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Sample records for solid propellant rocket

  1. Solid propellant rocket motor

    NASA Technical Reports Server (NTRS)

    Dowler, W. L.; Shafer, J. I.; Behm, J. W.; Strand, L. D. (Inventor)

    1973-01-01

    The characteristics of a solid propellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solid propellant igniter and a progressive burning main solid propellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

  2. 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).

  3. A miniature solid propellant rocket motor

    SciTech Connect

    Grubelich, M.C.; Hagan, M.; Mulligan, E.

    1997-08-01

    A miniature solid-propellant rocket motor has been developed to impart a specific motion to an object deployed in space. This rocket motor effectively eliminated the need for a cold-gas thruster system or mechanical spin-up system. A low-energy igniter, an XMC4397, employing a semiconductor bridge was used to ignite the rocket motor. The rocket motor was ground-tested in a vacuum tank to verify predicted space performance and successfully flown in a Sandia National Laboratories flight vehicle program.

  4. 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.

  5. High-pressure burning rate studies of solid rocket propellants

    NASA Astrophysics Data System (ADS)

    Atwood, A. I.; Ford, K. P.; Wheeler, C. J.

    2013-03-01

    Increased rocket motor performance is a major driver in the development of solid rocket propellant formulations for chemical propulsion systems. The use of increased operating pressure is an option to improve performance potentially without the cost of reformulation. A technique has been developed to obtain burning rate data across a range of pressures from ambient to 345 MPa. The technique combines the use of a low loading density combustion bomb with a high loading density closed bomb technique. A series of nine ammonium perchlorate (AP) based propellants were used to demonstrate the use of the technique, and the results were compared to the neat AP burning rate "barrier". The effect of plasticizer, oxidizer particle size, catalyst, and binder type were investigated.

  6. 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.

  7. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    NASA Technical Reports Server (NTRS)

    Moore, Carleton J.

    1988-01-01

    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

  8. 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.

  9. 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.

  10. Lidar measurements of solid rocket propellant fire particle plumes.

    PubMed

    Brown, David M; Brown, Andrea M; Willitsford, Adam H; Dinello-Fass, Ryan; Airola, Marc B; Siegrist, Karen M; Thomas, Michael E; Chang, Yale

    2016-06-10

    This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air. PMID:27409023

  11. 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.

  12. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2014-11-01

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

  14. Terminology and assessment methods of solid propellant rocket exhaust signatures

    NASA Astrophysics Data System (ADS)

    1993-02-01

    The Propulsion and Energetics Panel's Specialists' Meeting in autumn 1985 on Smokeless Propellants demonstrated that no common standard was available in this field and that the lack of common understanding led to misunderstanding amongst the NATO community. After some preparatory discussion, the Panel, therefore, formed Working Group Number 21 with the objectives of defining methods for the assessment of rocket motor exhaust optical properties in the visible and in the infrared range, and of recommending a terminology based on quantitative criteria. The Working Group discussed the subject in a total of eight sessions and prepared this Advisory Report. Following an Introduction and Summary there are six chapters, commencing with an Overview and continuing with Propellant Smoke Classification, Plume Primary Smoke, Plume Secondary Smoke, Plume Radiation and Plume Microwave Properties. In most cases, the conclusions and recommendations follow the chapters and are not repeated at the end of the report.

  15. A facility for testing the acoustic combustion instability characteristics of solid rocket propellants

    NASA Technical Reports Server (NTRS)

    Mathes, H. B.

    1980-01-01

    A facility is described that has been specifically designed for small-scale laboratory testing of solid rocket propellants. A description of the facility is provided which includes the general plan of the facility and features related to personnel safety. One of the major activities in the facility is testing solid rocket propellants for combustion response to acoustic perturbations. A detailed discussion of acoustic instability testing is given including specially designed combustion apparatus, data acquisition, and signal conditioning. Techniques of data reduction are reviewed and some of the instrumentation problems that arise in this type of testing are mentioned along with practical solutions.

  16. Modeling of heat generation in ammonia-treated solid rocket propellant

    SciTech Connect

    Raun, R.L.; Isom, K.B.

    1995-06-01

    With the end of the Cold War, safe, environmentally sound separation, recycling, and disposal of ingredients in solid rocket propellants and munitions has become a national priority. One approach to demilitarize solid rocket propellants is treatment with ammonia. Ammonia extracts the oxidizers ammonium perchlorate and HMX, yielding a solid reside that is more suitable for incineration and less sensitive to impact and other modes of accidental initiation. Ammonia treatment of nitroglycerin-containing propellants is complicated by an exothermic reaction between ammonia and nitroglycerin. If not removed, the heat generated by this reaction can cause propellant ignition. To help design safe treatment processes, a model for the ammonia-propellant reaction was developed, which integrates transient energy and species conservation equations to simulate ammonia diffusion, heat generation, and heat flow in a propellant and in the solid residue resulting from ammonia treatment. It was calibrated using residue thickness and thermocouple data for one propellant. The calibrated model was used to predict conditions leading to ignition of thin propellant strips. The results agree well with experimental observations.

  17. Studies of the exhaust products from solid propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Dawbarn, R.; Kinslow, M.

    1976-01-01

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

  18. 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.

  19. 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.

  20. Atmospheric Manmade Glowings Phenomena Observed During the Launches of Solid Propellant Rockets

    NASA Astrophysics Data System (ADS)

    Chernouss, S. A.; Platov, V. V.; Upspensky, M. V.; Alpatov, V. V.; Kirillov, A. S.

    2015-09-01

    Exotic types of luminosities observed in the upper atmosphere always take place during the launch and flight of solid-propellant rockets We consider a large-scale geometry and dynamic features of such phenomena also physics of the intense turquoise (blue-green) glow observed in twilight conditions in the region of missile flight. This study has been based on numerous observations of different rocket flights in the atmosphere over Russia and Scandinavia. Formation of the monoxide aluminum clouds observed in the upper atmosphere is a result of interaction of the exhausted propellant products with the atomic oxygen. The sunlight excited the monoxide aluminum EA1O*) resonance emissions in the atmosphere. Careful studies of spectra of the manmade luminosities during rocket launch/flight permit us to know chemical, thermal and mechanical processes in the atmosphere similar as it is doing in experiments with the artificial cloud release from sounding rockets in the high latitude atmosphere.

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

    NASA Technical Reports Server (NTRS)

    Baran, W. J.

    1974-01-01

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

  2. Features of optical phenomena connected with launches of solid-propellant ballistic rockets

    NASA Astrophysics Data System (ADS)

    Platov, Yu. V.; Chernouss, S. A.; Alpatov, V. V.

    2013-04-01

    Specific optical phenomena observed in the upper atmosphere layers and connected with launches of powerful solid-propellant rockets are considered: the development of spherically symmetric gas-dust formations having the shape of an extending torus in the image plane and the formation of regions with intense blue-green (turquoise) glow observed under twilight conditions along a rocket's flight path. The development of clouds can be represented by the model of a strong explosion occurring at the stage separation of solid-propellant rockets in the upper atmosphere. A turquoise glow arises as a result of resonance scattering of solar radiation on AlO molecules that are formed when metallic aluminum in the composition of fuel interacts with atmosphere components and combustion products.

  3. Solid rocket propellant waste disposal/ingredient recovery study

    NASA Technical Reports Server (NTRS)

    Mcintosh, M. J.

    1976-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  5. Measured particulate behavior in a subscale solid propellant rocket motor

    NASA Astrophysics Data System (ADS)

    Brennan, W. D.; Hovland, D. L.; Netzer, D. W.

    1992-10-01

    Particulate matter are sized in the exhaust nozzle and plume of small rocket motors of varying geometry to assess the effects of the expansion process on particle size. Both converging and converging-diverging nozzles are considered, and particle sizing is accomplished at pressures of up to 4.36 MPa with aluminum loadings of 2.0 and 4.7 percent. An instrument based on Fraunhofer diffraction is used to measure the particle-size distributions showing that: (1) high burning rates reduce particle agglomeration and increase C* efficiency; (2) high pressures lead to small and monomodal D32 entering the nozzle; and (3) D32 sizes increase appreciably at the tailoff. Some variations in plume signature are theorized to be caused by the tailoff phenomenon, and particle collisions and/or surface effects in the nozzle convergence are suggested by the reduced number of larger particles at the nozzle convergence.

  6. 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.

  7. Casting propellant in rocket engine

    NASA Technical Reports Server (NTRS)

    Roach, J. E.; Froehling, S. C. (Inventor)

    1976-01-01

    A method is described for casting a solid propellant in the casing of a rocket engine having a continuous wall with a single opening which is formed by leaves of a material which melt at a temperature of the propellant and with curved edges concentric to the curvature of the spherical casing. The leaves are inserted into the spherical casing through the opening forming a core having a greater width than the width of the single opening and with curved peripheral edges. The cast propellant forms a solid mass and then heated to melt the leaves and provide a central opening with radial projecting flutes.

  8. 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.

  9. 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.

  10. Effect of silicone oil on nozzle heat transfer in solid propellant rockets

    NASA Technical Reports Server (NTRS)

    Ramohalli, K.

    1978-01-01

    The feasibility of reducing heat transfer rates to inert nozzle walls in solid rockets by addition of silicone oil as a propellant ingredient is explored experimentally. A metallized control propellant and its counterpart with 1% silicone oil are fired in a 2 1/2-in. diameter end burning all-metal burner. Pressure-time histories are obtained; temperature-time data at the nozzle throat are obtained in some of the runs. The heat loss by the product gases to the cold inert walls of the nozzle lower the temperature of the gases leading to the condensation of aluminum oxide on the walls. This condensation is used as an index of the heat transfer rate to the nozzle. On a statistical basis it is found that the silicone propellant gives practically no condensation in 70% of the firings contrasted with 21% for its counterpart with no silicone oil in it. The gas temperature at the nozzle throat is experimentally seen to be generally higher with the silicone propellant than with its counterpart; this higher temperature is consistent with a preliminary analysis that is used here to relate the nozzle surface temperature to the heat transfer from the product gases.

  11. Solid propellant motor

    NASA Technical Reports Server (NTRS)

    Shafer, J. I.; Marsh, H. E., Jr. (Inventor)

    1978-01-01

    A case bonded end burning solid propellant rocket motor is described. A propellant with sufficiently low modulus to avoid chamber buckling on cooling from cure and sufficiently high elongation to sustain the stresses induced without cracking is used. The propellant is zone cured within the motor case at high pressures equal to or approaching the pressure at which the motor will operate during combustion. A solid propellant motor with a burning time long enough that its spacecraft would be limited to a maximum acceleration of less than 1 g is provided by one version of the case bonded end burning solid propellant motor of the invention.

  12. Nonideal detonation and initiation behavior of a composite solid rocket propellant. [HMX/AP/Al

    SciTech Connect

    Dick, J.J.

    1981-01-01

    Shock initiation and detonation behavior of an HMX/AP/Al rocket propellant were studied for nonideal character. Low detonation velocities and unusual shock initiation behavior were observed. Failure to propagate steady detonation in cylinders of the propellant was also noted.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Rotational inviscid flow in laterally burning solid-propellant rocket motors

    NASA Astrophysics Data System (ADS)

    Balakrishnan, G.; Linan, A.; Williams, F. A.

    1992-12-01

    A theoretical analysis to determine the effects of mass addition on the inviscid but rotational and compressible flowfield in a porous duct with the injection rate dependent on the local pressure is performed for large ratios of length-to-duct diameter. The problem of describing the flow is reduced to the solution of a single integral equation. The ratio of specific heat gamma, and a constant pressure exponent n, measuring the dependence of the rate of mass injection on the local pressure, are the parameters of the solutions. The integral equation is solved numerically, and parametric results are presented for gamma, varying from 1 to 5/3 and for n varying from 0 to 1. A choking phenomenon is exhibited at a critical length of the duct in the vicinity of which the Mach number approaches unity. The choking condition, which is relevant to the operation of nozzleless solid-propellant rocket motors, is obtained parametrically in the present study and compared with corresponding results for irrotational, quasi-one-dimensional flow. The rotationality reduces the choking pressure.

  20. Acceleration effects on the performance of solid-propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Lucy, M. H.; Jones, I. W.; Stephens, M. V.

    1976-01-01

    Some acceleration effects on rocket performance have been well publicized. The dynamic process, characterized by marked increases in 'localized' burning rate, produces excessive case heating, slag retention, pressure buildup, and/or internal flow alterations. Data are presented illustrating drastic effects at low accelerations for sustainer type propellants and its relevance to several recent failures. Normalized orientation dependence of rate augmentation appears coupled to acceleration level and base burning rate. Effects appear influenced by propellant composition. Predictions using subscale motor data show good agreement with observed performance for ground spin and flight tests. Subscale test methods and results are also discussed.

  1. Silicone containing solid propellant

    NASA Technical Reports Server (NTRS)

    Ramohalli, K. N. R. (Inventor)

    1980-01-01

    The addition of a small amount, for example 1% by weight, of a liquid silicone oil to a metal containing solid rocket propellant provides a significant reduction in heat transfer to the inert nozzle walls. Metal oxide slag collection and blockage of the nozzle are eliminated and the burning rate is increased by about 5% to 10% thus improving ballistic performance.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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

    2016-11-01

    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. PMID:27289269

  8. 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.

