Sample records for solid propellent rockets

  1. Solid propellant rocket motor

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


    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

    Jacox, James L.; Bradford, Daniel J.


    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. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  4. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  5. Specific Impulses Losses in Solid Propellant Rockets


    to use the collision function form proposed by Golovin to simplify this production term: 4C><=) <P- .: Accordingly: m hence, by integration: Now, we...November 21, 1940 in Paris, Seine. VFirst Thesis. "Contribution to the Study of Specific i Impulse Loss in Solid Propellant Rockets." Second Thesis

  6. Solid propellants for rockets. Rocket suishin yaku

    Kubota, N. (Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Physical and chemical ProPerties and combustion characteristics of propellants differ according to the combination of oxidizers and fuel components. Composite smoke propellant, having crystalline ammonium perchlorate as an oxidizer and hydrocarbon Polymer as a fuel, has higher specific impulse and improved mechanical properties compared to smokeless double base propellant consisting of nitroglycerin and nirocellulose. Double base propellants with low specific impulse are combined with nitramines( RDX or HMX ) to make composite modified double based( CMDB ) propellants, as a result the smokeless property of double base propellant is preserved and the combustion efficiency is increased. With the combination of oxidizing agents and fuels, formation of various high functional propellants has been possible and energetic azide polymers have provided possibilities for fuels of propellants. 3 refs., 6 figs., 3 tabs.

  7. Particle behavior in solid propellant rockets

    Netzer, D. W.; Diloreto, V. D.; Dubrov, E.


    The use of holography, high speed motion pictures, light scattering measurements, and post-fire particle collection/scanning electron microscopic examination to study the combustion of composite solid propellants is discussed. The relative advantages and disadvantages of the different experimental techniques for obtaining two-phase flow characteristics within the combustion environment of a solid propellant grain are evaluated. Combustion bomb studies using high speed motion pictures and post-fire residue analysis were completed for six low metal content propellants. Resolution capabilities and the relationships between post-fire residue and motion picture data are determined. Initial testing using a holocamera together with a 2D windowed motor is also described.

  8. Effect of Temperature on Mechanical Properties of Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical properties of solid rocket propellants are dependent on temperature. Any change in temperature brings significant change in the tensile strength, percentage elongation, and elastic modulus of the propellant. Different classes of operational solid rocket propellants namely extruded double-base propellants, composite, extruded composite and nitrarte ester polyester propellants were evaluated at different temperatures in the operating range of the rockets and missiles preferably in the range of –50 oC to +55 oC. It was observed that for each class of propellant, as temperature reduces, propellant becomes hard. This is depicted by increase in elastic modulus and tensile strength of the material. However, trend of percentage elongation is not very uniform. Extruded double-base propellants show less percentage elongation (around 1 per cent at reduced temperature (–50 oC probably due to brittleness. So is the trend with case-bonded composite propellants. However, reverse trend is exhibited by cartridge-loaded composite propellants and nitrate ester polyester propellants. Such propellants show higher percentage elongation (6 per cent for CLCP and 35 per cent for NEPE at reduced temperature (–50 oC. This makes such propellants tough and more area under stress-strain curve at reduced temperature is observed.Defence Science Journal, 2011, 61(6, pp.529-533, DOI:

  9. Burn Rate Modelling of Solid Rocket Propellants (Short Communication

    A.R. Kulkarni


    Full Text Available A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

  10. Sensitivity of solid rocket propellants for card gap test

    Kimura, Eishu; Oyumi, Yoshio (Japan Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solid propellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP- and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here. (orig.)

  11. Propellant development for the Advanced Solid Rocket Motor

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


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

  12. Solid propellants.

    Marsh, H. E., Jr.; Hutchison, J. J.


    The basic principles underlying propulsion by rocket motor are examined together with the configuration of a solid propellant motor. Solid propellants and their preparation are discussed, giving attention to homogeneous propellants, composite propellants, energetic considerations in choosing a solid propellant, the processing of composite propellants, and some examples of new developments. The performance of solid propellants is investigated, taking into account characteristics velocity, the specific impulse, and performance calculations. Aspects of propellant development considered include nonperformance requirements for solid propellants, the approach to development, propellant mechanical properties, and future trends.

  13. Rocket Solid Propellant Alternative Based on Ammonium Dinitramide

    Grigore CICAN


    Full Text Available Due to the continuous run for a green environment the current article proposes a new type of solid propellant based on the fairly new synthesized oxidizer, ammonium dinitramide (ADN. Apart of having a higher specific impulse than the worldwide renowned oxidizer, ammonium perchlorate, ADN has the advantage, of leaving behind only nitrogen, oxygen and water after decomposing at high temperatures and therefore totally avoiding the formation of hydrogen chloride fumes. Based on the oxidizer to fuel ratios of the current formulations of the major rocket solid booster (e.g. Space Shuttle’s SRB, Ariane 5’s SRB which comprises mass variations of ammonium perchlorate oxidizer (70-75%, atomized aluminum powder (10-18% and polybutadiene binder (12-20% a new solid propellant was formulated. As previously stated, the new propellant formula and its variations use ADN as oxidizer and erythritol tetranitrate as fuel, keeping the same polybutadiene as binder.

  14. MEMS-Based Solid Propellant Rocket Array Thruster

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

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

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

    Moore, Carleton J.


    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.

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

    Pobedonostsev, Y. A.


    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.

  17. Viscoelastic Modelling of Solid Rocket Propellants using Maxwell Fluid Model

    Himanshu Shekhar


    Full Text Available Maxwell fluid model consisting of a spring and a dashpot in series is applied for viscoelastic characterisation of solid rocket propellants. Suitable values of spring constant and damping coefficient wereemployed by least square variation of errors for generation of complete stress-strain curve in uniaxial tensile mode for case-bonded solid propellant formulations. Propellants from the same lot were tested at different strain rates. It was observed that change in spring constant, representing elastic part was very small with strain rate but damping constant varies significantly with variation in strain rate. For a typical propellant formulation, when strain rate was raised from 0.00037/s to 0.185/s, spring constant K changed from 5.5 MPato 7.9 MPa, but damping coefficient D was reduced from 1400 MPa-s to 4 MPa-s. For all strain rates, stress-strain curve was generated using Maxwell model and close matching with actual test curve was observed.This indicates validity of Maxwell fluid model for uniaxial tensile testing curves of case-bonded solid propellant formulations. It was established that at higher strain rate, damping coefficient becomes negligible as compared to spring constant. It was also observed that variation of spring constant is logarithmic with strain rate and that of damping coefficient follows power law. The correlation coefficients were introduced to ascertain spring constants and damping coefficients at any strain rate from that at a reference strain rate. Correlationfor spring constant needs a coefficient H, which is function of propellant formulation alone and not of test conditions and the equation developeds K2 = K1 + H ´ ln{(de2/dt/(de1/dt}. Similarly for damping coefficient D also another constant S is introduced and prediction formula is given by D2 = D1 ´ {(de2/dt/(de1/dt}S.Evaluating constants H and S at different strain rates validate this mathematical formulation for differentpropellant formulations

  18. Thermodynamic cycle analysis of solid propellant air-turbo-rocket

    CHEN Xiang; CHEN Yu-chun; TU Qiu-ye; ZHANG Hong; CAI Yuan-hu


    Solid propellant air-turbo-rocket (SPATR) is an air-breathing propulsion system. A numerical model of performance and characteristics analysis for SPATR was presented and the corresponding computer program was written according to the operation characteristics of SPATR. The influence on the SPATR performance at design point caused by the gas generator exit parameters and compressor pressure ratio had been computed and analyzed in detail. The off-design perform-ance of SPATR at sea level and high altitude had also been computed. The performance of thrust and specific impulse for SPATR with different solid propellant had been compared at off-design points, and the off-design performance comparison had been made between fuel-rich and oxygen-rich. The computation results indicated that SPATR operates within wide range of Maeh number (0 ~3) and altitude (0~12 km), and SPATR possesses high specific thrust (1 200 N/(kg/s)) and high specific impulse (7000 N/ (kg/s)) when fuel-air ratio of combustor equals fuel-air ratio.

  19. Solid propellant processing factor in rocket motor design


    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.

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

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


    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.




    Full Text Available In this study, three dimensional modelling of extrusion forming of a double base solid rocket propellant is performed on Ansys® finite element simulation package. For the purpose of initial model construction and later comparisons with elastoviscoplastik model, the solid propellant is assumed to obey the elastic-plastic material response during the direct extrusion process. Taking into account the contact surface behavior with Coulomb friction and geometric and material nonlinearities, an incremental large large strain solution methodology has been adapted in the simulation. The hydrostatic pressure, stress, strain, and displacement values during extrusion of the solid rocket propellant are obtained from the simulation.

