled the time evolution of resultant orbital populations using a historical database of SRM launches, propellant, and likely location and time of tail-off. This analysis indicates that SRM ejecta is a significant component of the debris environment.

Internal Flow Analysis of Large L/D Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Laubacher, Brian A.

2000-01-01

Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and more critically, motors with port to throat ratios near one, are covered.

AFOSR/ONR Contractors Meeting - Combustion, Rocket Propulsion, Diagnostics of Reacting Flow

DTIC Science & Technology

1990-06-15

GASIFICATION KINETICS OF SOLID BORON AND PYROLITIC GRAPHrrE Because of the energetic potential of boron as a solid fuel (or fuel additive) and the likely...87 Kinetic Studies of Metal Combustion in Propulsion, A. Fontijn, P. M. Futerko and A. G. Slavejkov .............................. 90...Measurements and Chemical Kinetic Simulation of the Structure of Model Propellant Flames, M. C. Branch and H. Dindi .......... 94 High-Rate Thermal

Workshop Report: Fundamental Reactions in Solid Propellant Combustion

DTIC Science & Technology

1979-05-01

combustion conditions. 6. What effect might a pressure-induced phase transition to a polymorph other than 6- HMX have on the pressure slope break during...pure HMX as well. Nevertheless, it is recommended that the high pressure polymorphs of HMX and RDX be determined. It was also felt that there...plateau burning phenomena E. Solid phase, surface, gas phase reactions F. Phase transitions : melting, vaporization, polymorphs G. Flame

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

NASA Technical Reports Server (NTRS)

Vonderesch, A. H.

1972-01-01

The factors affecting the choice of the 156 inch diameter, parallel burn, solid propellant rocket engine for use with the space shuttle booster are presented. Primary considerations leading to the selection are: (1) low booster vehicle cost, (2) the largest proven transportable system, (3) a demonstrated design, (4) recovery/reuse is feasible, (5) abort can be easily accomplished, and (6) ecological effects are minor.

Enhanced Thermal Decomposition Properties of CL-20 through Space-Confining in Three-Dimensional Hierarchically Ordered Porous Carbon.

PubMed

Chen, Jin; He, Simin; Huang, Bing; Wu, Peng; Qiao, Zhiqiang; Wang, Jun; Zhang, Liyuan; Yang, Guangcheng; Huang, Hui

2017-03-29

High energy and low signature properties are the future trend of solid propellant development. As a new and promising oxidizer, hexanitrohexaazaisowurtzitane (CL-20) is expected to replace the conventional oxidizer ammonium perchlorate to reach above goals. However, the high pressure exponent of CL-20 hinders its application in solid propellants so that the development of effective catalysts to improve the thermal decomposition properties of CL-20 still remains challenging. Here, 3D hierarchically ordered porous carbon (3D HOPC) is presented as a catalyst for the thermal decomposition of CL-20 via synthesizing a series of nanostructured CL-20/HOPC composites. In these nanocomposites, CL-20 is homogeneously space-confined into the 3D HOPC scaffold as nanocrystals 9.2-26.5 nm in diameter. The effect of the pore textural parameters and surface modification of 3D HOPC as well as CL-20 loading amount on the thermal decomposition of CL-20 is discussed. A significant improvement of the thermal decomposition properties of CL-20 is achieved with remarkable decrease in decomposition peak temperature (from 247.0 to 174.8 °C) and activation energy (from 165.5 to 115.3 kJ/mol). The exceptional performance of 3D HOPC could be attributed to its well-connected 3D hierarchically ordered porous structure, high surface area, and the confined CL-20 nanocrystals. This work clearly demonstrates that 3D HOPC is a superior catalyst for CL-20 thermal decomposition and opens new potential for further applications of CL-20 in solid propellants.

Solid rocket motors for the Space Shuttle booster.

NASA Technical Reports Server (NTRS)

Odom, J. B.

1972-01-01

The evolution of the space shuttle booster system is reviewed from its initial concepts based on liquid-propellant reusable boosters to the final selection of recoverable, solid-fuel rocket motors. The rationale associated with each of the several major decisions in the evolution process is discussed. It is shown that the external tank orbiter configuration emerging from the latest studies takes maximum advantage of the solid rocket motor development experience and promises to be the optimum configuration for fulfilling the paramount shuttle program requirements of minimum total development risk within acceptable costs.

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.

Random sphere packing model of heterogeneous propellants

NASA Astrophysics Data System (ADS)

Kochevets, Sergei Victorovich

It is well recognized that combustion of heterogeneous propellants is strongly dependent on the propellant morphology. Recent developments in computing systems make it possible to start three-dimensional modeling of heterogeneous propellant combustion. A key component of such large scale computations is a realistic model of industrial propellants which retains the true morphology---a goal never achieved before. The research presented develops the Random Sphere Packing Model of heterogeneous propellants and generates numerical samples of actual industrial propellants. This is done by developing a sphere packing algorithm which randomly packs a large number of spheres with a polydisperse size distribution within a rectangular domain. First, the packing code is developed, optimized for performance, and parallelized using the OpenMP shared memory architecture. Second, the morphology and packing fraction of two simple cases of unimodal and bimodal packs are investigated computationally and analytically. It is shown that both the Loose Random Packing and Dense Random Packing limits are not well defined and the growth rate of the spheres is identified as the key parameter controlling the efficiency of the packing. For a properly chosen growth rate, computational results are found to be in excellent agreement with experimental data. Third, two strategies are developed to define numerical samples of polydisperse heterogeneous propellants: the Deterministic Strategy and the Random Selection Strategy. Using these strategies, numerical samples of industrial propellants are generated. The packing fraction is investigated and it is shown that the experimental values of the packing fraction can be achieved computationally. It is strongly believed that this Random Sphere Packing Model of propellants is a major step forward in the realistic computational modeling of heterogeneous propellant of combustion. In addition, a method of analysis of the morphology of heterogeneous propellants is developed which uses the concept of multi-point correlation functions. A set of intrinsic length scales of local density fluctuations in random heterogeneous propellants is identified by performing a Monte-Carlo study of the correlation functions. This method of analysis shows great promise for understanding the origins of the combustion instability of heterogeneous propellants, and is believed to become a valuable tool for the development of safe and reliable rocket engines.

Covalent binding of nanoliposomes to the surface of magnetotactic bacteria for the synthesis of self-propelled therapeutic agents.

PubMed

Taherkhani, Samira; Mohammadi, Mahmood; Daoud, Jamal; Martel, Sylvain; Tabrizian, Maryam

2014-05-27

The targeted and effective delivery of therapeutic agents remains an unmet goal in the field of controlled release systems. Magnetococcus marinus MC-1 magnetotactic bacteria (MTB) are investigated as potential therapeutic carriers. By combining directional magnetotaxis-microaerophilic control of these self-propelled agents, a larger amount of therapeutics can be delivered surpassing the diffusion limits of large drug molecules toward hard-to-treat hypoxic regions in solid tumors. The potential benefits of these carriers emphasize the need to develop an adequate method to attach therapeutic cargos, such as drug-loaded nanoliposomes, without substantially affecting the cell's ability to act as delivery agents. In this study, we report on a strategy for the attachment of liposomes to MTB (MTB-LP) through carbodiimide chemistry. The attachment efficacy, motility, and magnetic response of the MTB-LP were investigated. Results confirm that a substantial number of nanoliposomes (∼70) are efficiently linked with MTB without compromising functionality and motility. Cytotoxicity assays using three different cell types (J774, NIH/3T3, and Colo205) reveal that liposomal attachments to MTB formulation improve the biocompatibility of MTB, whereas attachment does not interfere with liposomal uptake.

TNT equivalency study for space shuttle (EOS). Volume 1: Management summary report

NASA Technical Reports Server (NTRS)

Wolfe, R. R.

1971-01-01

The existing TNT equivalency criterion for LO2/LH2 propellant is reevaluated. It addresses the static, on-pad phase of the space shuttle launch operations and was performed to determine whether the use of a TNT equivalency criterion lower than that presently used (60%) could be substantiated. The large quantity of propellant on-board the space shuttle, 4 million pounds, was considered of prime importance to the study. A qualitative failure analysis of the space shuttle (EOS) on the launch pad was made because it was concluded that available test data on the explosive yield of LO2/LH2 propellant was insufficient to support a reduction in the present TNT equivalency value, considering the large quantity of propellant used in the space shuttle. The failure analysis had two objectives. The first was to determine whether a failure resulting in the total release of propellant could occur. The second was to determine whether, if such a failure did occur, ignition could be delayed long enough to allow the degree of propellant mixing required to produce an explosion of 60% TNT equivalency since the explosive yield of this propellant is directly related to the quantities of LH2 and LO2 mixed at the time of the explosion.

Design criteria monograph for pressurized metal cases

NASA Technical Reports Server (NTRS)

1972-01-01

Organiation and presentation of data pertaining to design of solid propellant rocket engine cases are discussed. Design criteria are presented in form of monograph based on accumulated experience and knowledge. Improvements in reliability, cost effectiveness, and engine efficiency are stressed.

SRB Safety Analysis

NASA Image and Video Library

2003-09-11

Jeff Thon, an SRB mechanic with United Space Alliance, is lowered into a mockup of a segment of a solid rocket booster. He is testing a technique for vertical SRB propellant grain inspection. The inspection of segments is required as part of safety analysis.