  9. Uncertainties in the characterization of the thermal environment of a solid rocket propellant fire

    SciTech Connect

    Diaz, J.C.

    1993-10-01

    There has been an interest in developing models capable of predicting the response of systems to Minuteman (MM) III third-stage solid propellant fires. Input parameters for such an effort include the boundary conditions that describe the fire temperature, heat flux, emissivity, and propellant burn rate. In this study scanning spectroscopy and pyrometry were used to infer plume temperatures. Each diagnostic system possessed strengths and weaknesses. The intention was to use various supportive methods to infer plume temperature and emissivity, because no one diagnostic had proven capabilities for determining temperature under these conditions. Furthermore, these diagnostics were being used near the limit of their applicability. All these points created some uncertainty in the data collected.

  10. Liquid propellant rocket combustion instability

    NASA Technical Reports Server (NTRS)

    Harrje, D. T.

    1972-01-01

    The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

  11. 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.

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

    NASA Technical Reports Server (NTRS)

    1975-01-01

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

  13. 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.

  14. Nuclear Magnetic Resonance Imaging of Solid Rocket Propellants at 14.1 T

    NASA Astrophysics Data System (ADS)

    Maas, W. E.; Merwin, L. H.; Cory, D. G.

    1997-11-01

    Proton NMR images of solid propellant materials, consisting of a polybutadiene binder material filled with 82% solid particles, have been obtained at a magnetic field strength of 14.1 T and at a resolution of 8.5 × 8.5 μm. The images are the first of elastomeric materials obtained at a proton frequency of 600 MHz and have the highest spatial resolution yet reported. The images display a high contrast and are rich in information content. They reveal the distribution of individual filler particles in the polymer matrix as well as a thin polymer film of about 10-30 μm which is found to surround some of the larger filler particles.

  15. Numerical investigation of the influence of crystallization of ultrafine particles of aluminum oxide on energy characteristics of solid-propellant rocket engine

    NASA Astrophysics Data System (ADS)

    Dyachenko, N. N.; Dyachenko, L. I.

    2014-08-01

    The results of numerical investigation of a multiphase flow considering coagulation, crushing and crystallization of the particles of polydispersed condensate in the nozzles of solid-propellant rocket engine are presented. The influence of particles crystallization on the energy characteristics of the engine is shown.

  16. 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.

  17. 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.

  18. Asbestos Free Insulation Development for the Space Shuttle Solid Propellant Rocket Motor (RSRM)

    NASA Technical Reports Server (NTRS)

    Allred, Larry D.; Eddy, Norman F.; McCool, A. A. (Technical Monitor)

    2000-01-01

    Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

  19. 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.

  20. 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).

  1. Performance of a UTC FW-4S solid propellant rocket motor under the command effects of simulated altitude and rotational spin

    NASA Technical Reports Server (NTRS)

    Merryman, H. L.; Smith, L. R.

    1974-01-01

    One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.

  2. Solid propellant environmental issues

    SciTech Connect

    Le, M.D.

    1998-07-01

    The objective of the Solid Propellant Environmental Issues (SPEI) project is to demonstrate environmentally acceptable technologies that will enhance the continued production of solid rocket motors (SRMs) by complying with current and anticipated environmental regulations. Phase 1 of the project identifies current and anticipated environmental regulations that may affect SRMs manufacturing in the future and identify emerging process technologies which comply with these regulations. Phase 2 of the project established a baseline database by fabricating a 363 kg motor using the current manufacturing process. In Phase 3, environmentally acceptable process technologies were evaluated, ranked, and selected for demonstration using criteria developed by the team. The results for Phase 1--3 have previously been presented. This paper will present data obtained to date on Phase 4. In Phase 4, the alternate process technologies were evaluated for compatibility, cleaning effectiveness, and waste minimization/pollution prevention. The best performing candidate for each application area was selected for demonstration. The selected process technologies will be inserted into the baseline manufacturing process from Phase 2. The new manufacturing process will be demonstrated and evaluated through the scale-up and fabrication of two 363 kg solid rocket motors.

  3. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

    The paper describes the Advanced Solid Rocket Motor (ASRM) that is being developed to replace, in 1997, the Redesigned Solid Rocket Motor which currently boosts the Space Shuttle. The ASRM will contain features to improve motor safety (fewer potential leak paths, improved seal materials, stronger case material, and fewer nozzle and case joints), an improved ignition system using through-bulkhead initiators, and highly reproducible manufacturing and inspection techniques with a large number of automated procedures. The ASRM will be able to deliver 12,000 lbs greater payloads to any given orbit of the Shuttle. There are also environmental improvements, realized by waste propellant recovery.

  4. Feasibility of rocket propellant production on Mars

    NASA Technical Reports Server (NTRS)

    Ash, R. L.; Dowler, W. L.; Varsi, G.

    1978-01-01

    In situ production of rocket propellant to reduce landed mass requirements for Mars return missions has been investigated. The analysis has shown that a system which utilizes atmospheric carbon dioxide and soil moisture to produce liquid methane-oxygen propellant requires a landed mass which is less than half the mass of the ascent vehicle it produces.

  5. Rocket thrust variation with foamed liquid propellants

    NASA Technical Reports Server (NTRS)

    Morrell, G

    1957-01-01

    An analysis is presented on a method for varying rocket thrust by varying the bulk density of the propellants. This density variation was accomplished by uniformly dispersing an inert, insoluble gas in the liquid propellants. Only qualitative agreement with theory was obtained from preliminary experiments with a 1000-pound-thrust ammonia - nitric acid rocket engine; the required experimental gas-flow rates were two to six times greater than those predicted by theory. It was demonstrated, however, that this method of rocket-thrust variation is feasible.

  6. 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

  7. The dynamics of a gas-dust cloud expansion in the upper atmosphere at a shutdown of solid-propellant rocket engines

    NASA Astrophysics Data System (ADS)

    Nikolaishvili, S. Sh.; Platov, Yu. V.; Chernouss, S. A.

    2015-09-01

    The velocity of spherical gas-dust cloud expansion in the situation when the stages of solid-propellant rocket separate in the upper atmosphere have been determined. The measured velocity vary from 2.5 to 7.5 km/s. The dispersed component accelerates at the front of a shock that develops at engine-thrust shutdown. The model calculations of the gas-dust cloud luminosity intensity qualitatively coincide with the photometric profiles of object images. Such formations can vary from almost homogeneous ball-shaped clouds to rather thin spherical shells depending on the gas-dust cloud mass and the matter distribution within this cloud.

  8. Advanced Solid Rocket Launcher and Its Evolution

    NASA Astrophysics Data System (ADS)

    Morita, Yasuhiro; Imoto, Takayuki; Habu, Hiroto; Ohtsuka, Hirohito; Hori, Keiichi; Koreki, Takemasa; Fukuchi, Apollo; Uekusa, Yasuyuki; Akiba, Ryojiro

    The research on next generation solid propellant rockets is actively underway in various spectra. JAXA is developing the Advanced Solid Rocket (ASR) as a successor to the M-V launch vehicle, which was utilized over past ten years for space science programs including planetary missions. ASR is a result of the development of the next generation technology including a highly intelligent autonomous check-out system, which is connected to not only the solid rocket but also future transportation systems. It is expected to improve the efficiency of the launch system and double the cost performance. Far beyond this effort, the passion of the volunteers among the industry-government-academia cooperation has been united to establish the society of the freewheeling thinking “Next generation Solid Rocket Society (NSRS)”. It aims at a larger revolution than what the ASR provides so that the order of the cost performance is further improved. A study of the Low melting temperature Thermoplastic Propellant (LTP) is now at the experimental stage, which is expected to reform the manufacturing process of the solid rocket propellant and lead to a significant increase in cost performance. This paper indicates the direction of the big flow towards the next generation solid-propellant rockets: the concept of the intelligent ASR under development; and the innovation behind LTP.

  9. 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.

  10. Space aging of solid rocket materials

    NASA Technical Reports Server (NTRS)

    Lester, Dean M.; Jones, Leon L.; Smalley, R. B., Jr.; Ord, R. Neil

    1991-01-01

    Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. The results of aging IPSM-II/PAM-DII space motor components are presented. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results are compared to data collected in a ground based vacuum aging program on similar components.

  11. Space Shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Hardy, G. B.

    1979-01-01

    Details of the design, operation, testing and recovery procedures of the reusable solid rocket boosters (SRB) are given. Using a composite PBAN propellant, they will provide the primary thrust (six million pounds maximum at 20 s after ignition) within a 3 g acceleration constraint, as well as thrust vector control for the Space Shuttle. The drogues were tested to a load of 305,000 pounds, and the main parachutes to 205,000. Insulation in the solid rocket motor (SRM) will be provided by asbestos-silica dioxide filled acrylonitrile butadiene rubber ('asbestos filled NBR') except in high erosion areas (principally in the aft dome), where a carbon-filled ethylene propylene diene monomer-neopreme rubber will be utilized. Furthermore, twenty uses for the SRM nozzle will be allowed by its ablative materials, which are principally carbon cloth and silica cloth phenolics.

  12. Solid Rocket Booster-Illustration

    NASA Technical Reports Server (NTRS)

    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.

  13. The Advanced Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Mitchell, Royce E.

    1992-08-01

    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  14. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  15. Solid rocket motor temperature sensitivity

    SciTech Connect

    Osborn, J.R.; Heister, S.D.

    1994-11-01

    The temperature sensitivity of the propellant and the solid rocket motor are described by several different temperature sensitivity coefficients. This enabled the derivation of three different relationships for the temperature sensitivity coefficient pi(sub K). To demonstrate this, two different propellants were used wherein the values of pi(sub K) were generated and compared. It was observed that the expressions are of equal complexity and offer ease of use. All involve only the burning rate data and the use of the parameters in St. Roberts burning rate low. It is also suggested that the most general expression for the sensitivity coefficient should be used since it is a true pi(sub K) relationship having the partial derivatives taken with the motor geometry held constant. 11 refs.

  16. 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.

  17. 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.

  18. General view of the Solid Rocket Booster's (SRB) Solid Rocket ...

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

    General view of the Solid Rocket Booster's (SRB) Solid Rocket Motor Segments in the Surge Building of the Rotation Processing and Surge Facility at Kennedy Space Center awaiting transfer to the Vehicle Assembly Building and subsequent mounting and assembly on the Mobile Launch Platform. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  19. Thiokol Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  20. Solid rocket motors

    NASA Technical Reports Server (NTRS)

    Carpenter, Ronn L.

    1993-01-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  1. Characterization of the exhaust particulates in the ground cloud and high-altitude plume of large solid-propellant booster rockets

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Bowyer, J. M.; Varsi, G.; Laue, E. G.; Gauldin, R.

    1980-01-01

    The report is concerned with the characterization of Al2O3 particles in the atmosphere. These particles comprise one of the major combustion products of the rocket propellant employed in the Space Shuttle boosters. A ground cloud and stratospheric plume are considered. It is concluded that the typical residence times in the atmosphere are much longer than earlier estimates have indicated.

  2. 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.

  3. 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.

  4. 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.

  5. Advances in Green Cryogenic Solid Propellant Propulsion

    NASA Astrophysics Data System (ADS)

    Lo, R. E.; Adirim, H.; Poller, S.; Glaeser, S.; Schoeyer, H.; Caramelli, F.

    2004-10-01

    The combustion of hydrocarbons with hydrogen peroxide or oxygen based oxidizers is known as the best possible realization of green bipropellants in the realm of conventional propellants. By the nature of these constituents, corresponding rocket motors are either hybrids or bi-liquids. This is advantageous in all applications requiring the merits of these categories, such as variations of the thrust - time profile (throttle-ability up to shut down and restart), or variable propellant loading and mixture ratio variation in liquid bipropellants. However, when it comes to thriving on the simplicity and reliability of solid propellant technology, it takes cryogenic solid propulsion (CSP) as enabling technology to make these normally liquid propellants available for many solid propellant applications, in particular for high thrust Earth-to-orbit boosting. It is obvious that proper CSP propellant selection yields solids that are as "green" as any chemical propellant combination can be. The paper describes recent advances in CSP technology related investigations sponsored by the German Aerospace Centre DLR and the European Space Agency ESA at AI/ICT.

  6. Ignition transient analysis of solid rocket motor

    NASA Astrophysics Data System (ADS)

    Han, Samuel S.

    A 1-D numerical model based on the SIMPLE is developed to predict the pressure and thrust behavior of space shuttle solid rocket motors. The present model solves the conservation equations through the attached nozzle as well as in the combustion chamber. Numerical results were seen to agree qualitatively well with the test data by controlling the wetted perimeter in the head-end star-section of the motor and the erosive burning rate of the solid propellent.

  7. 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.

  8. Ignition transient analysis of solid rocket motor

    NASA Technical Reports Server (NTRS)

    Han, Samuel S.