  2. Multi-Dimensional Combustion Instability Analysis of Solid Propellant Rocket Motors.


    RI D-R159 314 MULTI-DIMENSIONAL COMBUSTION INSTABLITY ANALYSIS OF 1/1 I SOLID PROPELLANT ROCK.. (U) ALABAMA UNIY IN HUNTSVILLE I DEPT OF MECHANICAL...STANDARDS MlICROCOPY RESOLUTION TEST CHART 0 0 0 03 V.%% f iSR.TR. 85-0567 NULTI-DIMNSIONAL COMBUSTION INSTABILITY ANALYSIS OF SOLID PROPELLANT ROCKET...analysis of solid propellant rocket motors. This research was motivated by the need for im- provement of the current practice in combustion instability

  3. Mathematical Modelling of In-Chamber Processes in Hydrocombined Propellant Solid Rocket Motors

    Nikolai A. Obukhov


    Full Text Available The special conditions of employment of commercial rockets in the sea environment has opened up new possibilities of improving motor performance. The interesting method suggests supplying water into the running motor. This paper reports the calculations and experiments carried out with solid propellant model setups. The results prove the validity of the proposed method and allow the refinement of calculation techniques for the prediction of solid rocket motor performance characteristics. The serviceability of the solid propellant charges working in combination with water is demonstrated. A mathematical model is proposed for the operation of a hydrocombined propellant motor with water and powdered additives applied to the combustion chamber."

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

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


    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.

  5. Estimation of Pressure Index and Temperature Sensitivity Coefficient of Solid Rocket Propellants by Static Evaluation

    Himanshu Shekhar


    Full Text Available Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T-To}]Pn was considered and a method was evolved for determination of pressure index (n and temperature sensitivity coefficient (p of burning rate for solid rocket propellants from static evaluation data. Effect of pressure index and temperature sensitivity coefficient on firing curve is also depicted. Propellant grain was fired in progressive mode to cover a very wide pressure range of 50 kg/cm2 to 250 kg/cm2 and propellant burning rate index was calculated to be 0.32 in the given pressure range. Propellant grain was fired at +35 °C and –20 °C temperatures and temperature sensitivity coefficient of burning rate was calculated to be 0.27 % per °C. Since both the values were evaluated from realised static evaluation curves, these are more realistic and accurate compared to data generated by conventional methods.Defence Science Journal, 2009, 59(6, pp.666-669, DOI:

  6. Design of Multi-Propellant Star Grains for Solid Propellant Rockets

    S. Krishnan


    Full Text Available A new approach to solve the geometry-problem of solid propellant star is presented. The basis of the approach is to take the web-thickness (a ballistic as well as a geometrical property as the characteristic length. The nondimensional characteristic parameters representing diameter, length, slenderness-ratio, and ignitor accommodation of the grain are all identified. Many particular cases of star configurations (from the configurations of single propellant to those of four different propellants can be analysed through the identified characteristic parameters. A better way of representing the single-propellant-star-performance in a design graph is explained. Two types of dual propellant grains are analysed in detail. The first type is characterised by its two distinct stages of burning (initially by single propellant burning and then by dual propellant burning; the second type has the dual propellant burning throughout. Suitability of the identified characteristic parameters to an optimisation study is demonstrated through examples.

  7. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.


    Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  8. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.; Borisov Boris V.; Bondarchuk Sergey S.; Zhukov Alexander S.


    On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  9. Development of high temperature materials for solid propellant rocket nozzle applications

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


    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.

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

    Baran, W. J.


    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.

  11. Radiographic Sensitivity of Flaw Detection in Solid Rocket Propellants

    G.G Puranik


    Full Text Available The sensitivity of flaw detection with x-ray radiographic methods is investigated here qualitatively in case of cast double base and cast composite propellants and for air pockets it is found to be 1.5 and 0.9 percent of the web respectively. General guidelines for the inspection of sustainer charges have also been laid down.

  12. Discrimination for ablative control mechanism in solid-propellant rocket nozzle


    The ablation in solid-propellant rocket nozzle is a coupling process resulted by chemistry, heat and mass transfer. Based on the heat and mass transfer theory, the aero-thermo-dynamic, and thermo-chemical kinetics, the thermal-chemical ablation model is established. Simulations are completed on the heat flow field and chemical ablation in the nozzle with different concentrations, frequency factors and activation energy of H2. The calculation results show that the concentration and the activation energy of H2 can provoke the transformation of control mechanism, whereas the influence brought by the frequency factor of H2 is feeble under a high-temperature and high-pressure combustion circumstance. The discrimination for ablative control mechanism is dependent on both concentration and activation energy of H2. This study will be useful in handling ablation and thermal protection problem in the design of solid-propellant rocket.

  13. Metallised Fuel rich Propellants for Solid Rocket Ramjet: A Review

    B. K. Athawale


    Full Text Available This paper reviews the research work carried out in the field of metallised fuel-rich propellants (FRP. Limitations and merits of various potential metals (Al, Mg, B, Be, or Zr as a component of FRP are discussed. The paper also includes a discussion on the combustion mechanism of metallised propellants, including problem areas and probable remedial measures. Zirconium and Ti appear to have potential to offer FRP with efficient combustion. Ideal performance is not achieved with current systems based on Al and B and further work is needed to develop FRP having all three desirable attributes, viz., ease of ignition, stable combustion and high specific impulse (I/sub sp/ in a single composition.

  14. Studies on Stress-Strain Curves of Aged Composite Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical property evaluation of composite solid rocket propellants is used as a quick quality control tool for propellant development and production. However, stress-strain curves from uni-axial tensile testing can be utilised to assess the shelf-life of propellants also. Composite propellants (CP of two varieties cartridge-loaded (CLCP and case-bonded (CBCP are utilized in rocket and missile applications. Both classes of propellants were evaluated for mechanical properties namely tensile strength, modulus and percentage elongation using specimens conforming to ASTM D638 type IV at different ageing time. Both classes of propellants show almost identical variation in various mechanical properties with time. Tensile strength increases with time for both classes of propellants and percentage elongation reduces. Initial modulus is also found to decrease with time. Tensile strength is taken as degradation criteria and it is observed that CLCP has slower degradation rate than CBCP. This is because of two facts–(i higher initial tensile strength of CLCP (1.39 MPa compared to CBCP (0.665 MPa and (ii lower degradation rate of CLCP (0.0014 MPa/day with respect to CBCP (0.0025 MPa/day. For the studied composite propellants, a degradation criterion in the form of percentage change in tensile strength is evaluated and shelf life for different degradation criteria is tabulated for quick reference.Defence Science Journal, 2012, 62(2, pp.90-94, DOI:

  15. Regarding the evaluation of the solid rocket propellant response function to pressure coupling

    Ioan ION


    Full Text Available High frequency combustion instabilities imply a major risk for the solid rocket motor stableworking and they are directly linked to the propellant response to chamber pressure coupling. Thisarticle discusses a laboratory testing method for the measurement and evaluation of the pressurecoupled response for non-metalized propellants in a first stage. Experimental researches were donewith an adequate setup, built and improved in our lab, able to evaluate the propellant response byinterpreting the pressure oscillations damping in terms of propellant response. Our paper aims atdefining a linearized one-dimensional flow study model to analyze the disturbed operation of the solidpropellant rocket motors. Based on the applied model we can assert that the real part of propellantresponse is a function of the oscillations damping, acoustic energy in the motor chamber and variouslosses in the burning chamber. The imaginary part of propellant response mainly depends on thenormalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency.Our research purpose was obviously to minimize the risk of the combustion instabilities effects on therocket motors working, by experimental investigations using jet modulating techniques and sustainedby an interesting study model based on the perturbation method.

  16. Mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene

    Gligorijević Nikola I.


    Full Text Available This paper presents the procedure of uniaxial mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene (HTPB, whose mechanical properties strongly depend on temperature, strain rate, natural aging and accumulated damage. A method of processing data is presented in order to determine time-temperature shift factor and master curves for tensile strength, ultimate strain and relaxation modulus, depending on reduced time. Functional dependences of these features represent an input for structural analysis of a rocket motor propellant grain. The effects of natural aging on the mechanical properties are also considered. [Projekat Ministarstva nauke Republike Srbije, br. TR 36050: Research and development of unmanned aircraft in support of traffic infrastructure monitoring

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


    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.