Inert Reassessment Document for Diethyl Phthalate - CAS No. 84-66-2

EPA Pesticide Factsheets

Diethyl phthalate used is as a plasticizer in a wide variety of consumer products, including plastic packaging film, automotive components, toys, cosmetic formulations, toiletries, medical tubing, solid rocket propellants, and as a ingredient in aspirin.

Large scale production of densified hydrogen to the triple point and below

NASA Astrophysics Data System (ADS)

Swanger, A. M.; Notardonato, W. U.; E Fesmire, J.; Jumper, K. M.; Johnson, W. L.; Tomsik, T. M.

2017-12-01

Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage technology at NASA Kennedy Space Center led to the production of large quantities of densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. Overall densification performance of the system is explored, and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing, and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. The phenomenon, observed at two fill levels, is detailed herein. The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Large Scale Production of Densified Hydrogen to the Triple Point and Below

NASA Technical Reports Server (NTRS)

Swanger, A. M.; Notardonato, W. U.; Fesmire, J. E.; Jumper, K. M.; Johnson, W. L.; Tomsik, T. M.

2017-01-01

Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage technology at NASA Kennedy Space Center led to the production of large quantities of densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. Overall densification performance of the system is explored, and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing, and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. The phenomenon, observed at two fill levels, is detailed herein. The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Controlled propulsion of artificial magnetic nanostructured propellers.

PubMed

Ghosh, Ambarish; Fischer, Peer

2009-06-01

For biomedical applications, such as targeted drug delivery and microsurgery, it is essential to develop a system of swimmers that can be propelled wirelessly in fluidic environments with good control. Here, we report the construction and operation of chiral colloidal propellers that can be navigated in water with micrometer-level precision using homogeneous magnetic fields. The propellers are made via nanostructured surfaces and can be produced in large numbers. The nanopropellers can carry chemicals, push loads, and act as local probes in rheological measurements.

Large-eddy simulation of propeller wake at design operating conditions

NASA Astrophysics Data System (ADS)

Kumar, Praveen; Mahesh, Krishnan

2016-11-01

Understanding the propeller wake is crucial for efficient design and optimized performance. The dynamics of the propeller wake are also central to physical phenomena such as cavitation and acoustics. Large-eddy simulation is used to study the evolution of the wake of a five-bladed marine propeller from near to far field at design operating condition. The computed mean loads and phase-averaged flow field show good agreement with experiments. The propeller wake consisting of tip and hub vortices undergoes streamtube contraction, which is followed by the onset of instabilities as evident from the oscillations of the tip vortices. Simulation results reveal a mutual induction mechanism of instability where instead of the tip vortices interacting among themselves, they interact with the smaller vortices generated by the roll-up of the blade trailing edge wake in the near wake. Phase-averaged and ensemble-averaged flow fields are analyzed to explain the flow physics. This work is supported by ONR.

Fabrication of copper-based anodes via atmosphoric plasma spraying techniques

DOEpatents

Lu, Chun [Monroeville, PA

2012-04-24

A fuel electrode anode (18) for a solid oxide fuel cell is made by presenting a solid oxide fuel cell having an electrolyte surface (15), mixing copper powder with solid oxide electrolyte in a mixing step (24, 44) to provide a spray feedstock (30,50) which is fed into a plasma jet (32, 52) of a plasma torch to melt the spray feed stock and propel it onto an electrolyte surface (34, 54) where the spray feed stock flattens into lamellae layer upon solidification, where the layer (38, 59) is an anode coating with greater than 35 vol. % based on solids volume.

NASA reload program

NASA Technical Reports Server (NTRS)

Byington, Marshall

1993-01-01

Atlantic Research Corporation (ARC) contracted with NASA to manufacture and deliver thirteen small scale Solid Rocket Motors (SRM). These motors, containing five distinct propellant formulations, will be used for plume induced radiation studies. The information contained herein summarizes and documents the program accomplishments and results. Several modifications were made to the scope of work during the course of the program. The effort was on hold from late 1991 through August, 1992 while propellant formulation changes were developed. Modifications to the baseline program were completed in late-August and Modification No. 6 was received by ARC on September 14, 1992. The modifications include changes to the propellant formulation and the nozzle design. The required motor deliveries were completed in late-December, 1992. However, ARC agreed to perform an additional mix and cast effort at no cost to NASA and another motor was delivered in March, 1993.

Quantitative computer simulations of extraterrestrial processing operations

NASA Technical Reports Server (NTRS)

Vincent, T. L.; Nikravesh, P. E.

1989-01-01

The automation of a small, solid propellant mixer was studied. Temperature control is under investigation. A numerical simulation of the system is under development and will be tested using different control options. Control system hardware is currently being put into place. The construction of mathematical models and simulation techniques for understanding various engineering processes is also studied. Computer graphics packages were utilized for better visualization of the simulation results. The mechanical mixing of propellants is examined. Simulation of the mixing process is being done to study how one can control for chaotic behavior to meet specified mixing requirements. An experimental mixing chamber is also being built. It will allow visual tracking of particles under mixing. The experimental unit will be used to test ideas from chaos theory, as well as to verify simulation results. This project has applications to extraterrestrial propellant quality and reliability.

EPDM rocket motor insulation

NASA Technical Reports Server (NTRS)

Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

2003-01-01

A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components and with appropriate adjustment of curing and other additives functionally-required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g. powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf life characteristics.

EPDM rocket motor insulation

NASA Technical Reports Server (NTRS)

Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

2008-01-01

A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

EPDM rocket motor insulation

NASA Technical Reports Server (NTRS)

Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

2004-01-01

A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

Space Shuttle Five-Segment Booster (Short Course)

NASA Technical Reports Server (NTRS)

Graves, Stanley R.; Rudolphi, Michael (Technical Monitor)

2002-01-01

NASA is considering upgrading the Space Shuttle by adding a fifth segment (FSB) to the current four-segment solid rocket booster. Course materials cover design and engineering issues related to the Reusable Solid Rocket Motor (RSRM) raised by the addition of a fifth segment to the rocket booster. Topics cover include: four segment vs. five segment booster, abort modes, FSB grain design, erosive burning, enhanced propellant burn rate, FSB erosive burning model development and hardware configuration.

Surface Piercing Propeller Performance

DTIC Science & Technology

2005-09-01

solid body ( hydrodynamic cavitation ) or by high-intensity sound waves (acoustic cavitation). A Research study done by Yin Lu Young at UT studied and...discusses the effect of hydrodynamic cavitation , which occurs when pressure drops below the saturated vapor pressure, consequently resulting in the

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.

Fluid-Structure Interaction Effects on Mass Flow Rates in Solid Rocket Motors

DTIC Science & Technology

2015-09-02

FEA ) is explored. A propellant flap in a cross flow is analyzed. Comparisons are made between an analytical solution, a solely CFD solution, a manual...finite element analysis ( FEA ) is explored.  A  propellant flap in a cross flow is analyzed.  Comparisons are made between an analytical  solution, a...Condition Zones ..................................................................... 11  Figure 6: Pressure Boundary Condition Applied to  FEA  model

Credit PSR. This view shows the southeast and northeast facades ...

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

Credit PSR. This view shows the southeast and northeast facades of building as seen when looking west (264°). The open double doors reveal the curing room, which was kept at ambient temperatures. A maximum of 10,000 pounds (4,545 Kg) of class 1.1 propellants were permitted in this room, along with a maximum of 4 people. A separate room at the west end of the building housed temperature control equipment. Note the lightning rods on roof corners - Jet Propulsion Laboratory Edwards Facility, Solid Propellant Conditioning Building, Edwards Air Force Base, Boron, Kern County, CA

The Guggenheim Aeronautics Laboratory at Caltech and the creation of the modern rocket motor (1936-1946): How the dynamics of rocket theory became reality

NASA Astrophysics Data System (ADS)

Zibit, Benjamin Seth

This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the creation of the jet Propulsion Laboratory, the founding of the Aerojet Corporation, and emphasizes the issue of JPL's close relation to military development of the rocket becomes a core subject of this thesis. Cooperation between engineers in an academic setting and the military was not merely inevitable in the 1940s---it was actively fostered and proved quite profitable to all concerned. The deep relationship between the Guggenheim Aeronautics Laboratory and the Army Air Force was one model of the evolution of a permanent institutional edifice, weaving academic research and military end-use together. The dissertation concludes that what began as a modest effort to understand rocket theory in greater depth led within ten years to both research and development tracks which have profoundly altered the technological and military definition of modern history.

A search for experiments to exploit the space shuttle environment, volume 1

NASA Technical Reports Server (NTRS)

Fenn, J. B.

1979-01-01

A search for worthwhile experiments in pure and applied physics and chemistry which might take advantage of conditions achievable aboard the space shuttle is documented. Of particular interest were the very large pumping speeds at high or ultra high vacuum, the highly nonequilibrium composition of the ambient atmosphere, and the relative absence of gravitational effects. Ideas and suggestions were solicated in the course of visits to 31 research establishments in Western Europe, India, and Japan; conversations with over 90 scientists; and presentations at 3 international meetings. Intriguing possibilities emerged in the following arenas: (1) spectroscopy of the transition state in chemical reactions; (2) flame structure and analysis; (3) solid propellant combustion; (4) analysis of atmospheric composition; (5) turbulence effects on aerosol coagulation.