    1991-01-01

    Measurement data on the performance of Space Shuttle Solid Rocket Motor show wide variations in the head-end pressure changes and the total thrust build-up during the ignition transient periods. To analyze the flow and thermal behavior in the tested solid rocket motors, a 1-dimensional, ideal gas flow model via the SIMPLE algorithm was developed. Numerical results showed that burning patterns in the star-shaped head-end segment of the propellant and the erosive burning rate are two important factors controlling the ignition transients. The objective of this study is to extend the model to include the effects of aluminum particle commonly used in solid propellants. To treat the effects of aluminum-oxide particles in the combustion gas, conservation of mass, momentum, and energy equations for the particles are added in the numerical formulation and integrated by an inter-phase-slip algorithm.

  9. The development of space solid rocket motors in China

    NASA Astrophysics Data System (ADS)

    Jianding, Huang; Dingyou, Ye

    1997-01-01

    China has undertaken to research and develop composite solid propellant rocket motors since 1958. At the request of the development of space technology, composite solid propellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

  10. Nuclear thermal rockets using indigenous extraterrestrial propellants

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert M.

    1990-01-01

    A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS.

  11. 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.

  12. Space aging of solid rocket materials

    NASA Technical Reports Server (NTRS)

    Lester, Dean M.; Jones, Leon L.; Smalley, R. B., Jr.; Ord, R. Neil

    1992-01-01

    Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. This paper will present the results of aging the Improved Performance Space Motor-II/Payload Assist Module-Delta II (IPSM-II/PAM-DII) space motor components. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results will be compared to data collected in a ground based vacuum aging program on similar components. The operation of the vacuum actuated venting valve and pressure actuated resealing of the container will also be addressed. The materials tested showed no significant changes due to space aging. These results indicate that properly designed solid rocket motors can be expected to perform reliably after extended periods of exposure to a space environment.

  13. 148. SKID 2 FOR LOADING ROCKET PROPELLANT AT EAST SIDE ...

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

    148. SKID 2 FOR LOADING ROCKET PROPELLANT AT EAST SIDE OF FUEL CONTROL ROOM (215), LSB (BLDG. 751) - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

  14. On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

    NASA Technical Reports Server (NTRS)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

  15. 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.

  16. Processing solid propellants for recycling

    SciTech Connect

    Whinnery, L.L.; Griffiths, S.K.; Handrock, J.L.; Lipkin, J.

    1994-05-01

    Rapid evolution in the structure of military forces worldwide is resulting in the retirement of numerous weapon systems. Many of these systems include rocket motors containing highly energetic propellants based on hazardous nitrocellulose/nitroglycerin (NC/NG) mixtures. Even as the surplus quantities of such material increases, however, current disposal methods -- principally open burning and open detonation (OB/OD) -- are coming under close scrutiny from environmental regulators. Environmentally conscious alternatives to disposal of propellant and explosives are thus receiving renewed interest. Recycle and reuse alternatives to OB/OD appear particularly attractive because some of the energetic materials in the inventories of surplus weapon systems represent potentially valuable resources to the commercial explosives and chemical industries. The ability to reclaim such resources is therefore likely to be a key requirement of any successful technology of the future in rocket motor demilitarization. This document consists of view graphs from the poster session.

  17. NASA's Advanced solid rocket motor

    NASA Astrophysics Data System (ADS)

    Mitchell, Royce E.

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

  18. NASA's Advanced solid rocket motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1993-01-01

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

  19. Small Solid Rocket Motor Test

    NASA Video Gallery

    It was three-two-one to brilliant fire as NASA's Marshall Space Flight Center tested a small solid rocket motor designed to mimic NASA's Space Launch System booster. The Mar. 14 test provides a qui...

  20. 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.

  1. Solid Hydrogen Formed for Atomic Propellants

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan A.

    2000-01-01

    Several experiments on the formation of solid hydrogen particles in liquid helium were recently conducted at the NASA Glenn Research Center at Lewis Field. The solid hydrogen experiments are the first step toward seeing these particles and determining their shape and size. The particles will ultimately store atoms of boron, carbon, or hydrogen, forming an atomic propellant. Atomic propellants will allow rocket vehicles to carry payloads many times heavier than possible with existing rockets or allow them to be much smaller and lighter. Solid hydrogen particles are preferred for storing atoms. Hydrogen is generally an excellent fuel with a low molecular weight. Very low temperature hydrogen particles (T < 4 K) can prevent the atoms from recombining, making it possible for their lifetime to be controlled. Also, particles that are less than 1 mm in diameter are preferred because they can flow easily into a pipe when suspended in liquid helium. The particles and atoms must remain at this low temperature until the fuel is introduced into the engine combustion (or recombination) chamber. Experiments were, therefore, planned to look at the particles and observe their formation and any changes while in liquid helium.

  2. Dynamic characterization and analysis of space shuttle SRM solid propellant

    NASA Technical Reports Server (NTRS)

    Hufferd, W. L.

    1979-01-01

    The dynamic response properties of the space shuttle solid rocket moter (TP-H1148) propellant were characterized and the expected limits of propellant variability were established. Dynamic shear modulus tests conducted on six production batches of TP-H1148 at various static and dynamic strain levels over the temperature range from 40 F to 90 F. A heat conduction analysis and dynamic response analysis of the space shuttle solid rocket motor (SRM) were also conducted. The dynamic test results show significant dependence on static and dynamic strain levels and considerable batch-to-batch and within-batch variability. However, the results of the SRM dynamic response analyses clearly demonstrate that the stiffness of the propellant has no consequential on the overall SRM dynamic response. Only the mass of the propellant needs to be considered in the dynamic analysis of the space shuttle SRM.

  3. Solid Rocket Motor Acoustic Testing

    SciTech Connect

    Rogers, J.D.

    1999-03-31

    Acoustic data are often required for the determination of launch and powered flight loads for rocket systems and payloads. Such data are usually acquired during test firings of the solid rocket motors. In the current work, these data were obtained for two tests at a remote test facility where we were visitors. This paper describes the data acquisition and the requirements for working at a remote site, interfacing with the test hosts.

  4. Propellant-Flow-Actuated Rocket Engine Igniter

    NASA Technical Reports Server (NTRS)

    Wollen, Mark

    2013-01-01

    A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0

  5. 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.

  6. 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.

  7. 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.

  8. Erosive burning of solid propellants

    NASA Technical Reports Server (NTRS)

    King, Merrill K.

    1993-01-01

    Presented here is a review of the experimental and modeling work concerning erosive burning of solid propellants (augmentation of burning rate by flow of product gases across a burning surface). A brief introduction describes the motor design problems caused by this phenomenon, particularly for low port/throat area ratio motors and nozzleless motors. Various experimental techniques for measuring crossflow sensitivity of solid propellant burning rates are described, with the conclusion that accurate simulation of the flow, including upstream flow development, in actual motors is important since the degree of erosive burning depends not only on local mean crossflow velocity and propellant nature, but also upon this upstream development. In the modeling area, a brief review of simplified models and correlating equations is presented, followed by a description of more complex numerical analysis models. Both composite and double-base propellant models are reviewed. A second generation composite model is shown to give good agreement with data obtained in a series of tests in which composite propellant composition and heterogeneity (particle size distribution) were systematically varied. Finally, the use of numerical models for the development of erosive burning correlations is described, and a brief discussion of scaling is presented.

  9. Dragonfly directional sensor versus rocket-propelled grenades

    NASA Astrophysics Data System (ADS)

    Geary, Joseph; Blackwell, Lisa

    2015-02-01

    The Dragonfly directional sensor was deployed at the Army's Yuma Proving Grounds for preliminary field tests against rocket-propelled grenades. This wide-field (nonimaging) sensor's purpose was to angularly locate the latter's launch plume. These tests successfully demonstrated proof-of-concept.

  10. Motion Analysis of a Rocket-Propelled Truck.

    ERIC Educational Resources Information Center

    Hitt, Darren L.; Lowe, Mary L.

    1996-01-01

    Describes an experiment to study the motion of a rocket-propelled vehicle over the entire duration of the engine burn using a video system with a frame-by-frame playback and a Sonic Ranger for ultrasonic position movements. Enables students to study the impulse-momentum principle and the effects of a time-varying force. (JRH)

  11. Solid rocket technology advancement for Space Tug and IUS applications. [Interim Upper Stage

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Two-burn restartable solid propellant rocket motors for the kick stage (auxiliary stage) of the Shuttle Tug, or Interim Upper Stage, are described, with details on features and test results of the ignition and quench (thrust termination) systems and procedures, fabrication of propellant and insulation, explosion hazards of propellants, and comparative data on present and future motor design. These rocket motor systems are designed for upper stage augmentation of launch vehicles and possible service in Shuttle-launched outer planet spacecraft.

  12. 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.

  13. Studies of solid propellant combustion with pulsed radiography

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

    Pulsed radiography was applied to observe solid propellant surface regression during rocket motor operation. Using a 150 KV flash X-ray system manufactured by the Field Emission Corporation and two kinds of film suppliers, images of the propellant surface of a 5 cm diameter end burning rocket motor were recorded on film. The repetition frame rate of 8 pulses per second and the pulse train length of 10 pulses are limited by the capability of the power supply and the heat build up within the X-ray tube, respectively. The experiment demonstrated the effectiveness of pulsed radiography for observing solid propellant surface regression. Measuring the position of burning surface images on film with a microdensitometer, quasi-instantaneous burning rate as a function of pressure and the variation of characteristic velocity with pressure and gas stay time were obtained. Other research items to which pulsed radiography can be applied are also suggested.

  14. Metallic Hydrogen: A Game Changing Rocket Propellant

    NASA Technical Reports Server (NTRS)

    Silvera, Isaac F.

    2016-01-01

    The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

  15. 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.

  16. Study of solid rocket motor for space, shuttle booster, volume 2, book 4 appendices B thru D

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The mass properties and related data for the solid propellant rocket engine for use with the space shuttle are presented. Data for three solid propellant rocket engines are provided. The three designs considered are: (1) baseline parallel burn, (2) optional parallel burn, and (3) baseline series burn. Layouts of the respective designs to show design and dimensional data are included.

  17. Solid rocket motor witness test

    NASA Technical Reports Server (NTRS)

    Welch, Christopher S.

    1991-01-01

    The Solid Rocket Motor Witness Test was undertaken to examine the potential for using thermal infrared imagery as a tool for monitoring static tests of solid rocket motors. The project consisted of several parts: data acquisition, data analysis, and interpretation. For data acquisition, thermal infrared data were obtained of the DM-9 test of the Space Shuttle Solid Rocket Motor on December 23, 1987, at Thiokol, Inc. test facility near Brigham City, Utah. The data analysis portion consisted of processing the video tapes of the test to produce values of temperature at representative test points on the rocket motor surface as the motor cooled down following the test. Interpretation included formulation of a numerical model and evaluation of some of the conditions of the motor which could be extracted from the data. These parameters included estimates of the insulation remaining following the tests and the thickness of the charred layer of insulation at the end of the test. Also visible was a temperature signature of the star grain pattern in the forward motor segment.

  18. Experimental investigation of a solid rocket combustion simulator

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1991-01-01

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

  19. Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

    NASA Technical Reports Server (NTRS)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

  20. Overview of the manufacturing sequence of the Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Chapman, John S.; Nix, Michael B.

    1992-01-01

    The manufacturing sequence of NASA's new Advanced Solid Rocket Motor, developed as a replacement of the Space Shuttle's existing Redesigned Solid Rocket Motor, is overviewed. Special attention is given to the case preparation, the propellant mix/cast, the nondestructuve evaluation, the motor finishing, and the refurbishment. The fabrication sequences of the case, the nozzle, and the igniter are described.

  1. Rocket Propellants Engine Design/Operations/Validation

    NASA Technical Reports Server (NTRS)

    Monk, Jan C.

    2002-01-01

    Lockheed Martin Astronautics Operations (LMA) was competitively awarded a contract May 21, 2001 for next generation launch system architecture definition and technology maturation. The Second Generation Launch Vehicle Program objectives include reducing the technical and programmatic risk of proceeding to full scale development of the system by establishing requirements for the next generation launch system and maturing critical technologies needed by the system. LMA will conduct analyses and trades to optimize the architecture ETO elements including configuration, conceptual designs, and preliminary operations definition. To fully understand the engine and propellant trades were conducted by LMA to yield the optimized architecture system from the operability, reliability, safety, and cost perspectives. A government/industry team addressed the required trade studies, the parameters and weighting factors, and the most critical trades were addressed. This report summarizes the participation of JCM Consulting, Inc. in the propellant trade study.

  2. Fluid thrust control system. [for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Howell, W. L.; Jansen, H. B.; Lehmann, E. N. (Inventor)

    1968-01-01

    A pure fluid thrust control system is described for a pump-fed, regeneratively cooled liquid propellant rocket engine. A proportional fluid amplifier and a bistable fluid amplifier control overshoot in the starting of the engine and take it to a predetermined thrust. An ejector type pump is provided in the line between the liquid hydrogen rocket nozzle heat exchanger and the turbine driving the fuel pump to aid in bringing the fluid at this point back into the regular system when it is not bypassed. The thrust control system is intended to function in environments too severe for mechanical controls.