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

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


    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.

  19. Solid propellant motor

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


    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.

  20. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  1. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  2. New discrimination method for ablative control mechanism in solid-propellant rocket nozzle


    A reasonable discrimination method for ablative control mechanism in solid-propellant rocket nozzle can improve the calculation accuracy of ablation rate. Based on the different rate constants for reactions of C with H2O and CO2,a new discrimination method for ablative control mechanism,which comprehensively considers the influence of nozzle surface temperature and gas component concentration,is presented. Using this new discrimination method,calculations were performed to simulate the nozzle throat insert ablation. The numerical results showed that the calculated ablation rate,which was more close to the measured values,was less than the value calculated by diffusion control mechanisms or by double control mechanisms. And H2O was proved to be the most detrimental oxidizing species in nozzle ablation.

  3. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

  4. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

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

    Heitkotter, Robert H


    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.

  6. Navier-Stokes calculation of solid-propellant rocket motor internal flowfields

    Hsieh, Kwang-Chung; Yang, Vigor; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the detailed physical and chemical processes involved in the combustion of homogeneous propellant in a rocket motor. The formulation is based on the time-dependent full Navier-Stokes equations, with special attention devoted to the chemical reactions in both gas and condensed phases. The turbulence closure is achieved using both the Baldwin-Lomax algebraic model and a modified k-epsilon two-equation scheme with a low Reynolds number and near-wall treatment. The effects of variable thermodynamic and transport properties are also included. The system of governing equations are solved using a multi-stage Runge-Kutta shceme with the source terms treated implicitly. Preliminary results clearly demonstrate the presence of various combustion regimes in the vicinity of propellant surface. The effects of propellant combustion on the motor internal flowfields are investigated in detail.

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

    Verderaime, V.


    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.

  8. Alternate Propellant Thermal Rocket Project

    National Aeronautics and Space Administration — The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by...

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

    Terry, Brandon C., E-mail: [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Sippel, Travis R. [Department of Mechanical Engineering, Iowa State University, 2025 Black Engineering, Ames, IA 50011 (United States); Pfeil, Mark A. [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Gunduz, I.Emre; Son, Steven F. [School of Mechanical Engineering, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States)


    Highlights: • Al-Li alloy propellant has increased ideal specific impulse over neat aluminum. • Al-Li alloy propellant has a near complete reduction in HCl acid formation. • Reduction in HCl was verified with wet bomb experiments and DSC/TGA-MS/FTIR. - Abstract: 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 (I{sub SP}). 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 I{sub SP} by ∼7 s 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.

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

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


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

  11. Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant%Nmerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

    SHAFQAT Wahab; XIE Kan; LIU Yu


    Multi-phase flow field simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM)is in- vestigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and form liquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flow simulations have been performed on a large scale motor,to observe the effect of the flowfield on the chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an expo- nential relationship exists between the droplet diameter,d and mass fraction of droplets with diameter greater than d have been used for the simulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-1 00μm are used,as being the most common droplets.In this approach the complete range of particle sizes is divided into a set of discrete size ranges,each to be defined by single stream that is part of the group.Roe scheme-flux differencing splitting based on approximate Riemann problem has been used to simulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect, Spalart-Allmaras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRM flow dynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the parti- cles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

  12. Solid Propellant Grain Structural Integrity Analysis


    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.

  13. Nuclear magnetic resonance imaging of solid rocket propellants at 14.1 T.

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


    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 x 8.5 micron. 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 micron which is found to surround some of the larger filler particles.

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

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


    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.

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


    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.

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

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


    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

    Bennett, Robert R.


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

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


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

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

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


    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.

  20. Composite Solid Propellant Predictability and Quality Assurance

    Ramohalli, Kumar


    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. Nanoparticles for solid rocket propulsion

    Galfetti, L [Politecnico di Milano, SPLab, Milan (Italy); De Luca, L T [Politecnico di Milano, SPLab, Milan (Italy); Severini, F [Politecnico di Milano, SPLab, Milan (Italy); Meda, L [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marra, G [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marchetti, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Regi, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Bellucci, S [INFN, Laboratori Nazionali di Frascati, Frascati (Italy)


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  2. Nanoparticles for solid rocket propulsion

    Galfetti, L.; DeLuca, L. T.; Severini, F.; Meda, L.; Marra, G.; Marchetti, M.; Regi, M.; Bellucci, S.


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

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

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


    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.

  4. Assessment of Poisson's Ratio for Hydroxy-terminated Polybutadine-based Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Poisson's ratio of hydroxy-terminated polybutadine (HTPB-based composite propellant is estimated from uni-axial tensile testing. Double dumbbell specimens as per ASTM D638 type IV standard were used and Poisson's ratio at break, obtained by change in volume of specimen, was calculated as approximately 0.25. It was also observed that Poisson's ratio is different along different lateral directions of the propellant specimen. Poisson's ratios in two orthogonal directions perpendicular to longitudinal axis were calculated as 0.17 and 0.30. As ASTM specimen has rectangular cross-section of approximate size 6 mm x 4 mm, the directional behaviour of Poisson's ratio may be attributed to initial dimensions. Prismatic propellant specimen with square cross-section and of 115 mm x 6 mm x 6 mm dimension do not show any variation wrt Young's modulus,tensile strength, and percentage elongation as compared to ASTM specimen. Directional behaviour of Poisson's ratio with almost similar numerical value was again observed, thus ruling out dependence of this behaviour on different initial dimensions of propellant cross-section. Further, Poisson's ratio varies linearly with strain even in linear portion of stress-strain curve in uni-axial tensile testing. The rate of reduction of Poisson's ratio with increase in strain is slower in linear region and it accelerates after dewetting due to formation of vacuoles. Variation of Poisson's ratio with strain has two different slopes in linear (slope = 0.3165 and nonlinear regions (slope = 0.61364. Numerical value of slope for variation of Poisson's ratio with strain almost doubles after dewetting. It must be noted that no change in volume does not necessarily indicate constant Poisson's ratioequal to 0.5. Composite propellants behave as compressible material in most of the regions and near-failure region or at higher strains; Poisson's ratio is not anywhere near to 0.5, instead it is near 0.25.Defence Science

  5. The mixing of solid propellant by an artificial muscle actuator

    岩崎, 祥大; 伴, 遼介; 吉浜, 舜; 中村, 太郎; 羽生, 宏人; Iwasaki, Akihiro; Ban, Ryosuke; Yoshihama, Shun; Nakamura, Taro; Habu, Hiroto


    This research aims to reduce the cost of the solid rocket motor production, mainly solid propellant. The production process of the solid rocket propellant are usually employed the multi-batch mixing. However, this study using a peristaltic pump as a mixer will lead to the continuous process. The pump system can mix the powder materials for propellant and we consider that it will make the slurry of the solid propellant efficiently by the mechanism of the fluid dynamics in the pump.

  6. Velocity-coupled flow oscillations in a simulated solid-propellant rocket environment

    Yang, Vigor; Hsieh, Kwang-Chung; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the unsteady flowfields in a simulated rocket-motor environment. The model is based on the time-dependent compressible Navier-Stokes equations with a two-equation turbulence closure scheme. Various important aspects of the coupling between acoustic oscillations and mean flowfields, including flow reversal, modification of transport properties, etc., are addressed. Results indicate that multi-dimensional effects play important roles in determining local flow structures and wave characteristics. In much of the domain, acoustic velocity nodal points are observed in the near-wall region. The classical one-dimensional theory fails to describe several important mechanisms associated with velocity-induced flow instabilities.

  7. Analysis of velocity-coupled response function data from the dual rotating valve. [combustion stability of solid rocket propellants

    Brown, R. S.; Waugh, R. C.


    The results of a re-evaluation of the propellant combustion data obtained using the dual valve approach for measuring velocity-coupling characteristics of solid propellants are presented. Data analysis and testing procedures are described. The velocity response is compared to pressure-coupled response data within the context of thermal wave response theory. This comparison shows important inconsistencies which cast doubt on inferring the velocity response from pressure-coupled response functions.

  8. Integrated model of a composite propellant rocket

    Miccio, Francesco


    The combustion of composite solid propellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

  9. Measuring Combustion Advance in Solid Propellants

    Yang, L. C.


    Set of gauges on solid-propellant rocket motor with electrically insulating case measures advance of combustion front and local erosion rates of propellant and insulation. Data furnished by gauges aid in motor design, failure analysis, and performance prediction. Technique useful in determining propellant uniformity and electrical properties of exhaust plum. Gauges used both in flight and on ground. Foilgauge technique also useful in basic research on pulsed plasmas or combustion of solids.