A theoretical analysis of vacuum arc thruster performance

NASA Technical Reports Server (NTRS)

Polk, James E.; Sekerak, Mike; Ziemer, John K.; Schein, Jochen; Qi, Niansheng; Binder, Robert; Anders, Andre

2001-01-01

In vacuum arc discharges the current is conducted through vapor evaporated from the cathode surface. In these devices very dense, highly ionized plasmas can be created from any metallic or conducting solid used as the cathode. This paper describes theoretical models of performance for several thruster configurations which use vacuum arc plasma sources. This analysis suggests that thrusters using vacuum arc sources can be operated efficiently with a range of propellant options that gives great flexibility in specific impulse. In addition, the efficiency of plasma production in these devices appears to be largely independent of scale because the metal vapor is ionized within a few microns of the cathode electron emission sites, so this approach is well-suited for micropropulsion.

Air-fluidized grains as a model system: Self-propelling and jamming

NASA Astrophysics Data System (ADS)

Daniels, Lynn J.

This thesis examines two concepts -- self-propelling and jamming -- that have been employed to unify disparate non-equilibrium systems, in the context of a monolayer of grains fluidized by a temporally and spatially homogeneous upflow of air. The first experiment examines the single particle dynamics of air-fluidized rods. For Brownian rods, equipartition of energy holds and rotational motion sets a timescale after which directional memory is lost. Air-fluidized rods no longer obey equipartion; they self-propel, moving preferentially along their long axis. We show that self-propelling can be treated phenomenologically as an enhanced memory effect causing directional memory to persist for times longer than expected for thermal systems. The second experiment studies dense collections of self-propelling air-fluidized rods. We observe collective propagating modes that give rise to anomalously large fluctuations in the local number density. We quantify these compression waves by calculating the dynamic structure factor and show that the wavespeed is weakly linear with increasing density. It has been suggested that the observed behavior might be explained using the framework put forth by Baskaran et al. [12]. The third experiment seeks to determine whether a force analogous to the critical Casimir force in fluids exists for a large sphere fluidized in the presence of a background of smaller spheres. The behavior of such a large sphere is fully characterized showing that, rather than behaving like a sphere driven by turbulence, the large ball self-propels. We also show that the background is responsible for the purely attractive, intermediate-ranged interaction force between two simultaneously-fluidized large balls. The final experiment seeks to determine what parameters control the diverging relaxation timescale associated with the jamming transition. By tilting our apparatus, we quantify pressure, packing fraction, and temperature simultaneously with dynamics as we approach jamming. We obtain an equation of state that agrees well with simulation and free volume theory. We collapse the relaxation time by defining a time- and energy-scale using pressure, consistent with recent simulation [82]. These experiments are further confirmation of the universality of the concepts of self-propelling and jamming.

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.

Liquid Oxygen Propellant Densification Production and Performance Test Results With a Large-Scale Flight-Weight Propellant Tank for the X33 RLV

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.; Meyer, Michael L.

2010-01-01

This paper describes in-detail a test program that was initiated at the Glenn Research Center (GRC) involving the cryogenic densification of liquid oxygen (LO2). A large scale LO2 propellant densification system rated for 200 gpm and sized for the X-33 LO2 propellant tank, was designed, fabricated and tested at the GRC. Multiple objectives of the test program included validation of LO2 production unit hardware and characterization of densifier performance at design and transient conditions. First, performance data is presented for an initial series of LO2 densifier screening and check-out tests using densified liquid nitrogen. The second series of tests show performance data collected during LO2 densifier test operations with liquid oxygen as the densified product fluid. An overview of LO2 X-33 tanking operations and load tests with the 20,000 gallon Structural Test Article (STA) are described. Tank loading testing and the thermal stratification that occurs inside of a flight-weight launch vehicle propellant tank were investigated. These operations involved a closed-loop recirculation process of LO2 flow through the densifier and then back into the STA. Finally, in excess of 200,000 gallons of densified LO2 at 120 oR was produced with the propellant densification unit during the demonstration program, an achievement that s never been done before in the realm of large-scale cryogenic tests.

Reliability of extended dorsal intercostal artery perforator propeller flaps for reconstruction of large myelomeningocele defects.

PubMed

Tenekeci, Goktekin; Basterzi, Yavuz

2017-01-01

Reconstruction of large myelomeningocele defects using extended (elongated beyond the lateral margin of the latissimus dorsi muscle) dorsal intercostal artery perforator (DICAP) propeller flaps is not recommended by previous studies. However, to provide tension-free and successful closure of a defect, the DICAP propeller flaps must sometimes be elongated beyond this margin. Our experience and results in this issue are discussed. In this article, reconstruction of 11 consecutive cases, with large myelomeningocele defects in which standard DICAP propeller flaps were incapable to close the defect, was achieved using extended DICAP propeller flaps between June 2013 and November 2015. At least two reliable perforators of the neighboring intervertebral spaces are included to supply the flap. Intramuscular dissection of perforators is performed to free the perforators from the surrounding muscle and to gain pedicle length as much as possible to prevent twisting and vascular compromise. All the flaps survived completely except for one patient who had superficial skin necrosis on the most distal part of the flap and had severe accompanying systemic disorders and died on postoperative 14th day. In 7 of 11 patients, venous congestion was noted, which resolved spontaneously. No hematoma or seroma formation was observed during the postoperative follow-up period. Dissection of multiple DICAPs supplying flaps enable us to harvest larger DICAP flaps possibly by providing better arterial supply and venous drainage. We use microsurgical instruments and 4.3× loupe magnification for pedicle dissection in this newborn population. This study shows the reliability of extended DICAP propeller flaps when multiple perforators at sixth or more cranial adjacent intercostal spaces are included in DICAP propeller flaps. Copyright © 2016 British Association of Plastic, Reconstructive and Aesthetic Surgeons. Published by Elsevier Ltd. All rights reserved.

ASRM propellant and igniter propellant development and process scale-up

NASA Technical Reports Server (NTRS)

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

1993-01-01

A program of formulation and process development for ANB-3652 motor propellant was conducted to validate design concepts and screen critical propellant composition and process parameters. Design experiments resulted in the selection of a less active grade of ferric oxide to provide better burning rate control, the establishment of AP fluidization conditions that minimized the adverse effects of particle attrition, and the selection of a higher mix temperature to improve mechanical properties. It is shown that the propellant can be formulated with AP and aluminum powder from various producers. An extended duration pilot plant run demonstrated stable equipment operation and excellent reproducibility of propellant properties. A similar program of formulation and process optimization culminating in large batch scaleup was conducted for ANB-3672 igniter propellant. The results for both ANB-3652 and ANB 37672 confirmed that their processing characteristics are compatible with full-scale production.

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.

Evaluation of Flygt Propeller Xixers for Double Shell Tank (DST) High Level Waste Auxiliary Solids Mobilization

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

PACQUET, E.A.

The River Protection Project (RPP) is planning to retrieve radioactive waste from the single-shell tanks (SST) and double-shell tanks (DST) underground at the Hanford Site. This waste will then be transferred to a waste treatment plant to be immobilized (vitrified) in a stable glass form. Over the years, the waste solids in many of the tanks have settled to form a layer of sludge at the bottom. The thickness of the sludge layer varies from tank to tank, from no sludge or a few inches of sludge to about 15 ft of sludge. The purpose of this technology and engineeringmore » case study is to evaluate the Flygt{trademark} submersible propeller mixer as a potential technology for auxiliary mobilization of DST HLW solids. Considering the usage and development to date by other sites in the development of this technology, this study also has the objective of expanding the knowledge base of the Flygt{trademark} mixer concept with the broader perspective of Hanford Site tank waste retrieval. More specifically, the objectives of this study delineated from the work plan are described.« less

Green Propellant Demonstration with Hydrazine Catalyst of F-16 Emergency Power Unit

NASA Technical Reports Server (NTRS)

Robinson, Joel W.; Brechbill, Shawn

2015-01-01

Some space vehicle and aircraft Auxiliary Power Units (APUs) use hydrazine propellant for generating power. Hydrazine is a toxic, hazardous fuel which requires special safety equipment and processes for handling and loading. In recent years, there has been development of two green propellants that could enable their use in APU's: the Swedish LMP-103S and the Air Force Research Laboratory (AFRL) AF-M315E. While there has been work on development of these propellants for thruster applications (Prisma and Green Propulsion Infusion Mission, respectively), there has been less focus on the application to power units. Beginning in 2012, an effort was started by the Marshall Space Flight Center (MSFC) on the APU application. The MSFC plan was to demonstrate green propellants with residual Space Shuttle hardware. The principal investigator was able to acquire a Solid Rocket Booster gas generator and an Orbiter APU. Since these test assets were limited in number, an Air Force equivalent asset was identified: the F-16 Emergency Power Unit (EPU). In June 2013, two EPU's were acquired from retired aircraft located at Davis Monthan Air Force Base. A gas generator from one of these EPU's was taken out of an assembly and configured for testing with a version of the USAF propellant with a higher water content (AF-M315EM) to reduce decomposition temperatures. Testing in November 2014 has shown that this green propellant is reactive with the Hydrazine catalyst (Shell 405) generating 300 psi of pressure with the existing F-16 EPU configuration. This paper will highlight the results of MSFC testing in collaboration with AFRL.