  3. 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.

  4. Adsorption and chemical reaction of gaseous mixtures of hydrogen chloride and water on aluminum oxide and application to solid-propellant rocket exhaust clouds

    NASA Technical Reports Server (NTRS)

    Cofer, W. R., III; Pellett, G. L.

    1978-01-01

    Hydrogen chloride (HCl) and aluminum oxide (Al2O3) are major exhaust products of solid rocket motors (SRM). Samples of calcination-produced alumina were exposed to continuously flowing mixtures of gaseous HCl/H2O in nitrogen. Transient sorption rates, as well as maximum sorptive capacities, were found to be largely controlled by specific surface area for samples of alpha, theta, and gamma alumina. Sorption rates for small samples were characterized linearly with an empirical relationship that accounted for specific area and logarithmic time. Chemisorption occurred on all aluminas studied and appeared to form from the sorption of about a 2/5 HCl-to-H2O mole ratio. The chemisorbed phase was predominantly water soluble, yielding chloride/aluminum III ion mole ratios of about 3.3/1 suggestive of dissolved surface chlorides and/or oxychlorides. Isopiestic experiments in hydrochloric acid indicated that dissolution of alumina led to an increase in water-vapor pressure. Dissolution in aqueous SRM acid aerosol droplets, therefore, might be expected to promote evaporation.

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

    NASA Astrophysics Data System (ADS)

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

    2016-02-01

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

  6. Characterization of rocket propellant combustion products

    SciTech Connect

    Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.

    1991-12-09

    The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

  7. Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

    NASA Technical Reports Server (NTRS)

    Meyer, Mike L.

    1993-01-01

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

  8. 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.

  9. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

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

  10. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

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

  11. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

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

  12. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

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

  13. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

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

  14. Development and Test of a Rocket Engine Using Environmentally Friendly Propellants

    NASA Technical Reports Server (NTRS)

    Webster, Kristi

    2009-01-01

    Develop and test a rocket engine that operates on environmentally friendly propellants; Liquid Oxygen (LOX) and Liquid Methane (LCH4). Due to modifications the rocket engine designed last summer (KJ_REX) is not the same rocket thruster tested this summer, but very similar. The new modified rocket thruster was built for NASA by Orion Propulsion Inc. (OPI), Huntsville, AL.

  15. ELIMINATION OF TOXIC MATERIALS AND SOLVENTS FROM SOLID PROPELLANT COMPONENTS (PP1058)

    EPA Science Inventory

    Solid rocket motor (srm) propellant used in military tactical and strategic missile weapon systems contain hazardous and toxic materials. Lead is used as a ballistic catalyst in minimum signature propellants. It's use results in approximately 161,000 pounds of lead compound usag...

  16. Radial slot flows in solid rocket motors

    NASA Astrophysics Data System (ADS)

    Hilbing, J. H.; Heister, S. D.

    1993-06-01

    A series of parametric numerical solutions have been generated to characterize the two-dimensional flowfield due to the presence of a radial slot in a solid rocket propellant grain. Results have been parameterized in terms of upstream core Mach number, slot contraction ratio, and slot-to-core mass flow and momentum ratios. Numerical solutions of the axisymmetric Euler equations have been obtained on a 'generic' slot geometry using a cell-centered, finite volume scheme. Results indicate that both the stagnation pressure loss and grain suction force on the propellant segment downstream of the slot correlate well with slot-to-core momentum ratio; a parameter which has not been used in previous studies. Significant differences (in stagnation pressure losses) have been identified between the 2-D numerical results and the 1-D methods applied in current state-of-the-art ballistics codes. We anticipate that the correlations derived through this parametric study can be used in preliminary performance and grain stress analyses performed during the motor development process.

  17. 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.

  18. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  19. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2007-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  20. Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2008-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  1. Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  2. Materials Problems in Chemical Liquid-Propellant Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, L. L.

    1959-01-01

    With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.

  3. Boundary cooled rocket engines for space storable propellants

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  4. Amateur Gas-Propelled Rocket Engine Development and Advanced Rocket Design

    NASA Astrophysics Data System (ADS)

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

    The paper describes the design and manufacturing of a gaseous propellant rocket engine. It is an undertaking of the authors, performed on project basis with fellow aerospace engineering students under auspices of DARE (Delft Aerospace Rocket Engineering). This paper describes the requirements, the engine development, and the design considerations and calculations as they were performed. Furthermore, the plans for engine tests and the parameters that will have to be measured during those tests are covered. The design process converged to a 1800 N thrust gaseous oxygen/methane (GOX/CH4) engine made of electrolytic copper. GOX/CH4 was selected based on its relatively high specific impulse, its availability and because of its potential as a green propellant. A test engine was produced with a specific impulse of 287 s and a propellant mass flow of 637 g/s. From a point of view of strength, the focus was mainly on robustness rather than light weight. The main aim now is to perform tests with the current engine, based on which the performance can be verified and vital information for future design efforts can be acquired. The ultimate goal is to have an operational rocket and to attempt an amateur altitude record.

  5. Acoustic Measurements for Small Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Vargas, Magda B.; Kenny, R. Jeremy

    2010-01-01

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

  6. 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.

  7. 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.

  8. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    NASA Technical Reports Server (NTRS)

    Crawford, T. Daniel

    2002-01-01

    The design of new high-energy density molecules as candidates for jet and rocket fuels is an important goal of modern chemical thermodynamics. The NASA Glenn Research Center is home to a database of thermodynamic data for over 2000 compounds related to this goal, in the form of least-squares fits of heat capacities, enthalpies, and entropies as functions of temperature over the range of 300 - 6000 K. The chemical equilibrium with applications (CEA) program written and maintained by researchers at NASA Glenn over the last fifty years, makes use of this database for modeling the performance of potential rocket propellants. During its long history, the NASA Glenn database has been developed based on experimental results and data published in the scientific literature such as the standard JANAF tables. The recent development of efficient computational techniques based on quantum chemical methods provides an alternative source of information for expansion of such databases. For example, it is now possible to model dissociation or combustion reactions of small molecules to high accuracy using techniques such as coupled cluster theory or density functional theory. Unfortunately, the current applicability of reliable computational models is limited to relatively small molecules containing only around a dozen (non-hydrogen) atoms. We propose to extend the applicability of coupled cluster theory- often referred to as the 'gold standard' of quantum chemical methods- to molecules containing 30-50 non-hydrogen atoms. The centerpiece of this work is the concept of local correlation, in which the description of the electron interactions- known as electron correlation effects- are reduced to only their most important localized components. Such an advance has the potential to greatly expand the current reach of computational thermochemistry and thus to have a significant impact on the theoretical study of jet and rocket propellants.

  9. Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

    NASA Technical Reports Server (NTRS)

    Casiano, Matthew; Hulka, James; Yang, Virog

    2009-01-01

    Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

  10. Modeling of nonlinear longitudinal instability in solid rocket motors

    NASA Astrophysics Data System (ADS)

    Baum, Joseph D.; Levine, Jay N.

    A comprehensive model of nonlinear longitudinal combustion instability in solid rocket motors has been developed. The two primary elements of this stability analysis are a finite difference solution of the two phase flow in the combustion chamber and a coupled solution of the nonlinear transient propellant burning rate. A new combination finite difference scheme gives the analysis the ability to treat the type of multiple travelling shock wave instabilities that are frequently observed in reduced smoke tactical solid rocket motors. Models for predicting the behavior of both gas ejection and solid ejecta pulses were developed and incorporated into the analysis. Extensive comparisons between model predictions and experimental data from pulsed solid rocket motor firings have been carried out. The nonlinear instability analysis was found to be capable of predicting the complete range of nonlinear behavior observed in actual motor firing data. Good agreement between measured and predicted initial pulse amplitude, pulse evolution, limit cycle amplitude and mean pressure shift was obtained. This investigation has also provided new insight into the nature of nonlinear pulse triggered instability and the factors which influence its occurrence and severity. This new instability analysis should significantly enhance our capability to design tactical solid rocket motors that are free from troublesome and expensive nonlinear combusion instability problems.

  11. Ignition transient analysis of solid rocket motor

    NASA Technical Reports Server (NTRS)

    Han, Samuel S.

    1990-01-01

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

  12. Acoustic Measurements of Small Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Vargas, Magda B.; Kenny, R. Jeremy

    2010-01-01

    Rocket acoustic noise can induce loads and vibration on the vehicle as well as the surrounding structures. Models have been developed to predict these acoustic loads based on scaling existing solid rocket motor data. The NASA Marshall Space Flight Center acoustics team has measured several small solid rocket motors (thrust below 150,000 lbf) to anchor prediction models. This data will provide NASA the capability to predict the acoustic environments and consequent vibro-acoustic response of larger rockets (thrust above 1,000,000 lbf) such as those planned for the NASA Constellation program. This paper presents the methods used to measure acoustic data during the static firing of small solid rocket motors and the trends found in the data.

  13. 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.

  14. NASA, ATK Successfully Test Solid Rocket Motor

    NASA Video Gallery

    With a loud roar and mighty column of flame, NASA and ATK Aerospace Systems successfully completed a two-minute, full-scale test of the largest and most powerful solid rocket motor designed for fli...

  15. Theoretical performance of lithium and fluorine as a rocket propellant

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Huff, Vearl N

    1951-01-01

    Theoretical performance for liquid lithium and liquid fluorine as a rocket propellant was calculated with assumptions both of equilibrium and frozen composition during expansion. Parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, composition, mean molecular weight, characteristic velocity, coefficient of thrust, and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, the maximum equilibrium specific impulse calculated was 335.5 pound-seconds per pound. The effect of ionization on calculated performance was shown to be negligible by comparison of values of various parameters calculated both with and without ionized products of combustion.

  16. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James

    2008-01-01

    The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

  17. Solid rocket booster retrieval operations

    NASA Technical Reports Server (NTRS)

    Rasmussen, A. M.

    1985-01-01

    Solid Rocket Booster Retrieval operations are discussed in detail. The recovery of expended boosters and associated hardware without damage attributable to retrieval procedures is the main goal. The retrieval force consists tof ship's personnel and retrieval team members, each of whom has been trained and is highly skilled in multi-faceted operations. The retrieval force is equipped with two specially-built, highly maneuverable ships outfitted with parachute reels, retrieval cranes, towing winches, large volume-low pressure air compressors, SCUBA diving gear, inflatable boats with outboard motors and diver-operated SRB dewatering devices. The two ships are deployed in sufficient time to conduct an electronic and visual search of the impact area prior to launch. Upon search completeion, each ship takes station a safe distance from the predetermined impact point initiating both visual and electronic search in the direction of flight path, ensuring SRB acquisition at splashdown. When safe, the ships enter the impact area and commence recovery of all floating flight hardware which is subsequently returned to the Disassembly Facility for refurbishment and reuse. Retrieval techniques have evolved in parallel with equipment and flight hardware configuration changes. Additional changes have been initiated to improve personnel safety.

  18. Solid rocket motor cost model

    NASA Technical Reports Server (NTRS)

    Harney, A. G.; Raphael, L.; Warren, S.; Yakura, J. K.

    1972-01-01

    A systematic and standardized procedure for estimating life cycle costs of solid rocket motor booster configurations. The model consists of clearly defined cost categories and appropriate cost equations in which cost is related to program and hardware parameters. Cost estimating relationships are generally based on analogous experience. In this model the experience drawn on is from estimates prepared by the study contractors. Contractors' estimates are derived by means of engineering estimates for some predetermined level of detail of the SRM hardware and program functions of the system life cycle. This method is frequently referred to as bottom-up. A parametric cost analysis is a useful technique when rapid estimates are required. This is particularly true during the planning stages of a system when hardware designs and program definition are conceptual and constantly changing as the selection process, which includes cost comparisons or trade-offs, is performed. The use of cost estimating relationships also facilitates the performance of cost sensitivity studies in which relative and comparable cost comparisons are significant.

  19. The use of electrical discharge for ignition and control of combustion of solid propellants

    NASA Technical Reports Server (NTRS)

    Tachibana, Takeshi; Kobayashi, Tsuruo; Matsuda, Takashi; Kimura, Itsuro

    1987-01-01

    As the first step of the study of the combustion control of solid propellants by electrical discharges, the effects of an arc discharge, which flows along the burning surface, on the burning rate and on the increase of enthalpy of the combustion product were investigated. For specially devised composite propellants, which are composed of Al and Teflon powders, it was shown that the combination can be controlled by an arc discharge; the combustion continues when the arc discharge is applied and is interrupted when the arc discharge breaks. In the present investigation, it was also shown that an arc discharge coupled with a high-frequency electrical discharge has potential as an effective ignition method for solid propellants. For the application of this type of combustion control to an ignitor for a solid propellant rocket motor or to a control rocket motor, this method lacks flexibility in the configuration scale and needs relatively high electric power at the present stage.

  20. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    NASA Astrophysics Data System (ADS)

    Elliott, T. S.; Majdalani, J.

    2014-11-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion.