  10. Ammonium nitrate: a promising rocket propellant oxidizer

    Oommen; Jain


    Ammonium nitrate (AN) is extensively used in the area of fertilizers and explosives. It is present as the major component in most industrial explosives. Its use as an oxidizer in the area of propellants, however, is not as extensive as in explosive compositions or gas generators. With the growing demand for environmental friendly chlorine free propellants, many attempts have been made of late to investigate oxidizers producing innocuous combustion products. AN, unlike the widely used ammonium perchlorate, produces completely ecofriendly smokeless products. Besides, it is one of the cheapest and easily available compounds. However, its use in large rocket motors is restricted due to some of its adverse characteristics like hygroscopicity, near room temperature phase transformation involving a volume change, and low burning rate (BR) and energetics. The review is an attempt to consolidate the information available on the various issues pertaining to its use as a solid propellant oxidizer. Detailed discussions on the aspects relating to phase modifications, decomposition chemistry, and BR and energetics of AN-based propellants, are presented. To make the review more comprehensive brief descriptions of the history, manufacture, safety, physical and chemical properties and various other applications of the salt are also included. Copyright 1999 Elsevier Science B.V.

  11. Enantioselective synthesis of aziridines using asymmetric transfer hydrogenation as a precursor for chiral derivatives used as bonding agent for rocket solid propellants

    Aparecida M. Kawamoto


    Full Text Available A rapid, expedient and enantioselective method for the synthesis of beta-hydroxy amines and monosubstituted aziridines in up to 99% e.e., via asymmetric transfer hydrogenation of a-amino ketones and cyclisation through treatment with tosyl chloride and base, is described. (1R,2R-N-(para-toluenesulfonyl-1,2-ethylenediamine with formic acid has been utilised as a ligand for the Ruthenium (II catalysed enantioselective transfer hydrogenation of the ketones.The chiral 2-methyl aziridine, which is a potentially more efficient bonding agent for Rocket Solid Propellant has been successfully achieved.

  12. Reducing Thrusts In Solid-Fuel Rockets

    Bement, Laurence J.


    Thrust-terminating system conceived to reduce thrust of solid-propellant rocket motor in controlled manner such that thrust loads not increased or decreased beyond predictable levels. Concept involves explosively cutting opposing venting pairs in case of rocket motor above nozzles to initiate venting of chamber and reduction of thrust. Vents sized and numbered to control amount and rate of reduction in thrust.

  13. Technology Method Design of Assembly and Testing for Solid Propellant Rocket Engine of Aviation Seat%航空座椅固体火箭发动机装配及检测工艺技术设计



    本文对航空座椅某型固体火箭发动机部装、总装及检测、试验、包装技术难点等进行了工艺分析;介绍了固体火箭发动机装配全过程工艺流程、检测、试验方法及注意事项等,对于同类及新型火箭发动机的装配制造过程具有良好的借鉴、推广应用意义。%Aiming at the difficulty of solid propellant rocket engine of aviation seat to assembly, testing and packaging technology, the assembly, testing process and method for solid propellant rocket engine were introduced. It can be regarded as reference with application for solid propellant rocket engine assembly process.

  14. Advanced Solid Rocket Launcher and Its Evolution

    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.

  15. Solid Rocket Booster-Illustration


    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.

  16. Propellant removal from rocket motors containing double-base compositions

    Whinnery, L.; Griffiths, S.; Hruby, J.; Larson, R.; Lipkin, J.; Long, B.; Schoenfelder, C.


    The uncertain environmental consequences and regulations associated with using open burning/open detonation for the disposal of energetic materials are forcing both manufacturers and users to examine alternative disposal technologies. In general, these alternatives involve a material removal operation followed by processing steps that lead to reuse of valuable constituents and/or disposal of waste. While a number of post-removal processing options appear to be viable, the initial step of removing an energetic material, such as a solid rocket motor propellant, from its container remains a significant technological challenge. Large rocket motors containing highly energetic propellant, hazard class 1.1, are of particular concern because of their inherent handling hazards. We will describe the results of a study using thermal cycling to increase the surface area of inert propellant formulations. The propellant removal method studied employs thermal cycling to cryogenic temperatures (cryocycling). Using inert propellants and liquid nitrogen we have demonstrated that this process produces multiple cracks throughout the bulk of the grain. The properties of the actual and inert propellants are being measured, and a model is being developed to relate experiments on inert material to actual propellant. Possible methods to increase thermal gradients, crack propagation and initiation are also presented.

  17. Dynamic mechanical analysis of double base rocket propellants

    Marcin Cegła


    Full Text Available The article presents dynamic mechanical analysis (DMA for solid rocket propellants testing. Principles of operation and measured values are briefly described. The authors refer to the previous research of PTFE material and literature data providing information about proper experimental conditions and influence of measurement frequency, load amplitude, and heating rate on the results of DMA tests. The experimental results of solid double-base rocket propellant testing obtained on the N Netzsch DMA 242 device are presented. Mechanical properties such as the dynamic storage modulus E´, the dynamic loss modulus E˝ and tan(δ were measured within temperature range from (–120°C to (+90°C at the heating rate of 1 K/min. The test sample was subjected to a dual cantilever multi-frequency test. Special attention was paid to determination of the glass transition temperature of the tested propellant in reference to the NATO standardization agreement 4540 as well as influence of the measurement frequency on the glass transition.[b]Keywords[/b]: Dynamic mechanical analysis, solid rocket propellants, glass transition temperature

  18. Solid Propellant Test Article (SPTA) Test Firing


    The Marshall Space Flight Center (MSFC) engineers test fired a 26-foot long, 100,000-pound-thrust solid rocket motor for 30 seconds at the MSFC east test area, the first test firing of the Modified NASA Motor (M-NASA Motor). The M-NASA Motor was fired in a newly constructed stand. The motor is 48-inches in diameter and was loaded with two propellant cartridges weighing a total of approximately 12,000 pounds. The purpose of the test was to learn more about solid rocket motor insulation and nozzle materials and to provide young engineers additional hands-on expertise in solid rocket motor technology. The test is a part of NASA's Solid Propulsion Integrity Program, that is to provide NASA engineers with the techniques, engineering tools, and computer programs to be able to better design, build, and verify solid rocket motors.

  19. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    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.

  20. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    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.

  1. Numerical Analysis on the Thermal Safety of Solid Rocket Motor Propellant%固体发动机装药热安全性数值分析

    刘文一; 焦冀光


    Objective To investigate the safety of solid rocket motor(SRM)when it was cook-off. Methods Finite element model of solid SRM was established, and the temperature distribution and the explosion delay time of propellant in fast cook-off mode and slow cook- off mode were computed. Results Propellant reached its critical temperature (352 ℃) after 47 h slow cook-off, while it reached its critical temperature (355 ℃) after 697 s fast cook-off. Conclusion It was proven that the thermal diffusivity in fast cook-off mode was greater than that in slow cook-off mode, while the temperature gradient had an opposite trend. The reaction position of propellant was different in the two different working modes when it reached critical temperature, and the thermal storage capacity of propellant was dependent on its thickness.%目的:研究固体火箭发动机遭受火烤时的安全性。方法建立发动机有限元模型,计算推进剂在慢速烤燃和快速烤燃工况下的温度分布和爆炸延迟时间。结果推进剂慢烤47 h后达到临界温度,其值为352℃;快烤推进剂加热697 s后达到临界温度,临界温度为355℃。结论推进剂在快速烤燃模式下的热扩散速率大于慢速烤燃工况下,但是温度梯度则相反。两种工况下推进剂达到临界温度后开始反应的位置不同,推进剂厚度决定了其储热能力。

  2. A research on polyether glycol replaced APCP rocket propellant

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


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

  3. Rheology of composite solid propellants during motor casting

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


    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.

  4. Method of injecting fluid propellants into a rocket combustion chamber

    Schneider, Steven J. (Inventor)


    A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connected to valves between these manifolds and propellant storage tanks. Additional valves are opened to furnish propellants to more of the manifolds when cryogenic propellant temperatures are sensed. Only a portion of the valves are opened to furnish propellants to some of the manifolds when lower temperatures are sensed.