State-of-the-art of high-speed propeller noise prediction - A multidisciplinary approach and comparison with measured data

NASA Technical Reports Server (NTRS)

Dunn, Mark H.; Farassat, F.

1990-01-01

The results of NASA's Propeller Test Assessment program involving extensive flight tests of a large-scale advanced propeller are presented. This has provided the opportunity to evaluate the current capability of advanced propeller noise prediction utilizing principally the exterior acoustic measurements for the prediction of exterior noise. The principal object of this study was to evaluate the state-of-the-art of noise prediction for advanced propellers utilizing the best available codes of the disciplines involved. The effects of blade deformation on the aerodynamics and noise of advanced propellers were also studied. It is concluded that blade deformation can appreciably influence propeller noise and aerodynamics, and that, in general, centrifugal and blade forces must both be included in the calculation of blade forces. It is noted that the present capability for free-field noise prediction of the first three harmonics for advanced propellers is fairly good. Detailed data and diagrams of the test results are presented.

Analysis of quasi-hybrid solid rocket booster concepts for advanced earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Zurawski, Robert L.; Rapp, Douglas C.

1987-01-01

A study was conducted to assess the feasibility of quasi-hybrid solid rocket boosters for advanced Earth-to-orbit vehicles. Thermochemical calculations were conducted to determine the effect of liquid hydrogen addition, solids composition change plus liquid hydrogen addition, and the addition of an aluminum/liquid hydrogen slurry on the theoretical performance of a PBAN solid propellant rocket. The space shuttle solid rocket booster was used as a reference point. All three quasi-hybrid systems theoretically offer higher specific impulse when compared with the space shuttle solid rocket boosters. However, based on operational and safety considerations, the quasi-hybrid rocket is not a practical choice for near-term Earth-to-orbit booster applications. Safety and technology issues pertinent to quasi-hybrid rocket systems are discussed.

Technology Area Roadmap for In-Space Propulsion Technologies

NASA Technical Reports Server (NTRS)

Johnson, Les; Meyer, Michael; Palaszewski, Bryan; Coote, David; Goebel, Dan; White, Harold

2012-01-01

The exponential increase of launch system size.and cost.with delta-V makes missions that require large total impulse cost prohibitive. Led by NASA fs Marshall Space Flight Center, a team from government, industry, and academia has developed a flight demonstration mission concept of an integrated electrodynamic (ED) tethered satellite system called PROPEL: \\Propulsion using Electrodynamics.. The PROPEL Mission is focused on demonstrating a versatile configuration of an ED tether to overcome the limitations of the rocket equation, enable new classes of missions currently unaffordable or infeasible, and significantly advance the Technology Readiness Level (TRL) to an operational level. We are also focused on establishing a far deeper understanding of critical processes and technologies to be able to scale and improve tether systems in the future. Here, we provide an overview of the proposed PROPEL mission. One of the critical processes for efficient ED tether operation is the ability to inject current to and collect current from the ionosphere. Because the PROPEL mission is planned to have both boost and deboost capability using a single tether, the tether current must be capable of flowing in both directions and at levels well over 1 A. Given the greater mobility of electrons over that of ions, this generally requires that both ends of the ED tether system can both collect and emit electrons. For example, hollow cathode plasma contactors (HCPCs) generally are viewed as state-of-the-art and high TRL devices; however, for ED tether applications important questions remain of how efficiently they can operate as both electron collectors and emitters. Other technologies will be highlighted that are being investigated as possible alternatives to the HCPC such as Solex that generates a plasma cloud from a solid material (Teflon) and electron emission (only) technologies such as cold-cathode electron field emission or photo-electron beam generation (PEBG) techniques

The Potential of Aluminium Metal Powder as a Fuel for Space Propulsion Systems

NASA Astrophysics Data System (ADS)

Ismail, A. M.; Osborne, B.; Welch, C. S.

Metal powder propulsion systems have been addressed intermittently since the Second World War, initially in the field of underwater propulsion where research in the application of propelling torpedoes continues until this day. During the post war era, researchers attempted to utilise metal powders as a fuel for ram jet applications in missiles. The 1960's and 1970's saw additional interest in the use of `pure powder' propellants, i.e. fluidised metal fuel and oxidiser, both in solid particulate form. Again the application was for employment in space-constrained missiles where the idea was to maximise the performance of high energy density powder propellants in order to enhance the missile's flight duration. Metal powder as possible fuel was investigated for in-situ resource utilisation propulsion systems post-1980's where the emphasis was on the use of gaseous oxygen or liquid oxygen combined with aluminium metal powder for use as a ``lunar soil propellant'' or carbon dioxide and magnesium metal powder as a ``Martian propellant''.Albeit aluminium metal powder propellants are lower in performance than cryogenic and Earth storable propellants, the former does have an advantage inasmuch that the propulsion system is generic, i.e. it can be powered with chemicals mined and processed on Earth, the Moon and Mars. Thus, due to the potential refuelling capability, the lower performing aluminium metal powder propellant would effectively possess a much higher change in velocity (V) for multiple missions than the cryogenic or Earth storable propellant which is only suitable for one planet or one mission scenario, respectively.One of the principal limitations of long duration human spaceflight beyond cis-lunar orbit is the lack of refuelling capabilities on distant planets resulting in the reliance on con- ventional non-cryogenic, propellants produced on Earth. If one could develop a reliable propulsion system operating on pro- pellants derived entirely of ingredients found on nearby plan- etary bodies, then not only could mission duration be extended, larger amounts of payload could be ferried to and from the destination and eventually the cost of transporting propellant ingredients from Earth could be reduced, if not eliminated.

Gas-core reactor power transient analysis

NASA Technical Reports Server (NTRS)

Kascak, A. F.

1972-01-01

The gas core reactor is a proposed device which features high temperatures. It has applications in high specific impulse space missions, and possibly in low thermal pollution MHD power plants. The nuclear fuel is a ball of uranium plasma radiating thermal photons as opposed to gamma rays. This thermal energy is picked up before it reaches the solid cavity liner by an inflowing seeded propellant stream and convected out through a rocket nozzle. A wall-burnout condition will exist if there is not enough flow of propellant to convect the energy back into the cavity. A reactor must therefore operate with a certain amount of excess propellant flow. Due to the thermal inertia of the flowing propellant, the reactor can undergo power transients in excess of the steady-state wall burnout power for short periods of time. The objective of this study was to determine how long the wall burnout power could be exceeded without burning out the cavity liner. The model used in the heat-transfer calculation was one-dimensional, and thermal radiation was assumed to be a diffusion process.

Hybrids - Best of both worlds. [liquid and solid propellants mated for safe reliable and low cost launch vehicles

NASA Technical Reports Server (NTRS)

Goldberg, Ben E.; Wiley, Dan R.

1991-01-01

An overview is presented of hybrid rocket propulsion systems whereby combining solids and liquids for launch vehicles could produce a safe, reliable, and low-cost product. The primary subsystems of a hybrid system consist of the oxidizer tank and feed system, an injector system, a solid fuel grain enclosed in a pressure vessel case, a mixing chamber, and a nozzle. The hybrid rocket has an inert grain, which reduces costs of development, transportation, manufacturing, and launch by avoiding many safety measures that must be taken when operating with solids. Other than their use in launch vehicles, hybrids are excellent for simulating the exhaust of solid rocket motors for material development.

76 FR 57103 - Office of Commercial Space Transportation (AST); Notice of Availability of the Supplemental...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-09-15

..., consisting of a two-stage Castor 120 solid-propellant rocket motor with the addition of up to six Castor IVA or Castor IVXL rocket motors strapped to the first stage. The 1995 EA analyzed the potential...

Thermal Insulation Chemical Composition and Method of Manufacture.

DTIC Science & Technology

conditions in high temperature solid propellant gas generators can be formed of an ethylene propylene, diene monomer ( EPDM )-neoprene rubber binders containing...silica powder filler and aramid fibers. The specific chemical constituents include EPDM elastomer, 2 Chlorobutadiene 1,3 elastomer, Silica hydrate

Hyperintense Vessels on T2-PROPELLER-FLAIR in Patients with Acute MCA Stroke: Prediction of Arterial Stenosis and Perfusion Abnormality.