  1. Shuttle solid rocket motor nozzle alternate ablative evaluation

    NASA Technical Reports Server (NTRS)

    Powers, L. B.; Bailey, R. L.; Morrison, B. H.

    1981-01-01

    A series of subscale tests are shown to suggest that a lower-cost ablative material than the rayon-based carbon ablative currently used in the Space Shuttle Solid Rocket Motor (SRM) may be used as a substitute. Six such ablatives with outstanding performance characteristics, using spun PAN and continuous pitch and PAN fibers instead of the present, continuous rayon, were identified in the course of tests with HTPB/AL/AP solid propellant grains with a burn time of 12 sec. The test nozzle features an initial throat diameter of 2.2 in. and a 6.1 expansion ratio. In addition to nozzle structural feature drawings, extensive test data tables and propellant formulation and properties tables are provided.

  2. Infrared Imagery of Solid Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Moran, Robert P.; Houston, Janice D.

    2011-01-01

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

  3. High burn rate solid composite propellants

    NASA Astrophysics Data System (ADS)

    Manship, Timothy D.

    High burn rate propellants help maintain high levels of thrust without requiring complex, high surface area grain geometries. Utilizing high burn rate propellants allows for simplified grain geometries that not only make production of the grains easier, but the simplified grains tend to have better mechanical strength, which is important in missiles undergoing high-g accelerations. Additionally, high burn rate propellants allow for a higher volumetric loading which reduces the overall missile's size and weight. The purpose of this study is to present methods of achieving a high burn rate propellant and to develop a composite propellant formulation that burns at 1.5 inches per second at 1000 psia. In this study, several means of achieving a high burn rate propellant were presented. In addition, several candidate approaches were evaluated using the Kepner-Tregoe method with hydroxyl terminated polybutadiene (HTPB)-based propellants using burn rate modifiers and dicyclopentadiene (DCPD)-based propellants being selected for further evaluation. Propellants with varying levels of nano-aluminum, nano-iron oxide, FeBTA, and overall solids loading were produced using the HTPB binder and evaluated in order to determine the effect the various ingredients have on the burn rate and to find a formulation that provides the burn rate desired. Experiments were conducted to compare the burn rates of propellants using the binders HTPB and DCPD. The DCPD formulation matched that of the baseline HTPB mix. Finally, GAP-plasticized DCPD gumstock dogbones were attempted to be made for mechanical evaluation. Results from the study show that nano-additives have a substantial effect on propellant burn rate with nano-iron oxide having the largest influence. Of the formulations tested, the highest burn rate was a 84% solids loading mix using nano-aluminum nano-iron oxide, and ammonium perchlorate in a 3:1(20 micron: 200 micron) ratio which achieved a burn rate of 1.2 inches per second at 1000

  4. 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.

  5. Binary Solid Propellants for Constant Momentum Missions

    SciTech Connect

    Pakhomov, Andrew V.; Mahaffy, Kevin E.

    2008-04-28

    A constant momentum mission is achieved when the speed of the vehicle in the inertial frame of reference is equal to the speed of exhaust relative to the vehicle. Due to 100% propulsive efficiency such missions are superior to traditional constant specific impulse missions. A new class of solid binary propellants for constant momentum missions is under development. A typical propellant column is prepared as a solid solution of two components, with composition gradually changing from 100% of a propellant of high coupling coefficient (C{sub m}) to one which has high specific impulse (I{sub sp}). The high coupling component is ablated first, gradually giving way to the high I{sub sp} component, as the vehicle accelerates. This study opens new opportunities for further design of complex propellants for laser propulsion, providing variable C{sub m} and I{sub sp} during missions.

  6. 24 Inch Reusable Solid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    2002-01-01

    A scaled-down 24-inch version of the Space Shuttle's Reusable Solid Rocket Motor was successfully fired for 21 seconds at a Marshall Space Flight Center (MSFC) Test Stand. The motor was tested to ensure a replacement material called Lycocel would meet the criteria set by the Shuttle's Solid Motor Project Office. The current material is a heat-resistant, rayon-based, carbon-cloth phenolic used as an insulating material for the motor's nozzle. Lycocel, a brand name for Tencel, is a cousin to rayon and is an exceptionally strong fiber made of wood pulp produced by a special 'solvent-spirning' process using a nontoxic solvent. It will also be impregnated with a phenolic resin. This new material is expected to perform better under the high temperatures experienced during launch. The next step will be to test the material on a 48-inch solid rocket motor. The test, which replicates launch conditions, is part of Shuttle's ongoing verification of components, materials, and manufacturing processes required by MSFC, which oversees the Reusable Solid Rocket Motor project. Manufactured by the ATK Thiokol Propulsion Division in Promontory, California, the Reusable Solid Rocket Motor measures 126 feet (38.4 meters) long and 12 feet (3.6 meters) in diameter. It is the largest solid rocket motor ever flown and the first designed for reuse. During its two-minute burn at liftoff, each motor generates an average thrust of 2.6 million pounds (1.2 million kilograms).

  7. Ignition transient analysis of a solid rocket motor using a one dimensional two fluid model

    NASA Astrophysics Data System (ADS)

    Pardue, Byron A.; Han, Samuel S.

    1992-07-01

    A one dimensional two fluid numerical model has been used to study the ignition transient stage of a Space Shuttle solid rocket motor. During the ignition phase of a solid rocket motor a pressure transient is induced by complex transport processes involving the igniter gas heat transfer to the propellant, chemical reactions at the propellant surface, and the interaction of the fluid with the attached rocket nozzle. One dimensional models used in the past neglected the aluminum oxide particles which are present in the combustion gases. The current model uses the IPSA (Inter-Phase-Slip-Algorithm) to solve the transient compressible flow equations for the rocket chamber and attached nozzle. Numerical results for head end pressure changes and overall thrust are compared with both measurement data and predictions of a one dimensional one fluid model.

  8. Numerical simulation of a liquid propellant rocket motor

    NASA Astrophysics Data System (ADS)

    Salvador, Nicolas M. C.; Morales, Marcelo M.; Migueis, Carlos E. S. S.; Bastos-Netto, Demétrio

    2001-03-01

    This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems. This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables, the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and external flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

  9. Experiments in thermosensitive cavitation of a cryogenic rocket propellant surrogate

    NASA Astrophysics Data System (ADS)

    Kelly, Sean Benjamin

    Cavitation is a phase-change phenomenon that may appear in practical devices, often leading to loss of performance and possible physical damage. Of particular interest is the presence of cavitation in rocket engine pumps as the cryogenic fluids cavitate in impellers and inducers. Unlike water, which has been studied exhaustively, cryogenic fluids undergo cavitation with significant thermal effect. Past attempts at analyzing this behavior in water have led to poor predictive capability due to the lack of data in the regime defined as thermosensitive cavitation. Fluids flowing near their thermodynamic critical point have a liquid-vapor density ratio that is orders of magnitude less than typical experimental fluids, so that the traditional equation-of-state and cavitation models do not apply. Thermal effects in cavitation have not been fully investigated due to experimental difficulties handling cryogenics. This work investigates the physical effects of thermosensitive cavitation in a model representative of a turbopump inducer in a modern rocket engine. This is achieved by utilizing a room-temperature testing fluid that exhibits a thermal effect equivalent to that experienced by cryogenic propellants. Unsteady surface pressures and high speed imaging collected over the span of thermophysical regimes ranging from thermosensitive to isothermal cavitation offer both quantitative and qualitative insight into the physical process of thermal cavitation. Physical and thermodynamic effects are isolated to identify the source of cavity conditions, oscillations and growth/collapse behavior. Planar laser imaging offers an instantaneous look inside the vapor cavity and at the behavior of the boundary between the two-phase region and freestream liquid. Nondimensional parameters are explored, with cavitation numbers, Reynolds Numbers, coefficient of pressure and nondimensional temperature in a broad range. Results in the form of cavitation regime maps, Strouhal Number of cavity

  10. Ignition transient calculations in the Space Shuttle solid rocket motor

    NASA Astrophysics Data System (ADS)

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

    1993-07-01

    The work presented is part of an effort to develop a multidimensional ignition transient model for large solid propellant rocket motors. On the Space Shuttle, the ignition transient in the slot is induced when the igniter, itself a small rocket motor, is fired into the head-end portion of the main rocket motor. The computational results presented in this paper consider two different igniter configurations. The first configuration is a simulated Space Shuttle RSRM igniter which has one central nozzle that is parallel to the centerline of the motor. The second igniter configuration has a nozzle which is canted at an angle of 45 deg from the centerline of the motor. This paper presents a computational fluid dynamic (CFD) analyses of certain flow field characteristics inside the solid propellant star grain slot of the Space Shuttle during the ignition transient period of operation for each igniter configuration. The majority of studies made to date regarding ignition transient performance in solid rocket motors have concluded that the key parameter to be determined is the heat transfer rate to the propellant surface and hence the heat transfer coefficient between the gas and the propellant. In this paper the heat transfer coefficients, pressure and velocity distributions are calculated in the star slot. In order to validate the computational method and to attempt to establish a correlation between the flow field characteristics and the heat transfer rates a series of cold flow experimental investigations were conducted. The results of these experiments show excellent qualitative and quantitative agreement with the pressure and velocity distributions obtained from the CFD analysis. The CFD analysis utilized a classical pipe flow type correlation for the heat transfer rates. The experimental results provide an excellent qualitative comparison with regard to spatial distribution of the heat transfer rates as a function of nozzle configuration and igniter pressure. The

  11. Ignition Transient Calculations in the Space Shuttle Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    The work presented is part of an effort to develop a multidimensional ignition transient model for large solid propellant rocket motors. On the Space Shuttle, the ignition transient in the slot is induced when the igniter, itself a small rocket motor, is fired into the head-end portion of the main rocket motor. The computational results presented in this paper consider two different igniter configurations. The first configuration is a simulated Space Shuttle RSRM igniter which has one central nozzle that is parallel to the centerline of the motor. The second igniter configuration has a nozzle which is canted at an angle of 45 deg from the centerline of the motor. This paper presents a computational fluid dynamic (CFD) analyses of certain flow field characteristics inside the solid propellant star grain slot of the Space Shuttle during the ignition transient period of operation for each igniter configuration. The majority of studies made to date regarding ignition transient performance in solid rocket motors have concluded that the key parameter to be determined is the heat transfer rate to the propellant surface and hence the heat transfer coefficient between the gas and the propellant. In this paper the heat transfer coefficients, pressure and velocity distributions are calculated in the star slot. In order to validate the computational method and to attempt to establish a correlation between the flow field characteristics and the heat transfer rates a series of cold flow experimental investigations were conducted. The results of these experiments show excellent qualitative and quantitative agreement with the pressure and velocity distributions obtained from the CFD analysis. The CFD analysis utilized a classical pipe flow type correlation for the heat transfer rates. The experimental results provide an excellent qualitative comparison with regard to spatial distribution of the heat transfer rates as a function of nozzle configuration and igniter pressure. The

  12. 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.

  13. Cubanes As Solid Propellant Ingredients

    NASA Astrophysics Data System (ADS)

    Cunkle, Glen T.; Willer, Rodney L.

    1988-05-01

    Nine diesters of cubane 1,4-dicarboxylic acid were synthesized and evaluated as advanced propellant plasticizers. Three energetic cubyl ammonium salts of amino and 1,4-diamino cubane were synthesized as models for more highly substituted cubyl ammonium salt oxidizers. Methyl carbamates of amino cubanes have been successfully nitrated with 100% HNO3/(CH3C0)20. These nitrated carbamates are precursors to unsubstituted nitramino cubanes.

  14. 14 CFR 420.65 - Handling of solid propellants.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ....65 Handling of solid propellants. (a) A launch site operator shall determine the maximum total quantity of solid propellants and other solid explosives by class and division, in accordance with 49 CFR... 14 Aeronautics and Space 4 2012-01-01 2012-01-01 false Handling of solid propellants....

  15. 14 CFR 420.65 - Handling of solid propellants.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ....65 Handling of solid propellants. (a) A launch site operator shall determine the maximum total quantity of solid propellants and other solid explosives by class and division, in accordance with 49 CFR... 14 Aeronautics and Space 4 2011-01-01 2011-01-01 false Handling of solid propellants....

  16. Analytical investigation of solid rocket nozzle failure

    NASA Technical Reports Server (NTRS)

    Mccoy, K. E.; Hester, J.

    1985-01-01

    On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

  17. Advanced Solid Rocket Motor case design status

    NASA Technical Reports Server (NTRS)

    Palmer, G. L.; Cash, S. F.; Beck, J. P.

    1993-01-01

    The Advanced Solid Rocket Motor (ASRM) case design aimed at achieving a safer and more reliable solid rocket motor for the Space Shuttle system is considered. The ASRM case has a 150.0 inch diameter, three equal length segment, and 9Ni-4CO-0.3C steel alloy. The major design features include bolted casebolted case joints which close during pressurization, plasma arc welded factory joints, integral stiffener for splash down and recovery, and integral External Tank attachment rings. Each mechanical joint has redundant and verifiable o-ring seals.