  5. Investigation of the role of aziridine bonding agents on the aging of the composite solid rocket propellant(CSRP)

    Amged A Ali; ZHANG Jian-wei; CAI Guo-biao


    The role of bonding agents on the aging characteristics is one of the important research topics.Aging program of the prepared propellant samples was conducting as follows:Five samples,two free of bonding agents,and three containing an aziridine based bonding agents (MAPO,HX-752,MAT4),four samples based on different bonding and curing agents all were aged at 70℃.The prepared bonding agent"MAT4"gave remarkable improvements and resulted in highly stable mechanical properties comparing with HX-752 or MAPO.The selected bonding agents family inhibited the rate of decomposition of the propellants during the aging periods and supported the propellant matrix against decomposition at the elevated temperatures.

  6. FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor%固体火箭发动机柔性喷管有限元建模及摆动分析

    王成林; 刘勇; 文立华


    The methods of FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor were explores. The FEM model of flexible nozzle using the equivalent model of flexible joint based on the three-direction custom spring elements is build; modify the FEM model of the flexible joint according to the experiment data; and carry out the oscillation analysis of solid propellant rocket motor.%研究了固态火箭发动机柔性喷管有限元建模及摆动分析方法.利用基于自定义三向弹簧单元的柔性接头线性等效模型,建立了发动机柔性喷管有限元模型.根据试验数据对柔性接头模型进行修正,并对发动机柔性喷管进行了摆动分析.

  7. Ignition transient analysis of solid rocket motor

    Han, Samuel S.


    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.

  8. Fuel-Cell Power Source Based on Onboard Rocket Propellants

    Ganapathi, Gani; Narayan, Sri


    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.

  9. The development of space solid rocket motors in China

    Jianding, Huang; Dingyou, Ye


    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

    Zubrin, Robert M.


    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. Solid Propellant Flame Spectroscopy


    400 jm to reach the maximum flame temperature, a distance that can be reduced by replacing the HTPB binder with a polyester or CMDB binder. The...the dark zone for propellants similar to HIX2 is 2-2.5 mm at 1.8 MPa (18 atm, 265 psia) (Ref. 22,187). In contrast, the dark zone for HMX CMDB ...propellants eliminates the dark zone is not surprising, since TMETN is a nitrate ester as was the double-base matrix of Kubota’s HMX CMDB propellant. A

  12. Processing solid propellants for recycling

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


    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.

  13. Innovative Metallized Formulations for Solid Rocket Propulsion

    Luigi T DeLUCA; Luciano GALFETTI; Filippo MAGGI; Giovanni COLOMBO; Alice REINA; Stefano DOSSI; Daniele CONSONNI; Melissa BRAMBILLA


    Several metallized solid rocket propellants,AP/Metal/HTPB in the ratio 68/18/1 4,were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano.Effects of the metals (micrometric and nanometric Al,B,Mg,and a variety of dual metals) on the performance of the propellant were studied and contrasted to a conventional micrometric aluminum (30 μm average grain size) taken as reference.It is shown that the propellant microstructure plays a fundamental role in controlling the critical aggregation/agglomeration phenomena occurring below and near the burning surface.Two specific effects of microstructure in terms of steady burning rate and average agglomerate size are illustrated.

  14. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    National Aeronautics and Space Administration — Spark ignition of a bi-propellant rocket engine is a classic, proven, and generally reliable process. However, timing can be critical, and the control logic,...

  15. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

  16. Atmospheric scavenging of solid rocket exhaust effluents

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


    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.

  17. NASA's Advanced solid rocket motor

    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

    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.

  19. Solid Hydrogen Formed for Atomic Propellants

    Palaszewski, Bryan A.


    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.

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

    Hufferd, W. L.


    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.

  1. Solid Propellant Test Article (SPTA) Test Stand


    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.

  2. Erosive burning of solid propellants

    King, Merrill K.


    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.

  3. Burn-back Equations for High Volumetric Loading Single-grain Dual-thrust Rocket Propellant Configuration (Review Paper

    Himanshu Shekhar


    Full Text Available Dual-thrust mode is adopted in solid propellant rocket propulsion through tailoring of burning area, nozzle, rocket motor chamber, propellant type, multiple propellant blocks. In the present study, mathematical formulation has been evolved for generation of burning surface area with web burnt for a simple central blind hole in a solid cylindrical propellant geometry with proper partial inhibition on external and lateral surfaces. The burn-back equation has been validated by static firing and parametric study was conducted to understand effect of various control geometrical parameters. The system is utilised for high volumetric loading, single propellant, single composition, single-chamber, single nozzle dual-thrust mode of burning profiles in rocket application.Defence Science Journal, 2011, 61(2, pp.165-170, DOI:

  4. Refinement of Propellant Strand Burning Method to Suit Aluminised Composite Rocket Propellant


    ABSTRACT An epoxy coating was trialled as an inhibitor for composite rocket propellant strands burned in a Crawford–style bomb . The epoxy coating...into the Crawford bomb and pressurised with nitrogen to the specified pressure. Once pressure and temperature were stable at the required set- point...aggressive propellant. This method involved the coating of strands of propellant with a two-part epoxy polymer. Strands were dipped into a bath of epoxy

  5. Combustion chemistry of solid propellants

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


    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.

  6. Hybrid Rocket Technology

    Sankaran Venugopal; K K Rajesh; V Ramanujachari


    With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems...

  7. History of solid propellants in the 20. century; Histoire des propergols solides au 20. siecle

    Jung, Ph.; Davenas, A.; McDonald, A.J.; Bret, P.; Moreau, J.P.; Boisson, J.; Kuentzmannn, P.; Maire, G.; Pontvianne, G.; Tranchant, J.; Evans, G.; Reydellet, D.; Vallet, G.; Eymard, M.; Pascal, Ph.; Kuentzmann, P.; Bonnevie, E.; Guery, J.F.; Lengelle, G.; Lhuillier, J.N.; Rat, R.; Keromnes, A.; Mathieu, D.; Simonetti, Ph.; Betin, P.; Thevenin, M.; Serra, J.J.; Delbac, P.; Lepeuple, G.; Miermont, H.; Guillot, J.; Vidal, M.; Citon, C.; Tauzia, J.M.; Chounet, G.; Cardin, J.; Longevialle, Y.; Uhrig, G.


    This colloquium has been jointly organized by the research center of history of sciences and techniques (CRHST) and the association of the friends of the gunpowder and pyrotechnical patrimony (A3P). It gathers historians of sciences and techniques and specialists of solid propellants and their applications who make a review of the approaches that have led to todays propellants efficiency and mastery. This books contains 2 introductive talks, 24 articles, a round table and some concluding remarks. The articles deal with: 1 - from the black powder rockets to the space shuttle: France, pioneer of solid propulsion, from Vaillant to Damblanc (1821-1938); the development of solid propellants in the 20. century; lessons learnt from the Challenger accident; 2 - the institutions: the laboratory of ballistics of Sevran-Livry (1945-1969); an historical overview of ONERA's researches on solid propellants; the cast propellants at the Direction of Explosives (1945-1955); 3 - the propellants: the manufacturing secrets of the extruded double base propellants; the development of cast double base propellants; the invention of composite propellants; 4 - space applications: the Diamant adventure; the solid propellant engines of Ariane 5, an endless story; P80, a new generation of solid propellant engines for space applications; 6 - physics and models: from ap{sup n} to 3-D simulations: the combustion of solid propellants in the 20. century; the mechanical behaviour of solid propellant loads (1960-70 years); composite propellants and static electricity (SE) or the occurrence of SE in the manufacturing and implementation of composite propellants; a priori calculation of the performances and synthesis of new energy materials for propellants; 6 - defense applications: French solid propellant rockets and missiles up to the 1960's; from PHI 1500 to PHI 1930 or the fabulous history of metallic and roving propulsion systems; the G2P, the propulsion system of the M4, the exploratory

  8. Solid Rocket Launch Vehicle Explosion Environments

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


    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.

  9. Study of Liquid Breakup Process in Solid Rocket Motor Nozzle


    34Chemical Erosion of Refractory - Metal Nozzle Inserts in Solid-Propellant Rocket Motors," J. Propulsion and Power, Vol. 25, no.1,, 2009. [4] E. Y. Wong...Paul A.;, "Gelcasting of Alumina," J. Am. Ceram . Soc. 74[3], pp. 612-618, 1991. [18] Blomquist , B. A.; Fink, J. K.; Leibowitz, L.;, "The

  10. Studies of solid propellant combustion with pulsed radiography

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


    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.