PubMed

Ahn, S J; Suh, S H; Lee, K-Y; Kim, J H; Seo, K-D; Lee, S

2015-11-01

Fluid-attenuated inversion recovery hyperintense vessels in stroke represent leptomeningeal collateral flow. We presumed that FLAIR hyperintense vessels would be more closely associated with arterial stenosis and perfusion abnormality in ischemic stroke on T2-PROPELLER-FLAIR than on T2-FLAIR. We retrospectively reviewed 35 patients with middle cerebral territorial infarction who underwent MR imaging. FLAIR hyperintense vessel scores were graded according to the number of segments with FLAIR hyperintense vessels in the MCA ASPECTS areas. We compared the predictability of FLAIR hyperintense vessels between T2-PROPELLER-FLAIR and T2-FLAIR for large-artery stenosis. The interagreement between perfusion abnormality and FLAIR hyperintense vessels was assessed. In subgroup analysis (9 patients with MCA horizontal segment occlusion), the association of FLAIR hyperintense vessels with ischemic lesion volume and perfusion abnormality volume was evaluated. FLAIR hyperintense vessel scores were significantly higher on T2-PROPELLER-FLAIR than on T2-FLAIR (3.50 ± 2.79 versus 1.21 ± 1.47, P < .01), and the sensitivity for large-artery stenosis was significantly improved on T2-PROPELLER-FLAIR (93% versus 68%, P = .03). FLAIR hyperintense vessels on T2-PROPELLER-FLAIR were more closely associated with perfusion abnormalities than they were on T2-FLAIR (κ = 0.64 and κ = 0.27, respectively). In subgroup analysis, FLAIR hyperintense vessels were positively correlated with ischemic lesion volume on T2-FLAIR, while the mismatch of FLAIR hyperintense vessels between the 2 sequences was negatively correlated with ischemic lesion volume (P = .01). In MCA stroke, FLAIR hyperintense vessels were more prominent on T2-PROPELLER-FLAIR compared with T2-FLAIR. In addition, FLAIR hyperintense vessels on T2-PROPELLER-FLAIR have a significantly higher sensitivity for predicting large-artery stenosis than they do on T2-FLAIR. Moreover, the areas showing FLAIR hyperintense vessels on T2-PROPELLER-FLAIR were more closely associated with perfusion abnormality than those on T2-FLAIR. © 2015 by American Journal of Neuroradiology.

Effects of normal acceleration on transient burning rate augmentation of an aluminized solid propellant

NASA Technical Reports Server (NTRS)

Northam, G. B.

1972-01-01

Instantaneous burning rate data for a polybutadiene acrylic acid propellant, containing 16 weight percent aluminum, were calculated from the pressure histories of a test motor with 96.77 sq cm of burning area and a 5.08-cm-thick propellant web. Additional acceleration tests were conducted with reduced propellant web thicknesses of 3.81, 2.54, and 1.27 cm. The metallic residue collected from the various web thickness tests was characterized by weight and shape and correlated with the instantaneous burning rate measurements. Rapid depressurization extinction tests were conducted in order that surface pitting characteristics due to localized increased burning rate could be correlated with the residue analysis and the instantaneous burning rate data. The acceleration-induced burning rate augmentation was strongly dependent on propellant distance burned, or burning time, and thus was transient in nature. The results from the extinction tests and the residue analyses indicate that the transient rate augmentation was highly dependent on local enhancement of the combustion zone heat feedback to the surface by the growth of molten residue particles on or just above the burning surface. The size, shape, and number density of molten residue particles, rather than the total residue weight, determined the acceleration-induced burning rate augmentation.

A Practical, Affordable Cryogenic Propellant Depot Based on ULA's Flight Experience

NASA Technical Reports Server (NTRS)

Kutter, Bernard F.; Zegler, Frank; O'Neil, Gary; Pitchford, Brian

2008-01-01

Mankind is embarking on the next step in the journey of human exploration. We are returning to the moon and eventually moving to Mars and beyond. The current Exploration architecture seeks a balance between the need for a robust infrastructure on the lunar surface, and the performance limitations of Ares I and V. The ability to refuel or top-off propellant tanks from orbital propellant depots offers NASA the opportunity to cost effectively and reliably satisfy these opposing requirements. The ability to cache large orbital quantities of propellant is also an enabling capability for missions to Mars and beyond. This paper describes an option for a propellant depot that enables orbital refueling supporting Exploration, national security, science and other space endeavors. This proposed concept is launched using a single EELV medium class rocket and thus does not require any orbital assembly. The propellant depot provides cryogenic propellant storage that utilizes flight proven technologies augmented with technologies currently under development. The propellant depot system, propellant management, flight experience, and key technologies are also discussed. Options for refueling the propellant depot along with an overview of Exploration architecture impacts are also presented.

Computational study of a self-cleaning process on superhydrophobic surface

NASA Astrophysics Data System (ADS)

Farokhirad, Samaneh

All substances around us are bounded by interfaces. In general, interface between different phases of materials are categorized as fluid-fluid, solid-fluid, and solid-solid. Fluid-fluid interfaces exhibit a distinct behavior by adapting their shape in response to external stimulus. For example, a liquid droplet on a substrate can undergo different wetting morphologies depending on topography and chemical composition of the surface. Fundamentally, interfacial phenomena arise at the limit between two immiscible phases, namely interface. The interface dynamic governs, to a great extent, physical processes such as impact and spreading of two immiscible media, and stabilization of foams and emulsions from break-up and coalescence. One of the recent challenging problems in the interface-driven fluid dynamics is the self-propulsion mechanism of droplets by means of different types of external forces such as electrical potential, or thermal Marangoni effect. Rapid removal of self-propelled droplet from the surface is an essential factor in terms of expense and efficiency for many applications including self-cleaning and enhanced heat and mass transfer to save energy and natural resources. A recent study on superhydrophobic nature of micro- and nanostructures of cicada wings offers a unique way for the self-propulsion process with no external force, namely coalescence-induced self-propelled jumping of droplet which can act effectively at any orientation. The biological importance of this new mechanism is associated with protecting such surfaces from long term exposure to colloidal particles such as microbial colloids and virus particles. Different interfacial phenomena can occur after out-of-plane jumping of droplet. If the departed droplet is landed back by gravity, it may impact and spread on the surface or coalesce with another droplet and again self-peopled itself to jump away from the surface. The complete removal of the propelled droplet to a sufficient distance beyond the boundary layer of the surface can be accomplished with a surface-parallel shear flow. This thesis presents an investigation of the physics involved in the mechanism of coalescence-induced self-propelled jumping of droplet with and without particle presence, through the use of numerical simulation. (Abstract shortened by ProQuest.).

Study of solid rocket motors for a space shuttle booster. Appendix E: Environmental impact statement, solid rocket motor, space shuttle booster

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the combustion products resulting from the solid propellant rocket engines of the space shuttle booster is presented. Calculation of the degree of pollution indicates that the only potentially harmful pollutants, carbon monoxide and hydrochloric acid, will be too diluted to constitute a hazard. The mass of products ejected during a launch within the troposphere is insignificant in terms of similar materials that enter the atmosphere from other sources. Noise pollution will not exceed that obtained from the Saturn 5 launch vehicle.

Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Martin, Heath T.

2017-01-01

An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

Microexplosions and ignition dynamics in engineered aluminum/polymer fuel particles

DOE PAGES

Rubio, Mario A.; Gunduz, I. Emre; Groven, Lori J.; ...

2016-11-11

Aluminum particles are widely used as a metal fuel in solid propellants. However, poor combustion efficiencies and two-phase flow losses result due in part to particle agglomeration. Engineered composite particles of aluminum (Al) with inclusions of polytetrafluoroethylene (PTFE) or low-density polyethylene (LDPE) have been shown to improve ignition and yield smaller agglomerates in solid propellants, recently. Reductions in agglomeration were attributed to internal pressurization and fragmentation (microexplosions) of the composite particles at the propellant surface. We explore the mechanisms responsible for microexplosions in order to better understand the combustion characteristics of composite fuel particles. Single composite particles of Al/PTFE andmore » Al/LDPE with diameters between 100 and 1200 µm are ignited on a substrate to mimic a burning propellant surface in a controlled environment using a CO 2 laser in the irradiance range of 78–7700 W/cm 2. Furthermore, the effects of particle size, milling time, and inclusion content on the resulting ignition delay, product particle size distributions, and microexplosion tendencies are reported. For example particles with higher PTFE content (30 wt%) had laser flux ignition thresholds as low as 77 W/cm 2, exhibiting more burning particle dispersion due to microexplosions compared to the other materials considered. Composite Al/LDPE particles exhibit relatively high ignition thresholds compared to Al/PTFE particles, and microexplosions were observed only with laser fluxes above 5500 W/cm 2 due to low LDPE reactivity with Al resulting in negligible particle self-heating. However, results show that microexplosions can occur for Al containing both low and high reactivity inclusions (LDPE and PTFE, respectively) and that polymer inclusions can be used to tailor the ignition threshold. Furthermore, this class of modified metal particles shows significant promise for application in many different energetic materials that use metal fuels.« less

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

Rubio, Mario A.; Gunduz, I. Emre; Groven, Lori J.

Aluminum particles are widely used as a metal fuel in solid propellants. However, poor combustion efficiencies and two-phase flow losses result due in part to particle agglomeration. Engineered composite particles of aluminum (Al) with inclusions of polytetrafluoroethylene (PTFE) or low-density polyethylene (LDPE) have been shown to improve ignition and yield smaller agglomerates in solid propellants, recently. Reductions in agglomeration were attributed to internal pressurization and fragmentation (microexplosions) of the composite particles at the propellant surface. We explore the mechanisms responsible for microexplosions in order to better understand the combustion characteristics of composite fuel particles. Single composite particles of Al/PTFE andmore » Al/LDPE with diameters between 100 and 1200 µm are ignited on a substrate to mimic a burning propellant surface in a controlled environment using a CO 2 laser in the irradiance range of 78–7700 W/cm 2. Furthermore, the effects of particle size, milling time, and inclusion content on the resulting ignition delay, product particle size distributions, and microexplosion tendencies are reported. For example particles with higher PTFE content (30 wt%) had laser flux ignition thresholds as low as 77 W/cm 2, exhibiting more burning particle dispersion due to microexplosions compared to the other materials considered. Composite Al/LDPE particles exhibit relatively high ignition thresholds compared to Al/PTFE particles, and microexplosions were observed only with laser fluxes above 5500 W/cm 2 due to low LDPE reactivity with Al resulting in negligible particle self-heating. However, results show that microexplosions can occur for Al containing both low and high reactivity inclusions (LDPE and PTFE, respectively) and that polymer inclusions can be used to tailor the ignition threshold. Furthermore, this class of modified metal particles shows significant promise for application in many different energetic materials that use metal fuels.« less

Study of solid rocket motors for a space shuttle booster. Volume 2, book 1: Analysis and design

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the factors which determined the selection of the solid rocket propellant engines for the space shuttle booster is presented. The 156 inch diameter, parallel burn engine was selected because of its transportability, cost effectiveness, and reliability. Other factors which caused favorable consideration are: (1) recovery and reuse are feasible and offer substantial cost savings, (2) abort can be easily accomplished. and (3) ecological effects are acceptable.

Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1997-01-01

Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

Homogenization Issues in the Combustion of Heterogeneous Solid Propellants

NASA Technical Reports Server (NTRS)

Chen, M.; Buckmaster, J.; Jackson, T. L.; Massa, L.

2002-01-01

We examine random packs of discs or spheres, models for ammonium-perchlorate-in-binder propellants, and discuss their average properties. An analytical strategy is described for calculating the mean or effective heat conduction coefficient in terms of the heat conduction coefficients of the individual components, and the results are verified by comparison with those of direct numerical simulations (dns) for both 2-D (disc) and 3-D (sphere) packs across which a temperature difference is applied. Similarly, when the surface regression speed of each component is related to the surface temperature via a simple Arrhenius law, an analytical strategy is developed for calculating an effective Arrhenius law for the combination, and these results are verified using dns in which a uniform heat flux is applied to the pack surface, causing it to regress. These results are needed for homogenization strategies necessary for fully integrated 2-D or 3-D simulations of heterogeneous propellant combustion.

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.

Study of the thermal degradation mechanism of a composite propellant. [using electron microscopes

NASA Technical Reports Server (NTRS)

Schmidt, W. G.

1975-01-01

The current experimental program was designed to systematically investigate the role of the oxidizer in the thermal degradation process of composite propellants. The scanning electron microscope (SEM) was used to examine the failure sites in thermally degraded propellant samples. The formulation variables tested were oxidizer purity, oxidizer particle size, and oxidizer to binder bonding agent. The binder, a saturated hydrocarbon, was kept constant throughout the experiments. The oxidizers were: AP, chlorate-doped AP, arsenate-doped AP, and phosphate-doped AP. The oxidizer particle size distribution was 60% of the large fraction and 40% of the small fraction. The bonding agent, when present, was used at the 0.15% level. The data showed that both the oxidizer purity and particle size had an important affect on the thermal degradation process. The affect of the oxidizer particle size was more noticeable at the higher temperature and stress levels. An examination of the failure site, by SEM, of propellants subject to these latter conditions indicated that the fracturing of the large oxidizer particles led to the propellant cracking.

Dorsal Intercostal Artery Perforator Propeller Flaps: A Reliable Option in Reconstruction of Large Meningomyelocele Defects.

PubMed

Basterzi, Yavuz; Tenekeci, Goktekin

2016-04-01

Several options have been reported for the reconstruction of myelomeningocele defects. In this article, we present our experience on soft tissue reconstruction of myelomeningocele defects by using island propeller dorsal intercostal artery perforator (DIAP) flaps. Between January 2008 and February 2014, all newborns with large myelomeningocele defects (13 newborns) were reconstructed with island propeller DIAP flaps. All flaps survived completely. In 8 patients out of 13, venous insufficiency was observed which then resolved spontaneously. Flap donor sites were closed primarily. Myelomeningocele defects with a diameter larger than 5 cm require reconstruction with flaps. To mobilize a well-vascularized tissue over the defect without tension in which the suture lines will not overlap over the midline where the dura is repaired and over the meninges is one of the goals of reconstruction for such defects. Perforator propeller flaps enable us to reach those goals. Use of perforator flaps provides 2 important advantages, namely, more predictability and also more freedom in mobilizing flaps toward the defect. This study proves the reliability of DIAP propeller flaps in the reconstruction of myelomeningocele defects.

Drag and Propulsive Characteristics of Air-Cooled Engine-Nacelle Installations for Large Airplanes, Special Report

NASA Technical Reports Server (NTRS)

Silverstein, Abe; Wilson, Herbert A., Jr.

1939-01-01

An investigation is in progress in the NACA full-scale wind tunnel to determine the drag and propulsive efficiency of nacelle sizes. In contrast with the usual tests with a single nacelle, these tests were conducted with nacelle-propeller installations on a large model of a 4-engine airplane. Data are presented on the first part of the investigation, covering seven nacelle arrangements with nacelle diameters from 0.53 to 1.5 times the wing thickness. These ratios are similar to those occurring on airplane weighing from about 20 to 100 tons. The results show that the drag, the propulsive efficiency, and the overall efficiency of the various nacelle arrangements as functions of the nacelle size, the propeller position, and the airplane lift coefficient. The effect of the nacelles on the aerodynamic characteristics of the model are shown for both propeller-removed and propeller-operating conditions.

U.S. Army Workshop on Solid-Propellant Ignition and Combustion Modeling.

DTIC Science & Technology

1997-07-01

saving tool in the design, development, testing, and evaluation of future gun-propulsion systems , and that, under current funding constraints, research...53 7.1 What systems are currently being addressed...9 ............. . . .. .. . . . . . . . . . . . . . . . . . . . . . . . . 56 7.5 What model systems might be valuable for

Acute Oral Toxicity of DIGL-RP Solid Propellant in Sprague-Dawley Rats

DTIC Science & Technology

1989-11-30

protein droplet and cast formation, glomeruli and cortical tubules Liver--diffuse vacuolation Stomach--multifocal, acute, necrotizing gastritis ...The choice of tissues examined histologically was biased by gross evaluation. Indications of renal protein loss were noted in five animals (casts and

Solid propellant grain design and internal ballistics

NASA Technical Reports Server (NTRS)

1972-01-01

The ballistic aspects of grain design were studied to outline the steps necessary to achieve a successful grain design. The relationships of the grain design to steady-state mass balance and erosive burning are considered. Grain design criteria is reviewed, and recommended design criteria are included.

14 CFR 420.63 - Explosive siting.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Explosive siting. 420.63 Section 420.63 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... launch site boundary; (2) A listing of the maximum quantities of liquid and solid propellants and other...

14 CFR 420.63 - Explosive siting.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Explosive siting. 420.63 Section 420.63 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... launch site boundary; (2) A listing of the maximum quantities of liquid and solid propellants and other...

14 CFR 420.63 - Explosive siting.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Explosive siting. 420.63 Section 420.63 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... launch site boundary; (2) A listing of the maximum quantities of liquid and solid propellants and other...

Ignition of Propellants Through Nanostructured Materials

DTIC Science & Technology

2016-03-31

the solid oxidizer, potassium permanganate (KMnO4) and boron–potassium nitrate (BKNO3) were used. While technically BKNO3 is not merely an oxidizer...unpurified SWCNT) EDS Energy Dispersive Spectroscopy Fe iron FPS frames per second GO graphene oxide KMnO4 potassium permanganate MIE

Primary Eye Irritation Potential of DIGL-RP Solid Propellant in Rabbits

DTIC Science & Technology

1989-10-01

0 Markedly deepened rugae , congestion, swelling., moderate circumiddial ".,peremia o. injection, any of these or...of the Iris (other than a slight deepening of the rugae or a slight hyperemia of the clrcumcomeal blood vessels); an obvious swelling in the

Multi-level of Fidelity Multi-Disciplinary Design Optimization of Small, Solid-Propellant Launch Vehicles

NASA Astrophysics Data System (ADS)

Roshanian, Jafar; Jodei, Jahangir; Mirshams, Mehran; Ebrahimi, Reza; Mirzaee, Masood

A new automated multi-level of fidelity Multi-Disciplinary Design Optimization (MDO) methodology has been developed at the MDO Laboratory of K.N. Toosi University of Technology. This paper explains a new design approach by formulation of developed disciplinary modules. A conceptual design for a small, solid-propellant launch vehicle was considered at two levels of fidelity structure. Low and medium level of fidelity disciplinary codes were developed and linked. Appropriate design and analysis codes were defined according to their effect on the conceptual design process. Simultaneous optimization of the launch vehicle was performed at the discipline level and system level. Propulsion, aerodynamics, structure and trajectory disciplinary codes were used. To reach the minimum launch weight, the Low LoF code first searches the whole design space to achieve the mission requirements. Then the medium LoF code receives the output of the low LoF and gives a value near the optimum launch weight with more details and higher fidelity.

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

NASA Technical Reports Server (NTRS)

Kibbey, Timothy P.