  18. Advanced Solid Rocket Motor case design status

    NASA Astrophysics Data System (ADS)

    Palmer, G. L.; Cash, S. F.; Beck, J. P.

    1993-06-01

    The Advanced Solid Rocket Motor (ASRM) case design aimed at achieving a safer and more reliable solid rocket motor for the Space Shuttle system is considered. The ASRM case has a 150.0 inch diameter, three equal length segment, and 9Ni-4CO-0.3C steel alloy. The major design features include bolted casebolted case joints which close during pressurization, plasma arc welded factory joints, integral stiffener for splash down and recovery, and integral External Tank attachment rings. Each mechanical joint has redundant and verifiable o-ring seals.

  19. [Progress in the protective medicine against [correction of aganist] rocket propellents].

    PubMed

    Hu, W X; Tan, C Y; Tan, S J; Jiang, J

    1999-12-01

    To review the progress in the major assignment, the organization and implementation of protection against liquid rocket propellent. The safety detection methods of the rocket [correction of rocked] propellent in the launching field were also discussed. Three steps of the sanitation and protection of the liquid propellent, the toxicity and the toxicology of hydrazine on central nervous system, blood circulatory system, assimilation system, respiratory system, immune system, liver, kidney, eye, skin and its hereditary toxicology were described. In addition, the clinical types of poisoning, the current principle and the common ways of the prevention and treatment of hydrazine and nitrogen oxides poisoning were summarized. PMID:12434814

  20. Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Meacham, Michael B.; Kennett, Andrew; Townsend, Derik J.; Marti, Benjamin

    2013-01-01

    Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

  1. 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.

  2. 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.

  3. 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.

  4. Development of an advanced rocket propellant handler's suit

    NASA Astrophysics Data System (ADS)

    Doerr, DonaldF.

    2001-08-01

    Most launch vehicles and satellites in the US inventory rely upon the use of hypergolic rocket propellants, many of which are toxic to humans. These fuels and oxidizers, such as hydrazine and nitrogen tetroxide have threshold limit values as low as 0.01 PPM. It is essential to provide space workers handling these agents whole body protection as they are universally hazardous not only to the respiratory system, but the skin as well. This paper describes a new method for powering a whole body protective garment to assure the safety of ground servicing crews. A new technology has been developed through the small business innovative research program at the Kennedy Space Center. Currently, liquid air is used in the environmental control unit (ECU) that powers the propellant handlers suit (PHE). However, liquid air exhibits problems with attitude dependence, oxygen enrichment, and difficulty with reliable quantity measurement. The new technology employs the storage of the supply air as a supercritical gas. This method of air storage overcomes all of three problems above while maintaining high density storage at relatively low vessel pressures (<7000 kPa or ˜1000 psi). A one hour prototype ECU was developed and tested to prove the feasibility of this concept. This was upgraded by the design of a larger supercritical dewar capable of holding 7 Kg of air, a supply which provides a 2 hour duration to the PHE. A third version is being developed to test the feasibility of replacing existing air cooling methodology with a liquid cooled garment for relief of heat stress in this warm Florida environment. Testing of the first one hour prototype yielded data comprobable to the liquid air powered predecessor, but enjoyed advantages of attitude independence and oxygen level stability. Thermal data revealed heat stress relief at least as good as liquid air supplied units. The application of supercritical air technology to this whole body protective ensemble marked an advancement in

  5. Development of an advanced rocket propellant handler's suit.

    PubMed

    Doerr, D F

    2001-01-01

    Most launch vehicles and satellites in the US inventory rely upon the use of hypergolic rocket propellants, many of which are toxic to humans. These fuels and oxidizers, such as hydrazine and nitrogen tetroxide have threshold limit values as low as 0.01 PPM. It is essential to provide space workers handling these agents whole body protection as they are universally hazardous not only to the respiratory system, but the skin as well. This paper describes a new method for powering a whole body protective garment to assure the safety of ground servicing crews. A new technology has been developed through the small business innovative research program at the Kennedy Space Center. Currently, liquid air is used in the environmental control unit (ECU) that powers the propellant handlers suit (PHE). However, liquid air exhibits problems with attitude dependence, oxygen enrichment, and difficulty with reliable quantity measurement. The new technology employs the storage of the supply air as a supercritical gas. This method of air storage overcomes all of three problems above while maintaining high density storage at relatively low vessel pressures (<7000 kPa or approximately 1000 psi). A one hour prototype ECU was developed and tested to prove the feasibility of this concept. This was upgraded by the design of a larger supercritical dewar capable of holding 7 Kg of air, a supply which provides a 2 hour duration to the PHE. A third version is being developed to test the feasibility of replacing existing air cooling methodology with a liquid cooled garment for relief of heat stress in this warm Florida environment. Testing of the first one hour prototype yielded data comparable to the liquid air powered predecessor, but enjoyed advantages of attitude independence and oxygen level stability. Thermal data revealed heat stress relief at least as good as liquid air supplied units. The application of supercritical air technology to this whole body protective ensemble marked an

  6. Inflatable mandrel fabrication technology - Advantages for the containment of rocket propellants

    NASA Astrophysics Data System (ADS)

    Moser, Daniel J.

    1992-07-01

    This paper discusses and compares the attributes of various mandrel types as they pertain to the fabrication of filament-wound composite containment vessels for rocket propellants, solid or liquid. The issues of dimensional conformity, processing parameters, unit costs, vessel performance, and development lead times are raised. The continuous fiber reinforced, reusable inflatable mandrel concept is explained and shown to have unique advantages over more traditional mandrel types. Burst pressure performance is equivalent to, or slightly better than, the results from vessels built on the more conventional net metal mandrel. The co-cured insulator process used in conjunction with the inflatable mandrel is shown to be superior in some respects. Some experimental findings are presented along with description of the processing parameters that must be understood when using inflatable mandrels.

  7. 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.

  8. Method for providing real-time control of a gaseous propellant rocket propulsion system

    NASA Technical Reports Server (NTRS)

    Morris, Brian G. (Inventor)

    1991-01-01

    The new and improved methods and apparatus disclosed provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants. Real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

  9. Characterization of booster-rocket propellants and their simulants

    SciTech Connect

    Weirick, L.J.

    1989-01-01

    A series of shock-loading experiments on a composite and an energietic propellant and there simulants was conducted on a light-gas gun. The initial objectives were to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall threshold at various impact velocities. The Hugoniot data measured for the propellants fit the Hugoniot curves provided by the manufacturer of the propellants extremely well and the Hugoniot curves developed for the simulants matched those of the propellants. Threshold pressures to initiate reactions in the composite and energetic propellants were found to be 40 and 3 kbars, respectively. In spall tests, the composite propellant and its simulant exhibited spall strengths around 0.25 and 0.18 kbar, respectively. The energetic propellant and its simulant were somewhat stronger with spall strengths just above 0.33 and 0.22 kbar. 12 refs., 6 figs., 6 tabs.

  10. New solid propellants based on energetic binders and HNF

    NASA Astrophysics Data System (ADS)

    Gadiot, G. M. H. J. L.; Mul, J. M.; Meulenbrugge, J. J.; Korting, P. A. O. G.; Schnorkh, A. J.; Schöyer, H. F. R.

    1993-09-01

    In view of Europe's need for more energetic as well as clean storable solid rocket propellants, formulations based on ammonium nitrate (AN), ammonium dinitramine (ADN) and hydrazinium nitroformate (HNF) as clean oxidizers, glycidyl azide polymer (GAP), polyglycidyl nitrate (PGN), polynitromethyloxetane (PLN) and 3,3-bis(azydomethyl)oxetane (BAMO) as energetic binders and aluminum (A1) as an additional fuel are looking promising in terms of theoretical performance. These propellant formulations do not contain chlorine or chlorine compounds and are, as compared to conventional ammonium perchlorate (AP) based propellants, with respect to environmental aspects virtually non-polluting. For a number of oxidizer fuel combinations theoretical performance figures are given for various oxidizer-to-fuel (O/F) ratios using the aluminum content as a parameter. Performance figures show that the combination HNF/A1/GAP, in particular, is a very promising one. The ingredients HNF and GAP were synthesized in small quantities and their relevant properties were determined. A high performance HNF/A1/GAP formulation was characterized and tested in a laboratorium scale burner yielding pressure vs burning rate data.

  11. 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.

  12. Improved method of measuring pressure coupled response for composite solid propellants

    NASA Astrophysics Data System (ADS)

    Su, Wanxing; Wang, Ningfei; Li, Junwei; Zhao, Yandong; Yan, Mi

    2014-04-01

    Pressure coupled response is one of the main causes of combustion instability in the solid rocket motor. It is also a characteristic parameter for predicting the stability. The pressure coupled response function is usually measured by different methods to evaluate the performance of new propellant. Based on T-burner and "burning surface doubled and secondary attenuation", an improved method for measuring the pressure coupled response of composite propellant is introduced in this article. A computational fluid dynamics (CFD) study has also been conducted to validate the method and to understand the pressure oscillation phenomenon in T-burner. Three rounds of tests were carried out on the same batch of aluminized AP/HTPB composite solid propellant. The experimental results show that the sample propellant had a high response function under the conditions of high pressure (~11.5 MPa) and low frequency (~140 Hz). The numerically predicted oscillation frequency and amplitude are consistent with the experimental results. One practical solid rocket motor using this sample propellant was found to experience pressure oscillation at the end of burning. This confirms that the sample propellant is prone to combustion instability. Finally, acoustic pressure distribution and phase difference in T-burner were analyzed. Both the experimental and numerical results are found to be associated with similar acoustic pressure distribution. And the phase difference analysis showed that the pressure oscillations at the head end of the T-burner are 180° out of phase from those in the aft end of the T-burner.

  13. Facility for cold flow testing of solid rocket motor models

    NASA Astrophysics Data System (ADS)

    Bacchus, D. L.; Hill, O. E.; Whitesides, R. Harold

    1992-02-01

    A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solid propellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The

  14. 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.

  15. Combustion characteristics of a solid propellant with a charring binder

    NASA Technical Reports Server (NTRS)

    Udlock, D. E.; Strand, L. D.

    1973-01-01

    A brief investigation of the combustion characteristics of a solid propellant containing a binder which chars, as opposed to melting or volatizing, has been made. The burning rate of the propellant with the charring binder was significantly higher than similar propellants containing non-charring binders. High speed motion pictures of the burning propellant showed that the aluminum burned on the regressing surface, rather than a short distance from it as is typical with composite propellants.

  16. Designing for solid rocket booster reusability

    NASA Technical Reports Server (NTRS)

    Nevins, C. D.

    1975-01-01

    The Solid Rocket Booster (SRB) of the Space Shuttle Program has been designed to be recovered, refurbished, and reused up to 19 times on subsequent launches. The design modifications to the SRB to incorporate this capability include the addition of a parachute recovery system and minor structural modifications to the nose cone, forward skirt, aft skirt, and solid rocket motor structures for withstanding the water impact loading. In addition, the detail design of subsystem hardware requires careful attention to corrosion prevention and facility of refurbishment operations. A cost analysis comparing the reusable SRB with an expendable version shows a substantial reduction in Space Shuttle cost per flight is achieved with the reusable design approach.

  17. Design issues for lunar in situ aluminum/oxygen propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Meyer, Michael L.

    1992-01-01

    Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

  18. IUS solid rocket motor contamination prediction methods

    NASA Technical Reports Server (NTRS)

    Mullen, C. R.; Kearnes, J. H.

    1980-01-01

    A series of computer codes were developed to predict solid rocket motor produced contamination to spacecraft sensitive surfaces. Subscale and flight test data have confirmed some of the analytical results. Application of the analysis tools to a typical spacecraft has provided early identification of potential spacecraft contamination problems and provided insight into their solution; e.g., flight plan modifications, plume or outgassing shields and/or contamination covers.

  19. Solid rocket booster water impact test

    NASA Technical Reports Server (NTRS)

    Bugg, F.

    1982-01-01

    Water impact drop tests were performed on the space shuttle solid rocket boosters (SRB). Peak water impact pressures and pressure/time traces were measured for various impact velocities using a two-dimensional, full-scale SRB aft skirt internal ring model. Passive burst disc-type pressure transducers were calibrated for use on flight SRB's. The effects on impact pressure of small ring configuration changes and application of thermal protection system cork layers were found to be negligible.

  20. Space Shuttle Reusable Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Moore, Dennis; Phelps, Jack; Perkins, Fred

    2010-01-01

    RSRM is a highly reliable human-rated Solid Rocket Motor: a) Largest diameter SRM to achieve flight status; b) Only human-rated SRM. RSRM reliability achieved by: a)Applying special attention to Process Control, Testing, and Postflight; b) Communicating often; c) Identifying and addressing issues in a disciplined approach; d) Identifying and fully dispositioning "out-of-family" conditions; e) Addressing minority opinions; and f) Learning our lessons.