  11. Ultrasonic investigation of mechanical properties of double base rocket propellants

    Schroeff, J.A. van der; Boer, R.S. de


    For a series of double base rocket propellants and for poly-methylmethacrylate (PMMA) the longitudinal and transverse sound wave velocities are measured at a frequency of 0.351 MHz in t h e temperature range of −40°C to +60°C. The relations between these acoustic properties and mechanical properties

  12. Magnetic resonance imaging (MRI) of solid rocket components

    Wallner, A.S. [Missouri Western State College, St. Joseph, MO (United States); Nissan, R.A.; Merwin, L.H. [Naval Air Warfare Center, China Lake, CA (United States)] [and others


    The evaluation of solid rocket components has become an area of great interest. Studying these materials with MRI offers a great advantage to observe knit lines, regions of inhomogeneity, voids, defects, plasticizer rich/poor areas and solids distribution because of the nondestructive nature of the technique. Aspects of sample preparation, spectroscopic relaxation studies, and MRI as a method of studying these systems will be discussed. Initial images show the ability to image propellant, liner, and explosive materials with an in-plane resolution of 70 {mu}m/pixel. These initial images show that MRI can be developed as a viable nondestructive evaluation method of solid rocket components.

  13. Metallic Hydrogen: A Game Changing Rocket Propellant

    Silvera, Isaac F.


    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.

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

    Chapman, John S.; Nix, Michael B.


    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.

  15. Advanced Computer Science on Internal Ballistics of Solid Rocket Motors

    Shimada, Toru; Kato, Kazushige; Sekino, Nobuhiro; Tsuboi, Nobuyuki; Seike, Yoshio; Fukunaga, Mihoko; Daimon, Yu; Hasegawa, Hiroshi; Asakawa, Hiroya

    In this paper, described is the development of a numerical simulation system, what we call “Advanced Computer Science on SRM Internal Ballistics (ACSSIB)”, for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solid propellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.

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

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


    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.

  17. Scaling theory for liquid propellant rocket thrust chambers

    C. M. Sethna


    Full Text Available With the advent of the very large liquid propellant rocket, it has become necessary, if possible, to derive a rational scaling theory for combustion chamber design so as to enable relatively simple and economical initial tests to be carried out on small scaled models using scaled parameters of propellant mass flows, pressure etc., and from these to predict operating and design data for the full scale rocket. Owing to the complex and interdependent nature of the aerothermo-chemical processes in the chamber involving evaporation, diffusion and chemical reaction, the similarity criteria must necessarily extend over several, non-dimensional parameters, but it is still possible to evolve relatively simple rules for correlating the design and performance of the model and large scale motors-as shown by Penner-Tsien and Crocco. The paper concludes with a discussion of the accuracy and practical feasibility of such scaling rules.

  18. Characterization of rocket propellant combustion products

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


    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.

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

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


    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.

  20. Laboratory test methods for combustion stability properties of solid propellants

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


    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.

  1. Internal Flow Simulation of High-Performance Solid Rockets using a k-ωTurbulence Model



    @@ For technological reasons many high-performance solid rocket motors are made from segmented propellant grains with non-uniform port geometry. In this paper parametric studies have been carried out to examine the geometric dependence of transient flow features in solid rockets with non-uniform ports. Numerical computations have been carried out in an inert simulator of solid propellant rocket motor with the aid of a standard k-ω turbulence model. It was seen that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that the damping of the flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the starting transient of solid rockets.

  2. Thermodynamic cycle analysis on solid propellant air-turbo-rocket%固体推进剂吸气式涡轮火箭发动机的气动热力循环分析

    屠秋野; 丁朝霞; 陈玉春; 蔡元虎


    A numerical model for calculating performance of solid propellant air-turbo-rocket at design point was set up,and a relationship expression of fuel-air ratio of combustion based on compressor pressure ratio,tubine intet total temperature and turbine expansion ratio was put forward.And a relationship between the turbine expansion ratio and the bypass ratio was given.The effects of compressor pressure ratio, turbine inlet temperature,bypass ratio/turbine expansion ratio and flight Mach number on the specific thrust and specific impulse were analyzed quantitatively.%建立了固体推进剂吸气式涡轮火箭发动机的设计状态数值模型,提出了基于压气机增压比、涡轮前温度和涡轮落压比关系的燃烧室燃气与空气配比表达式,以及涡轮落压比和发动机涵道比的匹配关系.定量分析了压气机增压比、涡轮进口燃气总温、涵道比/涡轮落压比和飞行马赫数对固体推进剂吸气式涡轮火箭发动机的单位推力和比冲的影响.

  3. Development of hydrazinium nitroformate based solid propellants

    Schöyer, H.F.R.; Schnorhk, A.J.; Korting, P.A.O.G.; Lit, P.J. van; Mul, J.M.; Gadiot, G.; Meulenbrugge, J.J.


    The development of new high-performance propellant combinations requires the establishment of safety and handling characteristics and thermodynamic decomposition and explosive properties. This paper addresses the early development phases of a new composite solid propellant based on HNF as oxidizer a

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

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


    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.

  5. On the hydrodynamics of rocket propellant engine inducers and turbopumps

    d'Agostino, L.


    The lecture presents an overview of some recent results of the work carried out at Alta on the hydrodynamic design and rotordynamic fluid forces of cavitating turbopumps for liquid propellant feed systems of modern rocket engines. The reduced order models recently developed for preliminary geometric definition and noncavitating performance prediction of tapered-hub axial inducers and centrifugal turbopumps are illustrated. The experimental characterization of the rotordynamic forces acting on a whirling four-bladed, tapered-hub, variable-pitch high-head inducer, under different load and cavitation conditions is presented. Future perspectives of the work to be carried out at Alta in this area of research are briefly illustrated.

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

    Sutton, George P.

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

  7. Development of a solid propellant viscoelastic dynamic model

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


    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.

  8. Development of small solid rocket boosters for the ILR-33 sounding rocket

    Nowakowski, Pawel; Okninski, Adam; Pakosz, Michal; Cieslinski, Dawid; Bartkowiak, Bartosz; Wolanski, Piotr


    This paper gives an overview of the development of a 6000 Newton-class solid rocket motor for suborbital applications. The design configuration and results of interior ballistics calculations are given. The initial use of the motor as the main propulsion system of the H1 experimental in-flight test platform, within the Polish Small Sounding Rocket Program, is presented. Comparisons of theoretical and experimental performance are shown. Both on-ground and in-flight tests are discussed. A novel composite-case manufacturing technology, which enabled to reach high propellant mass fractions, was validated and significant cost-reductions were achieved. This paper focuses on the process of adapting the design for use as the booster stage of the ILR-33 sounding rocket, under development at the Institute of Aviation in Warsaw, Poland. Parallel use of two of the flight-proven rocket motors along with the main stage is planned. The process of adapting the rocket motor for booster application consists of stage integration, aerothermodynamics and reliability analyses. The separation mechanism and environmental impact are also discussed within this paper. Detailed performance analysis with focus on propellant grain geometry is provided. The evolution of the design since the first flights of the H1 rocket is covered and modifications of the manufacturing process are described. Issues of simultaneous ignition of two motors and their non-identical performance are discussed. Further applications and potential for future development are outlined. The presented results are based on the initial work done by the Rocketry Group of the Warsaw University of Technology Students' Space Association. The continuation of the Polish Small Sounding Rocket Program on a larger scale at the Institute of Aviation proves the value of the outcomes of the initial educational project.

  9. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    Crawford, T. Daniel


    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.

  10. High burn rate solid composite propellants

    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

  11. Storage of solid propellants in space

    Udlock, D. E.


    A test program is described which determines the extent of physical property changes that result from extended space exposure. Primary emphasis was placed on determining the effects of space vacuum. Solid propellants were stored and their physical properties tested in a vacuum and in a dry environment. The storage caused significantly greater increases in the propellants' modulus and maximum tensile strength than occurred in parallel ambient stored samples. The data indicate that the loss of trace amounts of residual moisture from cured propellant is the apparent cause of the observed propellant property changes. Therefore, initial screening tests were carried out under dry storage conditions. Upon completion of the dry storage tests, appropriate propellant samples are exposed to an actual space environment using the Long Duration Exposure Facility (LDEF).

  12. Ballistic anomalies in solid rocket motors due to migration effects

    Pröbster, M.; Schmucker, R. H.