2015-01-01

The ideal rocket equation is manipulated to demonstrate the essential link between propellant density and specific impulse as the two primary stage performance drivers for a launch vehicle. This is illustrated by examining volume-limited stages such as first stages and boosters. This proves to be a good approximation for first-order or Phase A vehicle design studies for solid rocket motors and for liquid stages, except when comparing to hydrogen-fueled stages. A next-order mass model is developed that is able to model the mass differences between hydrogen-fueled and other stages. Propellants considered range in density from liquid methane to inhibited red fuming nitric acid. Calculated comparisons are shown for solid rocket boosters, liquid first stages, liquid upper stages, and a balloon-deployed single-stage-to-orbit concept. The derived relationships are ripe for inclusion in a multi-stage design space exploration and optimization algorithm, as well as for single-parameter comparisons such as those shown herein.

Inverse Leidenfrost effect: self-propelling drops on a bath

NASA Astrophysics Data System (ADS)

Gauthier, Anais; van der Meer, Devaraj; Lohse, Detlef; Physics of Fluids Team

2017-11-01

When deposited on very hot solid, volatile drops can levitate over a cushion of vapor, in the so-called Leidenfrost state. This phenomenon can also be observed on a hot bath and similarly to the solid case, drops are very mobile due to the absence of contact with the substrate that sustains them. We discuss here a situation of ``inverse Leidenfrost effect'' where room-temperature drops levitate on a liquid nitrogen pool - the vapor is generated here by the bath sustaining the relatively hot drop. We show that the drop's movement is not random: the liquid goes across the bath in straight lines, a pattern only disrupted by elastic bouncing on the edges. In addition, the drops are initially self-propelled; first at rest, they accelerate for a few seconds and reach velocities of the order of a few cm/s, before slowing down. We investigate experimentally the parameters that affect their successive acceleration and deceleration, such as the size and nature of the drops and we discuss the origin of this pattern.

Flocking Transition in Confluent Tissues

NASA Astrophysics Data System (ADS)

Paoluzzi, Matteo; Giavazzi, Fabio; Macchi, Marta; Scita, Giorgio; Cerbino, Roberto; Manning, Lisa; Marchetti, Cristina

The emerging of collective migration in biological tissues plays a pivotal role in embryonic morphogenesis, wound healing and cancer invasion. While many aspects of single cell movements are well established, the mechanisms leading to coherent displacements of cohesive cell groups are still poorly understood. Some of us recently proposed a Self-Propelled Voronoi (SPV) model of dense tissues that combines self-propelled particle models and vertex models of confluent cell layers and exhibits a liquid-solid transition as a function of cell shape and cell motility. We now examine the role of cell polarization on collective cell dynamics by introducing an orientation mechanism that aligns cell polarization with local cell motility. The model predicts a density-independent flocking transition tuned by the strength of the aligning interaction, with both solid and liquid flocking states existing in different regions of parameter space. MP and MCM were supported by the Simons Foundation Targeted Grant in the Mathematical Modeling of Living Systems Number: 342354 and by the Syracuse Soft Matter Program.

Simple-1: Development stage of the data transmission system for a solid propellant mid-power rocket model

NASA Astrophysics Data System (ADS)

Yarce, Andrés; Sebastián Rodríguez, Juan; Galvez, Julián; Gómez, Alejandro; García, Manuel J.

2017-06-01

This paper presents the development stage of a communication module for a solid propellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

Applied algorithm in the liner inspection of solid rocket motors

NASA Astrophysics Data System (ADS)

Hoffmann, Luiz Felipe Simões; Bizarria, Francisco Carlos Parquet; Bizarria, José Walter Parquet

2018-03-01

In rocket motors, the bonding between the solid propellant and thermal insulation is accomplished by a thin adhesive layer, known as liner. The liner application method involves a complex sequence of tasks, which includes in its final stage, the surface integrity inspection. Nowadays in Brazil, an expert carries out a thorough visual inspection to detect defects on the liner surface that may compromise the propellant interface bonding. Therefore, this paper proposes an algorithm that uses the photometric stereo technique and the K-nearest neighbor (KNN) classifier to assist the expert in the surface inspection. Photometric stereo allows the surface information recovery of the test images, while the KNN method enables image pixels classification into two classes: non-defect and defect. Tests performed on a computer vision based prototype validate the algorithm. The positive results suggest that the algorithm is feasible and when implemented in a real scenario, will be able to help the expert in detecting defective areas on the liner surface.

Computation of turbulent reacting flow in a solid-propellant ducted rocket

NASA Astrophysics Data System (ADS)

Chao, Yei-Chin; Chou, Wen-Fuh; Liu, Sheng-Shyang

1995-05-01

A mathematical model for computation of turbulent reacting flows is developed under general curvilinear coordinate systems. An adaptive, streamline grid system is generated to deal with the complex flow structures in a multiple-inlet solid-propellant ducted rocket (SDR) combustor. General tensor representations of the k-epsilon and algebraic stress (ASM) turbulence models are derived in terms of contravariant velocity components, and modification caused by the effects of compressible turbulence is also included in the modeling. The clipped Gaussian probability density function is incorporated in the combustion model to account for fluctuations of properties. Validation of the above modeling is first examined by studying mixing and reacting characteristics in a confined coaxial-jet problem. This is followed by study of nonreacting and reacting SDR combustor flows. The results show that Gibson and Launder's ASM incorporated with Sarkar's modification for compressible turbulence effects based on the general curvilinear coordinate systems yields the most satisfactory prediction for this complicated SDR flowfield.

Computation of turbulent reacting flow in a solid-propellant ducted rocket

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

Chao, Y.; Chou, W.; Liu, S.

1995-05-01

A mathematical model for computation of turbulent reacting flows is developed under general curvilinear coordinate systems. An adaptive, streamline grid system is generated to deal with the complex flow structures in a multiple-inlet solid-propellant ducted rocket (SDR) combustor. General tensor representations of the k-epsilon and algebraic stress (ASM) turbulence models are derived in terms of contravariant velocity components, and modification caused by the effects of compressible turbulence is also included in the modeling. The clipped Gaussian probability density function is incorporated in the combustion model to account for fluctuations of properties. Validation of the above modeling is first examined bymore » studying mixing and reacting characteristics in a confined coaxial-jet problem. This is followed by study of nonreacting and reacting SDR combustor flows. The results show that Gibson and Launder`s ASM incorporated with Sarkar`s modification for compressible turbulence effects based on the general curvilinear coordinate systems yields the most satisfactory prediction for this complicated SDR flowfield. 36 refs.« less

Glow-to-arc transition in plasma-assisted combustion at 100 MPa

NASA Astrophysics Data System (ADS)

Larsson, A.; Andreasson, S.

2015-04-01

Electric energy can be added to the combustion of solid propellants in a gun in order to augment and to control parts of the internal ballistic cycle of the launch of a projectile. The pressure in the chamber and bore during launch is typically several hundred megapascal and the electric energy must be delivered to the flame at such a pressure level. To increase the understanding of the interaction between a flame and an electrical discharge at elevated pressure, experiments have been performed at 100 MPa in a combustion chamber where electric current has been conducted through the flame of a solid propellant. Pressure, voltage and current have been measured. The measured signals have been analysed and interpreted. The sequence of events has been interpreted as an initial formation of a glow-like discharge in the flame followed by a discharge mode transition to a filamentary arc discharge. The transition is shown to be dependent on the flame conductivity. For the test propellant used (Nzk5230 doped with 5% potassium nitrate), the flame conductivity is calculated to be 0.84 S m-1 and the discharge mode transition is found to occur after a dissipation of 0.2-0.4 kJ, or 11-22 kJ m-1 of electric energy, at an electric power of 0.1-0.5 MW.

Applications of High-speed motion analysis system on Solid Rocket Motor (SRM)

NASA Astrophysics Data System (ADS)

Liu, Yang; He, Guo-qiang; Li, Jiang; Liu, Pei-jin; Chen, Jian

2007-01-01

High-speed motion analysis system could record images up to 12,000fps and analyzed with the image processing system. The system stored data and images directly in electronic memory convenient for managing and analyzing. The high-speed motion analysis system and the X-ray radiography system were established the high-speed real-time X-ray radiography system, which could diagnose and measure the dynamic and high-speed process in opaque. The image processing software was developed for improve quality of the original image for acquiring more precise information. The typical applications of high-speed motion analysis system on solid rocket motor (SRM) were introduced in the paper. The research of anomalous combustion of solid propellant grain with defects, real-time measurement experiment of insulator eroding, explosion incision process of motor, structure and wave character of plume during the process of ignition and flameout, measurement of end burning of solid propellant, measurement of flame front and compatibility between airplane and missile during the missile launching were carried out using high-speed motion analysis system. The significative results were achieved through the research. Aim at application of high-speed motion analysis system on solid rocket motor, the key problem, such as motor vibrancy, electrical source instability, geometry aberrance, and yawp disturbance, which damaged the image quality, was solved. The image processing software was developed which improved the capability of measuring the characteristic of image. The experimental results showed that the system was a powerful facility to study instantaneous and high-speed process in solid rocket motor. With the development of the image processing technique, the capability of high-speed motion analysis system was enhanced.

Propeller swirl effect on single-engine general-aviation aircraft stall-spin tendencies

NASA Technical Reports Server (NTRS)

Katz, Joseph; Feistel, Terry W.