  1. Design and Experimental Study on Spinning Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Xue, Heng; Jiang, Chunlan; Wang, Zaicheng

    The study on spinning solid rocket motor (SRM) which used as power plant of twice throwing structure of aerial submunition was introduced. This kind of SRM which with the structure of tangential multi-nozzle consists of a combustion chamber, propellant charge, 4 tangential nozzles, ignition device, etc. Grain design, structure design and prediction of interior ballistic performance were described, and problem which need mainly considered in design were analyzed comprehensively. Finally, in order to research working performance of the SRM, measure pressure-time curve and its speed, static test and dynamic test were conducted respectively. And then calculated values and experimental data were compared and analyzed. The results indicate that the designed motor operates normally, and the stable performance of interior ballistic meet demands. And experimental results have the guidance meaning for the pre-research design of SRM.

  2. Evidence of erosive burning in shuttle solid rocket motor

    NASA Technical Reports Server (NTRS)

    Martin, C. L.

    1983-01-01

    Known models of Shuttle Solid Rocket Motor (SRM) performance have failed to produce pressure-time traces which accurately matched actual motor performance, especially during the first 5 seconds after ignition and during the last quarter of web burn time. Efforts to compensate for these differences in model reconstruction and actual performance resulted in resorting to the use of a Burning Anomaly Rate Function (BARF). It was suspected that propellant erosive burning was primarily responsible for the variation of model from actual results. The three dimensional Hercules Grain Design and Internal Ballistics Evaluation Program was made operational and slightly modified and an extensive trial and error effort was begun to test the hypothesis of erosive burning as an explanation of the burning anomaly. It was found that introduction of erosive burning (using Green's erosive burning equation) over portions of the aft segment grain and above a threshold gas Mach number did, in fact, give excellent agreement with the actual motor trace.

  3. 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.

  4. Five-Segment Solid Rocket Motor Development Status

    NASA Technical Reports Server (NTRS)

    Priskos, Alex S.

    2012-01-01

    In support of the National Aeronautics and Space Administration (NASA), Marshall Space Flight Center (MSFC) is developing a new, more powerful solid rocket motor for space launch applications. To minimize technical risks and development costs, NASA chose to use the Space Shuttle s solid rocket boosters as a starting point in the design and development. The new, five segment motor provides a greater total impulse with improved, more environmentally friendly materials. To meet the mass and trajectory requirements, the motor incorporates substantial design and system upgrades, including new propellant grain geometry with an additional segment, new internal insulation system, and a state-of-the art avionics system. Significant progress has been made in the design, development and testing of the propulsion, and avionics systems. To date, three development motors (one each in 2009, 2010, and 2011) have been successfully static tested by NASA and ATK s Launch Systems Group in Promontory, UT. These development motor tests have validated much of the engineering with substantial data collected, analyzed, and utilized to improve the design. This paper provides an overview of the development progress on the first stage propulsion system.

  5. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt ...

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

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt sitting on ground support equipment in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center while being prepared for mating with the Frustum-Nose Cap Assembly and the Forward Rocket Motor Segment. The prominent feature in this view is the electrical, data, telemetry and safety systems terminal which connects to the Aft Skirt Assembly systems via the Systems Tunnel that runs the length of the Rocket Motor. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  6. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Thomas, Matt; Bossard, John; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

  7. 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.

  8. 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

  9. Space Shuttle solid rocket booster dewatering system

    NASA Technical Reports Server (NTRS)

    Fishel, K. R.

    1982-01-01

    After the launch of the Space Shuttle, the two solid rocket boosters (SRB's) are jettisoned into the ocean where they float in a spar (vertical) mode. It is cost effective to recover the SRB's. A remote controlled submersible vehicle has been developed to aid in their recovery. The vehicle is launched from a support ship, maneuvered to the SRB, then taken to depth and guided into the rocket nozzle. It then dewaters the SRB, using compressed air from the ship, and seals the nozzle. When dewatered, the SRB floats in a log (horizontal) mode and can be towed to port for reuse. The design of the remote controlled vehicle and its propulsion system is presented.

  10. 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.

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

    NASA Technical Reports Server (NTRS)

    1991-01-01

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

  12. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt, ...

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

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt, Frustum and Nose Cap mated assembly undergoing final preparations in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center. In this view the access panel on the Forward Skirt is removed and you can see a small portion of the interior of the Forward Skirt. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  13. Closeup view of the Solid Rocket Booster Frustum and Nose ...

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

    Close-up view of the Solid Rocket Booster Frustum and Nose Cap assembly undergoing preparations and close-out procedures in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center. The Nose Cap contains the Pilot and Drogue Chutes and the Frustum contains the three Main Parachutes, Altitude Switches and forward booster Separation Motors. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  14. 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.

  15. 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.

  16. Cape Canaveral Air Force Station, Launch Complex 39, Solid Rocket ...

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

    Cape Canaveral Air Force Station, Launch Complex 39, Solid Rocket Booster Disassembly & Refurbishment Complex, Thrust Vector Control Deservicing Facility, Hangar Road, Cape Canaveral, Brevard County, FL

  17. Propellant acquisition for single-stage rocket technology

    NASA Astrophysics Data System (ADS)

    Grayson, G.; Distefano, E.

    1993-06-01

    An analytical design approach defines the liquid oxygen (LO2) propellant-acquisition system for the vertical-take-off and vertical-landing DC-X flight test vehicle. The DC-X trajectory includes a rotation maneuver in which the vehicle pitches from a nose-first to a tail-first orientation. The 8-ft diameter LO2 tank accommodates outflow rates as high as 120 lbm/sec, while vehicle drag forces result in a lateral acceleration approaching 25 ft/sq sec during rotation. FLOW-3D computational results show that, without propellant control, these conditions cause the LO2 to slosh within the tank and expose the outlet to ullage gas during the rotation maneuver. Using FLOW 3-D simulations of several baffle configurations, one propellant acquisition concept is selected. In this concept, a simple cone-shaped containment baffle compartmentalizes the tank and an outlet extension offsets the outlet to where the propellant is biased. The final design incorporates a factor of safety greater than 2 resulting in continuous coverage of the outlet by propellant as validated by FLOW-3D simulation.

  18. Radiation from advanced solid rocket motor plumes

    NASA Technical Reports Server (NTRS)

    Farmer, Richard C.; Smith, Sheldon D.; Myruski, Brian L.

    1994-01-01

    The overall objective of this study was to develop an understanding of solid rocket motor (SRM) plumes in sufficient detail to accurately explain the majority of plume radiation test data. Improved flowfield and radiation analysis codes were developed to accurately and efficiently account for all the factors which effect radiation heating from rocket plumes. These codes were verified by comparing predicted plume behavior with measured NASA/MSFC ASRM test data. Upon conducting a thorough review of the current state-of-the-art of SRM plume flowfield and radiation prediction methodology and the pertinent data base, the following analyses were developed for future design use. The NOZZRAD code was developed for preliminary base heating design and Al2O3 particle optical property data evaluation using a generalized two-flux solution to the radiative transfer equation. The IDARAD code was developed for rapid evaluation of plume radiation effects using the spherical harmonics method of differential approximation to the radiative transfer equation. The FDNS CFD code with fully coupled Euler-Lagrange particle tracking was validated by comparison to predictions made with the industry standard RAMP code for SRM nozzle flowfield analysis. The FDNS code provides the ability to analyze not only rocket nozzle flow, but also axisymmetric and three-dimensional plume flowfields with state-of-the-art CFD methodology. Procedures for conducting meaningful thermo-vision camera studies were developed.

  19. Detailed modal testing of a solid rocket motor using a portable test system

    NASA Technical Reports Server (NTRS)

    Glozman, Vladimir; Brillhart, Ralph D.

    1990-01-01

    Modern analytical techniques have expended the ability to evaluate solid rocket motors used in launch vehicles. As more detailed models of solid rocket motors were developed, testing methods were required to verify the models. Experimental modal analysis (modal testing) of space structures and launch vehicles has been a requirement for model validation for many years. However, previous testing of solid rocket motors has not typically involved dynamic modal testing of full scale motors for verification of solid propellant or system assembly properties. Innovative approaches to the testing of solid rocket motors were developed and modal testing of a full scale, two segment Titan 34D Solid Rocket Motor (SRM) was performed to validate detailed computer modeling. Special modifications were made to convert an existing facility into a temporary modal test facility which would accommodate the test article. The assembly of conventional data acquisition equipment into a multiple channel count portable system has made modal testing in the field feasible. Special purpose hydraulic exciters were configured to apply the dynamic driving forces required. All instrumentation and data collection equipment were installed at the test site for the duration of the test program and removed upon completion. Conversion of an existing test facility into a temporary modal test facility, and use of a multiple channel count portable test data acquisition system allowed all test objectives to be met and resulted in validation of the computer model in a minimum time.

  20. 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.

  1. Environmental impact statement Space Shuttle advanced solid rocket motor program

    NASA Technical Reports Server (NTRS)

    1989-01-01

    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. Sites being considered for the new facilities include John C. Stennis Space Center, Hancock County, Mississippi; the Yellow Creek site in Tishomingo County, Mississippi, which is currently in the custody and control of the Tennessee Valley Authority; and John F. Kennedy Space Center, Brevard County, Florida. TVA proposes to transfer its site to the custody and control of NASA if it is the selected site. All facilities need not be located at the same site. Existing facilities which may provide support for the program include Michoud Assembly Facility, New Orleans Parish, Louisiana; and Slidell Computer Center, St. Tammany Parish, Louisiana. NASA's preferred production location is the Yellow Creek site, and the preferred test location is the Stennis Space Center.

  2. 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.

  3. Ariane 5 solid rocket booster dynamic behavior with respect to pressure oscillations

    NASA Astrophysics Data System (ADS)

    Durin, G.; Bouvier, F.; Mastrangelo, G.; Robert, E.

    2011-10-01

    Numerical simulations are performed to simulate the dynamic behavior of the Ariane 5 Solid Rocket Booster (SRB) excited by internal pressure oscillations. These pressure oscillations have sliding frequencies close to acoustic natural frequencies and changing amplitudes. Because the modal behavior of the booster is also continuously changing during the flight due to propellant burning, predictions with finite-element (FE) modeling (modal analyses, harmonic and transient responses) are necessary to predict and understand the dynamic behavior of the booster.

  4. 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.

  5. The scaling of the threshold conditions for solid propellant erosive burning

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Nguyen, M. H.; Cohen, N. S.

    1988-01-01

    Rocket test firings were performed to measure the transition length threshold conditions while systematically varying various rocket motor parameters. These include the crossflow velocity, the chamber pressure, the propellant nonerosive burning rate, the propellant surface roughness, and the motor port diameter. The erosive burning trends with varying propellant burning rate, motor chamber pressure, and mass flow rate are consistent with published results.

  6. Effects of propellant composition variables on acceleration-induced burning-rate augmentation of solid propellants

    NASA Technical Reports Server (NTRS)

    Northam, G. B.

    1972-01-01

    This work was conducted to define further the effects of propellant composition variables on the acceleration-induced burning rate augmentation of solid propellants. The rate augmentation at a given acceleration was found to be a nonlinear inverse function of the reference burning rate and not controlled by binder or catalyst type at a given reference rate. A nonaluminized propellant and a low rate double-base propellant exhibited strong transient rate augmentation due to surface pitting resulting from the retention of hot particles on the propellant surface.

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

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

  8. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt ...

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

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt sitting on ground support equipment in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center while being prepared for mating with the Frustum-Nose Cap Assembly and the Forward Rocket Motor Segment. The prominent feature in this view is the Forward Thrust Attach Fitting which mates up with the Forward Thrust Attach Fitting of the External Tank (ET) at the ends of the SRB Beam that runs through the ET's Inter Tank Assembly. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  9. Embeddable structural sensors for SHM of solid rocket grains

    NASA Astrophysics Data System (ADS)

    Chelner, Herb; Buswell, Jim

    2006-05-01

    Structural Health Monitoring is essential in that any event that may compromise the solid rocket motor must be detected. The magnitude, position and time of any imposed event that may damage the propellant grain, bonding system or integrity of the case must be detected and identified for safe operation of the motor. The current embedded sensor technology has been developed to monitor the effect of these events on the propellant grain. Normal bond stress and temperature can be measured using DBST sensors and the output interpreted to confirm integrity. It has been shown that the presences of de-bond and cracks can be determined. Current work is in progress to establish if these embedded sensors can be used to determine position and size of such defects. The stress distribution in a typical propellant grain also has a shear component particularly near the ends and around any flaps, slots or stress relieving devices. This can be the critical stress under certain loads and in complex geometries. Therefore, a recent addition to the range is a sensor to measure shear stress in the same body as the DBST. Motors can be stored for long periods before being used so the sensor system must also be reliable and stable for at least twenty years of operation. Similar sensors stored for ten years have shown little change and tests are being undertaken to establish the confidence that the sensor system will last the life of the motor. This paper will review the recent development and testing of these embeddable sensors, and results to date will be discussed.