    Double base and composite propellants are generally used for rocket motors, whereby double base propellants basically consist of nitrocellulose plasticized with an explosive plasticizer, mostly nitroglycerine, and in some cases with an additional inert plasticizer and ballistic additives. Composite propellants consist of an oxidizer like ammonium perchlorate and of aluminum, binder and plasticizer and often contain liquid or solid burning rate catalysts. A common feature of both propellants is that they contain smaller or larger amounts of chemically unbonded liquid species which tend to migrate. If these propellants loose part of the plasticizer by migration into the insulation layer, not only will there be a change in mechanical propellant properties but also the bond between propellant and insulation may degrade. However, depending on the severity of these effects, the change in the ballistic properties of the propellant grain caused by plasticizer migration may be of even more importance. In the past, most emphasis was placed on the behaviour of end-burning configurations. However, more recent theoretical and experimental studies revealed that not only for end-burning grain configurations but also for internal burning configurations there is a common effect which is responsible for ballistic anomalies: migration of liquid species from the propellant into the insulation. By using a plasticizer equilibrated insulation in an internal burning configuration the liquid species migration and thus the previously observed ballistic anomalies are avoided. Using this approach for end-burning configurations provides similar positive results. The various factors affecting plasticizer migration are studied and discussed, and several methods to prevent liquid species migration are described as well as methods to obtain plasticizer resistant insulations.

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

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


    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.

  14. Rheokinetic Analysis of Hydroxy Terminated Polybutadiene Based Solid Propellant Slurry

    Abhay K Mahanta


    Full Text Available The cure kinetics of propellant slurry based on hydroxy-terminated polybutadiene (HTPB and toluene diisocyanate (TDI polyurethane reaction has been studied by viscosity build up method. The viscosity (ɳ–time (t plots conform to the exponential function ɳ = aebt, where a & b are empirical constants. The rate constants (k for viscosity build up at various shear rate (rpm, evaluated from the slope of dɳ/dt versus ɳ plots at different temperatures, were found to vary from 0.0032 to 0.0052 min-1. It was observed that the increasing shear rate did not have significant effect on the reaction rate constants for viscosity build up of the propellant slurry. The activation energy (Eɳ, calculated from the Arrhenius plots, was found to be 13.17±1.78 kJ mole-1, whereas the activation enthalpy (∆Hɳ* and entropy (∆Sɳ* of the propellant slurry, calculated from Eyring relationship, were found to be 10.48±1.78 kJ mole-1 and –258.51± 5.38 J mole-1K-1, respectively. The reaction quenching temperature of the propellant slurry was found to be -9 ° C, based upon the experimental data. This opens up an avenue for a “freeze-and-store”, then “warm-up and cast”, mode of manufacturing of very large solid rocket propellant grains.

  15. Shuttle Redesigned Solid Rocket Motor aluminum oxide investigations

    Blomshield, Fred S.; Kraeutle, Karl J.; Stalnaker, Richard A.


    During the launch of STS-54, a 15 psi pressure blip was observed in the ballistic pressure trace of one of the two Space Shuttle Redesigned Solid Rocket Motors (RSRM). One possible scenario for the observed pressure increase deals with aluminum oxide slag formation in the RSRM. The purpose of this investigation was to examine changes which may have occurred in the aluminum oxide formation in shuttle solid propellant due to changes in the ammonium perchlorate. Aluminum oxide formation from three propellants, all having the same formulation, but containing ammonium perchlorate from different manufacturers, will be compared. Three methods have been used to look for possible differences among the propellants. The first method was to examine window bomb movies of the propellants burning at 100, 300 and 600 psia. The motor operating pressure during the pressure blip was around 600 psia. The second method used small samples of propellant which were fired in a combustion bomb which quenched the burning aluminum particles soon after they left the propellant surface. The bomb was fired in both argon and Nitrogen atmospheres at various pressures. Products from this device were examined by optical microscopy. The third method used larger propellant samples fired into a particle collection device which allowed the aluminum to react and combust more completely. This device was pressurized with Nitrogen to motor operating pressures. The collected products were subdivided into size fractions by screening and sedimentation and analyzed optically with an optical microscope. the results from all three methods indicate very small changes in the size distribution of combustion products.

  16. 14 CFR 420.65 - Handling of solid propellants.


    ... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Handling of solid propellants. 420.65....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...

  17. Fault Detection and Diagnosis Techniques for Liquid-Propellant Rocket Propellant Engines

    Wua, Jianjun; Tanb, Songlin


    Fault detection and diagnosis plays a pivotal role in the health-monitoring techniques for liquid- propellant rocket engines. This paper firstly gives a brief summary on the techniques of fault detection and diagnosis utilized in liquid-propellant rocket engines. Then, the applications of fault detection and diagnosis algorithms studied and developed to the Long March Main Engine System(LMME) are introduced. For fault detection, an analytical model-based detection algorithm, a time-series-analysis algorithm and a startup- transient detection algorithm based on nonlinear identification developed and evaluated through ground-test data of the LMME are given. For fault diagnosis, neural-network approaches, nonlinear-static-models based methods, and knowledge-based intelligent approaches are presented. Keywords: Fault detection; Fault diagnosis; Health monitoring; Neural networks; Fuzzy logic; Expert system; Long March main engines Contact author and full address: Dr. Jianjun Wu Department of Astronautical Engineering School of Aerospace and Material Engineering National University of Defense Technology Changsha, Hunan 410073 P.R.China Tel:86-731-4556611(O), 4573175(O), 2219923(H) Fax:86-731-4512301

  18. Unsteady Processes in Solid Propellant Combustion,


    0—AflO ~5a INSTITUTO NACIONAL DE TECNICA AEROESPACIAL MADRID (SPAIN) F/S 21/9.2UNSTEADY PROCESSES IN SOLID PROPELLANT COMBUSTION . (U) MAY...PRO C E SS E S IN SOLID P R O P E L L A N T C O M B U S T I O N H A. Crespo and M. Kindelán Instituto Nacional de Técnica Aeroespacial Madrid , Spain j

  19. CFD modelling of solid propellant ignition

    Lowe, C


    Solid propellant is the highly energetic fuel burnt in the combustion chamber of ballistic weapons. It is manufactured, for this purpose, in either granular or stick form. Internal ballistics describes the behavior within the combustion chamber throughout the ballistic cycle upto projectile exit from the muzzle of the gun barrel. Over the last twenty years this has been achieved by modelling the process using two-phase flow equations. The solid granules or sticks constitute ...

  20. Combustion Instabilities In Solid Propellant Rocket Motors


    34 AIAA Paper No. 98{3218. Avalon, G., Ugurtas, B. Grisch, F. and Bresson , F. (2000) \\Numerical Computations and Visualization Tests of the Flow Inside...Burners and Related Devices," Prog. in Comb. Energy and Science, Vol. 19, No. 4, pp. 313{364. Roberts , A.K. and Brownlee, W.G. (1971) \\Nonlinear

  1. Unsteady Aerodynamic Investigation of the Propeller-Wing Interaction for a Rocket Launched Unmanned Air Vehicle

    G. Q. Zhang


    Full Text Available The aerodynamic characteristics of propeller-wing interaction for the rocket launched UAV have been investigated numerically by means of sliding mesh technology. The corresponding forces and moments have been collected for axial wing placements ranging from 0.056 to 0.5D and varied rotating speeds. The slipstream generated by the rotating propeller has little effects on the lift characteristics of the whole UAV. The drag can be seen to remain unchanged as the wing's location moves progressively closer to the propeller until 0.056D away from the propeller, where a nearly 20% increase occurred sharply. The propeller position has a negligible effect on the overall thrust and torque of the propeller. The efficiency affected by the installation angle of the propeller blade has also been analyzed. Based on the pressure cloud and streamlines, the vortices generated by propeller, propeller-wing interaction, and wing tip have also been captured and analyzed.

  2. Scale Effects on Solid Rocket Combustion Instability Behaviour

    David R. Greatrix


    Full Text Available The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combustion response to pressure wave activity above the burning propellant surface, is applied to the investigation of scale effects (motor size, i.e., grain length and internal port diameter on influencing instability-related behaviour in a cylindrical-grain motor. The results of this investigation reveal that the motor’s size has a significant influence on transient pressure wave magnitude and structure, and on the appearance and magnitude of an associated base pressure rise.

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

    Doerr, DonaldF.


    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 (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 the state-of-the-art in personal protective equipment. Not only was long duration environmental control provided, but it was done without a high pressure vessel. The unit met human performance needs for attitude independence, oxygen stability, and relief of heat stress. This supercritical air (and oxygen) technology is suggested for microgravity applications in life support such as the Extravehicular Mobility Unit.

  4. The University of Arizona program in solid propellants

    Ramohalli, Kumar


    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.

  5. Numerical Simulation of a Liquid Propellant Rocket Motor

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


    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 extemal 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.