1987-01-01

An investigation is conducted of the effect of a single engine, untapered low wing general aviation aircraft propeller's swirl on the craft's stall pattern. The asymmetrical character of the propeller's swirl can trigger an early stall of one of the wings, aggravating the spin-entry condition. It is shown that the combination of this propeller-induced effect with adverse sideslip can result in large and abrupt changes in the rolling moment, in such conditions as uncoordinated low speed turning maneuvers where the pilot yaws the aircraft with wings level, rather than rolling it.

Space Shuttle propulsion performance reconstruction from flight data

NASA Technical Reports Server (NTRS)

Rogers, Robert M.

1989-01-01

The aplication of extended Kalman filtering to estimating Space Shuttle Solid Rocket Booster (SRB) performance, specific impulse, from flight data in a post-flight processing computer program. The flight data used includes inertial platform acceleration, SRB head pressure, and ground based radar tracking data. The key feature in this application is the model used for the SRBs, which represents a reference quasi-static internal ballistics model normalized to the propellant burn depth. Dynamic states of mass overboard and propellant burn depth are included in the filter model to account for real-time deviations from the reference model used. Aerodynamic, plume, wind and main engine uncertainties are included.

Environmentally Sound Processing Technology: JANNAF Safety and Environmental Protection Subcommittee and Propellant Development and Characterization Subcommittee Joint Workshop

NASA Technical Reports Server (NTRS)

Pickett, Lorri A. (Editor)

1995-01-01

Topics covered include: Risk assessment of hazardous materials, Automated systems for pollution prevention and hazardous materials elimination, Study design for the toxicity evaluation of ammonium perchlorate, Plasma sprayed bondable stainless surface coatings, Development of CFC-free cleaning processes, New fluorinated solvent alternatives to ozone depleting solvents, Cleaning with highly fluorinated liquids, Biotreatment of propyleneglycol nitrate by anoxic denitrification, Treatment of hazardous waste with white rot fungus, Hydrothermal oxidation as an environmentally benign treatment technology, Treatment of solid propellant manufacturing wastes by base hydrolysis, Design considerations for cleaning using supercritical fluid technology, and Centrifugal shear carbon dioxide cleaning.

Propulsion System Development for the Iodine Satellite (iSAT) Demonstration Mission

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Stephen R.; Seixal, Joao F.; Mauro, Stephanie L.; Lewis, Brandon L.; Jerman, Gregory A.; Calvert, Derek H.; Dankanich, John; Kamhawi, Hani; Hickman, Tyler A.;

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

DTIC Science & Technology

2010-05-01

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

Draft environmental impact statement: Space Shuttle Advanced Solid Rocket Motor Program

NASA Technical Reports Server (NTRS)

1988-01-01

The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site.

Chemical Laser Solid Fuels Program

DTIC Science & Technology

1976-12-01

liquids. Solid propellant gas generators which can supply all of the ^(Tl-*fs DD , FORM w73 JAN 71 I"* EDITION OF 1 NOV SS IS OBSOLETE...seven tests, the mean weight yield was 13.24 ± 0.09 percent which is 97.72 percent of the theoretical weight yield of 13. SS percent for this...early in the test and peaks as the deuteriun flow rate is dropping at the burnout of the grain. The pressure differential across the filter discs

Nitramine Composite Solid Propellant Modelling

DTIC Science & Technology

1989-07-01

nonshaded areas, Sox values of zero are calculated for the various configurations indicated. The upper shaded area corresponds to one- half cf the...results: QDX ST Q X P F -XP QX "X TsX - To Gs X+ + Cf l sT),X +(l-Xf) ,X (134) In the above equations, and in the ones that follow, the solid phase...Surface Warfare Center, White Oak Laboratory, Silver Spring Code R10, S. Jacobs (1) Code R16, Dr. G. B. Wilmot (1) 1 Naval Underwater Systems Center

Solid Propellant Subscale Burning Rate Analysis Methods for US and Selected NATO Facilities

DTIC Science & Technology

2002-01-01

impossibility of the center of a particle lying closer than its radius from a solid boundary, * Due to surface tension and sedimentation (tends to level...34 effect (for bottom cast or bayonet cast grains) may consist of sedimentation of larger particles against the walls during casting flow, with the...February 2000. 91 Ratti A., "Metodi di Riduzione Dati Balistici per i Boosters a Propellente Solido di Ariane-4 e di Ariane-5," M.Sc. Thesis in Aerospace

Combustion of Solid Propellants (La Combustion des Propergols Solides)

DTIC Science & Technology

1991-07-01

cin~tiques initiales. Il relatives A Ia granulom ~trie ct la surface eat s~me possible dWaller plus loin et sp~cifique des catalyseurs existent, il est...grand nombro do vari~t~s granulom ~triques des proporgols. On pout ainsi observer uno mont donc utilis~es induatriellemont pour notte influence du temps...et do la ajuster la vitosso des vari~tds do tempdrature do laminage sur la diminution granulom ~trie moyenno 400, 200, 100, 10, 3 de l’exposant do

DEVELOPMENT OF FLEXIBLE INSULATION FOR SOLID PROPELLANT ROCKET MOTOR CASES

DTIC Science & Technology

acrylonitrile-phenol furfural -asbestos composition. Other promising materials which are reported are based on two types of liquid butadiene/styrene cbers. The...This material was based on a butadiene/acrylonitrile-phenol furfural -asbestos composition. Other promising materials which are reported are based on two

Large-eddy simulation of propeller noise

NASA Astrophysics Data System (ADS)

Keller, Jacob; Mahesh, Krishnan

2016-11-01

We will discuss our ongoing work towards developing the capability to predict far field sound from the large-eddy simulation of propellers. A porous surface Ffowcs-Williams and Hawkings (FW-H) acoustic analogy, with a dynamic endcapping method (Nitzkorski and Mahesh, 2014) is developed for unstructured grids in a rotating frame of reference. The FW-H surface is generated automatically using Delaunay triangulation and is representative of the underlying volume mesh. The approach is validated for tonal trailing edge sound from a NACA 0012 airfoil. LES of flow around a propeller at design advance ratio is compared to experiment and good agreement is obtained. Results for the emitted far field sound will be discussed. This work is supported by ONR.

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Aspect-Ratio-10 Wing, Two Propellers, and Blowing Flaps

NASA Technical Reports Server (NTRS)

Griffin, Roy N., Jr.; Holzhauser, Curt A.; Weiberg, James A.

1958-01-01

An investigation was made to determine the lifting effectiveness and flow requirements of blowing over the trailing-edge flaps and ailerons on a large-scale model of a twin-engine, propeller-driven airplane having a high-aspect-ratio, thick, straight wing. With sufficient blowing jet momentum to prevent flow separation on the flap, the lift increment increased for flap deflections up to 80 deg (the maximum tested). This lift increment also increased with increasing propeller thrust coefficient. The blowing jet momentum coefficient required for attached flow on the flaps was not significantly affected by thrust coefficient, angle of attack, or blowing nozzle height.

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Kamhawi, Hani

2015-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at -75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Numerical Modeling of Conjugate Heat Transfer in Fluid Network

NASA Technical Reports Server (NTRS)

Majumdar, Alok

2004-01-01

Fluid network modeling with conjugate heat transfer has many applications in Aerospace engineering. In modeling unsteady flow with heat transfer, it is important to know the variation of wall temperature in time and space to calculate heat transfer between solid to fluid. Since wall temperature is a function of flow, a coupled analysis of temperature of solid and fluid is necessary. In cryogenic applications, modeling of conjugate heat transfer is of great importance to correctly predict boil-off rate in propellant tanks and chill down of transfer lines. In TFAWS 2003, the present author delivered a paper to describe a general-purpose computer program, GFSSP (Generalized Fluid System Simulation Program). GFSSP calculates flow distribution in complex flow circuit for compressible/incompressible, with or without heat transfer or phase change in all real fluids or mixtures. The flow circuit constitutes of fluid nodes and branches. The mass, energy and specie conservation equations are solved at the nodes where as momentum conservation equations are solved at the branches. The proposed paper describes the extension of GFSSP to model conjugate heat transfer. The network also includes solid nodes and conductors in addition to fluid nodes and branches. The energy conservation equations for solid nodes solves to determine the temperatures of the solid nodes simultaneously with all conservation equations governing fluid flow. The numerical scheme accounts for conduction, convection and radiation heat transfer. The paper will also describe the applications of the code to predict chill down of cryogenic transfer line and boil-off rate of cryogenic propellant storage tank.

A Summary of the Slush Hydrogen Technology Program for the National Aero-Space Plane

NASA Technical Reports Server (NTRS)

Mcnelis, Nancy B.; Hardy, Terry L.; Whalen, Margaret V.; Kudlac, Maureen T.; Moran, Matthew E.; Tomsik, Thomas M.; Haberbusch, Mark S.

1995-01-01

Slush hydrogen, a mixture of solid and liquid hydrogen, offers advantages of higher density (16 percent) and higher heat capacity (18 percent) than normal boiling point hydrogen. The combination of increased density and heat capacity of slush hydrogen provided a potential to decrease the gross takeoff weight of the National Aero-Space Plane (NASP) and therefore slush hydrogen was selected as the propellant. However, no large-scale data was available on the production, transfer and tank pressure control characteristics required to use slush hydrogen as a fuel. Extensive testing has been performed at the NASA Lewis Research Center K-Site and Small Scale Hydrogen Test Facility between 1990 and the present to provide a database for the use of slush hydrogen. This paper summarizes the results of this testing.

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.