  10. Rocket propellant inhalation in the Apollo-Soyuz astronauts.

    PubMed

    DeJournette, R L

    1977-10-01

    Acute exposure to monomethylhydrazine and dinitrogen tetroxide, the principal toxic irritants in rocket fuels, is described with particular attention to the development of pulmonary edema as a herbinger of more severe central nervous system toxicity. An acute respiratory embarrassment is documented and possible means of therapy based on animal experimental models is suggested. Early clinical and radiographic examination as a baseline for further evaluation is essential, with follow-up radiographs recommended for assessment of possible developing chronic lung disease. PMID:897171

  11. Combustion response modeling for composite solid propellants

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A computerized mathematical model of the combustion response function of composite solid propellants was developed with particular attention to the contributions of the solid phase heterogeneity. The one-dimensional model treats the solid phase as alternating layers of ammonium perchlorate and binder, with an exothermic melt layer at the surface. Solution of the Fourier heat equation in the solid provides temperature and heat flux distributions with space and time. The problem is solved by conserving the heat flux at the surface from that produced by a suitable model of the gas phase. An approximation of the BDP flame model is utilized to represent the gas phase. By the use of several reasonable assumptions, it is found that a significant portion of the problem can be solved in closed form. A method is presented by which the model can be applied to tetramodal particle size distributions. A computerized steady-state version of the model was completed, which served to validate the various approximations and lay a foundation for the combustion response modeling. The combustion response modeling was completed in a form which does not require an iterative solution, and some preliminary results were acquired.

  12. 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.

  13. 14 CFR 420.65 - Handling of solid propellants.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... quantity of solid propellants and other solid explosives by class and division, in accordance with 49 CFR... only of a public highway or railroad line. (4) A launch site operator may use linear interpolation...

  14. General view of a fully assembled Solid Rocket Booster sitting ...

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

    General view of a fully assembled Solid Rocket Booster sitting atop the Mobile Launch Platform in the Vehicle Assembly Building at Kennedy Space Center - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  15. System for imposing directional stability on a rocket-propelled vehicle

    NASA Technical Reports Server (NTRS)

    Perkins, H. (Inventor)

    1976-01-01

    An improved system for use in imposing directional stability on a rocket-propelled vehicle is described. The system includes a pivotally supported engine-mounting platform, a gimbal ring mounted on the platform and adapted to pivotally support a rocket engine and an hydraulic actuator connected to the platform for imparting selected pivotal motion. An accelerometer and a signal comparator circuit for providing error intelligence indicative of aberration in vehicle acceleration is included along with an actuator control circuit connected with the actuator and responsive to error intelligence for imparting pivotal motion to the platform. Relocation of the engine's thrust vector is thus achieved for imparting directional stability to the vehicle.

  16. Transferring jet engine diagnostic and control technology to liquid propellant rocket engines

    SciTech Connect

    Alcock, J.F.; Hagar, S.K.

    1989-01-01

    This paper presents the methodology for developing a diagnostic and control system for a current, operational jet engine. A description is given of each development stage, the system components and the technologies which could be transferred to liquid propellant rocket engines. Finally, the operational impact is described in terms of cost and maintenance based on actual jet engine experience. Efforts are continuing to develop new diagnostic techniques under IR D for application on the advanced technical fighter. Already improved techniques and application methods are becoming available. This technology is being evaluated and may also be transferred to rocket engine diagnostic and control system development.

  17. 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.

  18. Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

    NASA Technical Reports Server (NTRS)

    Hulka, James R.; Jones, Gregg W.

    2010-01-01

    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using

  19. The method of solid rocket motors firings environmental engineering model

    NASA Astrophysics Data System (ADS)

    Pang, Baojun; Xu, Ke; Peng, Keke; Mi, Yaoqi

    The solid rocket motors firings is one main source of space debris, the solid rocket motors firings model is a part of space debris engineering model. In this paper, researching the NASA and ESA model to achieve an appropriate firing model, using the discrete method to model the solid rocket motors firings; application of the long-term approximation orbit evolution algorithm to calculate the evolution of firings generated by a single solid rocket motors ignition event in space; finally, application of space debris environment space debris density algorithm to calculate the distribution of firings generated by a single solid rocket motors ignition event in space, analysing the influence on the space environment and spacecraft.

  20. The Chameleon Solid Rocket Propulsion Model

    SciTech Connect

    Robertson, Glen A.

    2010-01-28

    The Khoury and Weltman (2004a and 2004b) Chameleon Model presents an addition to the gravitation force and was shown by the author (Robertson, 2009a and 2009b) to present a new means by which one can view other forces in the Universe. The Chameleon Model is basically a density-dependent model and while the idea is not new, this model is novel in that densities in the Universe to include the vacuum of space are viewed as scalar fields. Such an analogy gives the Chameleon scalar field, dark energy/dark matter like characteristics; fitting well within cosmological expansion theories. In respect to this forum, in this paper, it is shown how the Chameleon Model can be used to derive the thrust of a solid rocket motor. This presents a first step toward the development of new propulsion models using density variations verse mass ejection as the mechanism for thrust. Further, through the Chameleon Model connection, these new propulsion models can be tied to dark energy/dark matter toward new space propulsion systems utilizing the vacuum scalar field in a way understandable by engineers, the key toward the development of such systems. This paper provides corrections to the Chameleon rocket model in Robertson (2009b).

  1. Computational analysis of liquid hypergolic propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Krishnan, A.; Przekwas, A. J.; Gross, K. W.

    1992-01-01

    The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

  2. A study on various methods of supplying propellant to an orbit insertion rocket engine

    NASA Technical Reports Server (NTRS)

    Boretz, J. E.; Huniu, S.; Thompson, M.; Pagani, M.; Paulsen, B.; Lewis, J.; Paul, D.

    1980-01-01

    Various types of pumps and pump drives were evaluated to determine the lightest weight system for supplying propellants to a planetary orbit insertion rocket engine. From these analyses four candidate propellant feed systems were identified. Systems Nos. 1 and 2 were both battery powered (lithium-thionyl-chloride or silver-zinc) motor driven pumps. System 3 was a monopropellant gas generator powered turbopump. System 4 was a bipropellant gas generator powered turbopump. Parameters considered were pump break horsepower, weight, reliability, transient response and system stability. Figures of merit were established and the ranking of the candidate systems was determined. Conceptual designs were prepared for typical motor driven pumps and turbopump configurations for a 1000 lbf thrust rocket engine.

  3. Prediction of explosive yield and other characteristics of liquid rocket propellant explosions

    NASA Technical Reports Server (NTRS)

    Farber, E. A.; Smith, J. H.; Watts, E. H.

    1973-01-01

    Work which has been done at the University of Florida in arriving at credible explosive yield values for liquid rocket propellants is presented. The results are based upon logical methods which have been well worked out theoretically and verified through experimental procedures. Three independent methods to predict explosive yield values for liquid rocket propellants are described. All three give the same end result even though they utilize different parameters and procedures. They are: (1) mathematical model; (2) seven chart approach; and (3) critical mass method. A brief description of the methods, how they were derived, how they were applied, and the results which they produced are given. The experimental work used to support and verify the above methods both in the laboratory and in the field with actually explosive mixtures are presented. The methods developed are used and their value demonstrated in analyzing real problems, among them the destruct system of the Saturn 5, and the early configurations of the space shuttle.

  4. Biogenic technology for recultivation of lands contaminated due to rocket propellant spillage

    NASA Astrophysics Data System (ADS)

    Kovshov, S. V.; Garkushev, A. U.; Sazykin, A. M.

    2015-04-01

    This article describes the problem of soil properties deterioration due to rocket propellant spillage. Melange and samin are considered to be the main pollutants. Provision is made for assessment of the existing mechanisms for monitoring of quality and recultivation of lands disturbed by rocket propellant spills. Some major disadvantages of currently used standard recultivation technologies are listed. An alternative is the use of more environmentally safe and cost effective methods aimed at disturbed lands biological restoration. An example of such a technology is covering the affected area with a biogenic mixture consisting of biohumus and sodium carboxymethyl cellulose followed by seeding it with specially selected herbal mixtures. It was found out that the most rational parameters of such protective layer is its thickness of 3 cm, and 99:1 ratio of its constituent components.

  5. Regeneratively cooled rocket engine for space storable propellants

    NASA Technical Reports Server (NTRS)

    Wagner, W. R.

    1973-01-01

    Analysis, design, fabrication, and test efforts were performed for the existing OF2/B2H6 regeneratively cooled lK (4448 N) thrust chamber to illustrate simultaneous B2H6 fuel and OF2 oxidizer cooling and to provide results for a gaseous propellant condition injected into the combustion chamber. Data derived from performance, thermal and flow measurements confirmed predictions derived from previous test work and from concurrent analytical study. Development data derived from the experimental study were indicated to be sufficient to develop a preflight thrust chamber demonstrator prototype for future space mission objectives.

  6. Advanced Solid Rocket Motor nozzle development status

    NASA Astrophysics Data System (ADS)

    Kearney, W. J.; Moss, J. D.

    1993-06-01

    This paper presents a status update of the design and development of an improved nozzle for the Advanced Solid Rocket Motor (ASRM). The ASRM nozzle incorporates advanced state-of-the-art design features and materials which contribute to enhanced safety, reliability, performance, and producibility for the space shuttle boosters. During 1992 the nozzle design progressed through a successful Preliminary Design Review (PDR). An improved ablative material development program also culminated in the selection of new standard and low density carbon cloth phenolic prepreg offering reduced variability and improved process attributes. A subscale motor test series to evaluate new materials and design features was also completed. An overview update of the matured design characteristics, supporting analysis, key development-program results and program status and plans is reported.

  7. Ocean recovery of Shuttle Solid Rocket Boosters.

    NASA Technical Reports Server (NTRS)

    Junker, L. J.

    1973-01-01

    Cost effective recovery of the expended Space-Shuttle Solid Rocket Boosters (SRB) from the ocean will result in significant overall cost savings to the Space Shuttle Program. The ocean recovery mission begins with the dispatching of the recovery team to the predicted splashdown area. The SRBs, drogue parachutes and main parachutes must be tracked, located, retrieved, and transported to land where they will be refurbished and recycled for reuse. Trade studies to be conducted will consider the recovery mission requirements and weigh the advantages, disadvantages and costs of various candidate recovery systems. Major parameters effecting the selection of the final system will ensure that the system will meet overall objectives. Large- and small-scale SRB model testing has been conducted to establish characteristics of SRBs during water entry, floating free and under tow.

  8. Solid rocket motor nozzle flexseal design sensitivity

    NASA Astrophysics Data System (ADS)

    Donat, James R.

    1993-02-01

    On solid rocket motors, direction is controlled by controlling the thrust vector. To achieve this, the nozzle usually incorporates a flexseal that allows the nozzle to vector (or rotate) in any direction. The flexseal has a core of alternating layers of elastomer pads and metal or composite shims. Flexseal core design is an iterative process. An estimate of the flexseal core geometry is made. The core is then analyzed for performance characteristics such as stress, weight, and the torque required to vector the core. Based on a comparison between the requirements/constraints and analysis results, another estimate of the geometry is then made. Understanding the effects changes in the core geometry have on the performance characteristics greatly decreases the number of iterations and time required to optimize the design. This paper documents a study undertaken to better understand these effects and how sensitive performance characteristics are to core geometry changes.

  9. Attenuation studies of booster-rocket propellants and their simulants

    SciTech Connect

    Weirick, L.J.

    1990-08-01

    A series of impact experiments on a composite propellant, an energetic propellant, and their simulants was recently completed using a light-gas gun. Previous experiments were done to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall damage at various impact velocities. The present studies measured the attenuation of shock waves in these materials, completing the shock characterization needed for material modeling. An initial impulse of 2.0 GPa magnitude and {approximately}0.6 {mu}s duration was imposed upon samples of various thicknesses. VISAR was used to measure the free-surface velocity at the back of the samples; these data were used to generate a curve of shock-wave attenuation versus sample thickness for each material. Results showed that all four materials attenuated the shock wave very similarly. Material thicknesses of 3.0, 7.62, 12.7, and 19.0 mm attenuated the shock wave {approximately}16%, 33%, 50%, and 66% respectively. 14 refs., 12 figs., 4 tabs.

  10. Regeneratively cooled rocket engine for space storable propellants

    NASA Technical Reports Server (NTRS)

    Wagner, W. R.; Waldman, B. J.

    1973-01-01

    Analyses and experimental studies were performed with the OF2 (F2/O2)/B2H6 propellant combination over a range in operating conditions to determine suitability for a space storable pressure fed engine configuration for an extended flight space vehicle configuration. The regenerative cooling mode selected for the thrust chamber was explored in detail with the use of both the fuel and oxidizer as coolants in an advanced milled channel construction thrust chamber design operating at 100 psia chamber pressure and a nominal mixture ratio of 3.0 with a 60:1 area ratio nozzle. Benefits of the simultaneous cooling as related to gaseous injection of both fuel and oxidizer propellants were defined. Heat transfer rates, performance and combustor stability were developed for impinging element triplet injectors in uncooled copper calorimeter hardware with flow, pressure and temperature instrumentation. Evaluation of the capabilities of the B2H6 and OF2 during analytical studies and numerous tests with flow through electrically heated blocks provided design criteria for subsequent regenerative chamber design and fabrication.