  6. Rocket propellant reorientation and fluid management used in space commercialization

    Hung, R. J.; Lee, C. C.; Shyu, K. L.


    In a spacecraft design, the requirements of settled propellant are different for tank pressurization, engine restart, venting, or propellant transfer. The requirement to settle or to position liquid fuel over the outlet end of the spacecraft propellant tank prior main engine restart possess a microgravity fluid behavior problem. In this paper, the dynamical behavior of liquid propellant, fluid reorientation, and propellant resettling have been carried out.

  7. Introduction to rocket science and engineering

    Taylor, Travis S


    What Are Rockets? The History of RocketsRockets of the Modern EraRocket Anatomy and NomenclatureWhy Are Rockets Needed? Missions and PayloadsTrajectoriesOrbitsOrbit Changes and ManeuversBallistic Missile TrajectoriesHow Do Rockets Work? ThrustSpecific ImpulseWeight Flow RateTsiolkovsky's Rocket EquationStagingRocket Dynamics, Guidance, and ControlHow Do Rocket Engines Work? The Basic Rocket EngineThermodynamic Expansion and the Rocket NozzleExit VelocityRocket Engine Area Ratio and LengthsRocket Engine Design ExampleAre All Rockets the Same? Solid Rocket EnginesLiquid Propellant Rocket Engines

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


    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.

  9. Method and device for protecting objects against Rocket Propelled Grenades (RPGs)

    Wentzel, C.M.


    A device and a method for protecting objects against rocket-propelled grenades having a hollow nose cone includes a netting of knotted and coated superstrong fibers disposed in front of the object, in such a manner that the nose cone of a rocket caught in the netting will penetrate one of the meshes

  10. Method and device for protecting objects against Rocket Propelled Grenades (RPGs)

    Wentzel, C.M.


    A device and a method for protecting objects against rocket-propelled grenades having a hollow nose cone includes a netting of knotted and coated superstrong fibers disposed in front of the object, in such a manner that the nose cone of a rocket caught in the netting will penetrate one of the meshes

  11. Solid rocket motor fire tests: Phases 1 and 2

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


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

  12. Solid State MEMS Thrusters Using Electrically Controlled Extinguishable Solid Propellant Project

    National Aeronautics and Space Administration — ET Materials, LLC developed the first ever electrically controlled extinguishable solid propellant (ECESP). The original propellant developed under Air Force SBIR...

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

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


    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.

  14. Characterization of booster-rocket propellants and their simulants

    Weirick, L.J.


    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.

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

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


    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

  16. Solid Hydrogen Experiments for Atomic Propellants

    Palaszewski, Bryan


    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.

  17. Feasibility Study and Demonstration of an Aluminum and Ice Solid Propellant

    Timothee L. Pourpoint


    Full Text Available Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solid propellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

  18. Study on the Detonation Danger of Solid Propellants

    黄风雷; 张宝(金平)


    A measurement system to study shock initiation behavior of solid propellants was established experimentally. By using this system, the study on shock initiation to the recovered solid propellants with micro damage was performed, especially on the deflagration to denonation transition (DDT) process of solid propellants under both the strong and weak conditions of restriction. The experimental results show that there is a fully compression region in DDT process.

  19. First Stage Solid Propellant Multiply Debris Thermal Analysis

    Toleman, Benjamin M.


    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.

  20. First Stage Solid Propellant Multi Debris Thermal Analysis

    Toleman, Benjamin M.


    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

  1. The Dynamics of Solid Propellant Combustion.


    gases of average molecular weight ’F. 3. One-phase, laminar, nonviscous, strongly subsonic flow. 4. Adiabatic, except at the burning surface. 5. No...Comosite Solid Rocket !i ropellantl", XXXIV Congresso Nazionalc Associaziono Term-a tecnica italiana (ATI) , Palcrino, 8-12 October 1979. Proceed- * i-ngs...1.000 E-04 cal/cm K s Average product molecular weight, 2.600 E+01 g/mole Optical surface reflectivity, rX 3.800 % Optical volumetric scattering, sX 0 an

  2. The 260: The Largest Solid Rocket Motor Ever Tested

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


    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.

  3. Design and Experimental Study on Spinning Solid Rocket Motor

    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.

  4. Ariane-5 solid-propellant stage development

    Gigou, Jacques


    The development status of the solid propellant engine (P230) of the Ariane-5 launcher is described. Large new industrial plants were built in Europe and Guiana for the development and manufacture of the solid-booster stage and are now operational. A product assurance policy, specific and common to the companies that are involved in the engine's development, was defined and will be implemented. The paper describes the production cycles for the charged segments, the igniter, and the nozzle for P230 engine, as well as the process of engine integration and testing. Consideration is also given to the engine thrust capability, the launcher flight control, and the interfaces. The the major engine development tests are described.

  5. An Evaluation Of Rocket Parameters

    J. N. Beri


    Full Text Available The dependence of conventional parameters of internal ballistics of Solid Propellant Rockets using external burning cruciform charge, on the geometry of charge aad rocket motor is discussed and results applied in a special case.

  6. Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

    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.

  7. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    National Aeronautics and Space Administration — Under a Phase 1 effort, IES successfully developed and demonstrated a spark ignition concept where propellant flow drives a very simple fluid mechanical oscillator...

  8. Advances in the research on the solid propellant properties abroad

    Du, Lei; Jiang, Zhirong


    The recent research on the mechanical properties, burning behavior and processing technology of solid propellants abroad was reviewed. There are some available results in predicting theoretically the mechanical and rheological properties of solid propellants. In order to reduce the cost and increase the reliability in propellants processing, there is great demand on the design and manufacture of continuous mixer of high efficiency and safety. The research on the thermoplastic elastomers used as a kind of future binder of solid propellants has attracted more and more attention of many relevant experts.

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

    Northam, G. B.


    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.

  10. Performance evaluation of commercial copper chromites as burning rate catalyst for solid propellants

    Milton Faria Diniz


    Full Text Available Copper chromites are well known as burning rate catalysts for the combustion of composite solid propellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET. The samples were then evaluated as burning rate catalyst in a typical composite propellant formulation based on HTPB binder, ammonium perchlorate and aluminum. The obtained surface area values are very close to those informed by the catalyst suppliers. The propellant processing as well as its mechanical properties were not substantially affected by the type of catalyst. Some copper chromite catalysts caused an increase in the propellant burning rate in comparison to the iron oxide catalyst. The results show that in addition to the surface area, other parameters like chemical composition, crystalline structure and the presence of impurities might be affecting the catalyst performance. All evaluated copper chromite samples may be used as burning rate catalyst in composite solid propellant formulations, with slight advantages for the SX14, Cu-0202P and Cu-1800P samples, which led to the highest burning rate propellants.

  11. Combustion response modeling for composite solid propellants


    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. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

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


    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.

  13. Natural Rubber Based Fuel Rich Propellant for Ramjet Rocket

    Solomon Thomas; T. L. Varghese; Gupta, S. K.; T. S. Ram; V. N. Krishnamurthy


    Development of fuel rich propellants for air-breathing propulsion is one of the frontier areas of research. This carries less oxidiser than the normal propellants, uses ram air for complete combustion and is heavily metallised. Studies were conducted at VSSC for the development of hydroxyl-terminated natural rubber (HTNR)-based Mg and Mg-Al alloy-filled polyurethane systems. The HTNR prepolymer was modified to make it free of acid functionality so as to avoid reaction with the metals a...

  14. Laser-induced fluorescence in high pressure solid propellant flames.

    Edwards, T; Weaver, D P; Campbell, D H


    The application of laser-induced fluorescence (LIF) to the study of high pressure solid propellant flames is described. The distribution of the OH and CN radicals was determined in several solid propellant flames at pressures up to 3.5 MPa. The greatest difficulty in these measurements was the separation of the desired LIF signals from the large scattering at the laser wavelength from the very optically thick propellant flames. Raman experiments using 308-nm excitation were also attempted in the propellant flames but were unsuccessful due to LIF interferences from OH and NH.

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


    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.

  16. Studies on composite solid propellant with tri-modal ammonium perchlorate containing an ultrafine fraction

    K.V. Suresh Babu


    composite solid propellant is prepared by using burn rate modifiers Copper chromite and Iron oxide. Addition of Copper chromite and Iron oxide has enhanced the burn rate of tri-modal AP based composite solid propellant. The catalytic propensity of copper chromite is higher than that of iron oxide. The pressure exponent increased with the catalyst concentration and the values obtained are compatible for solid rocket motor applications.

  17. Natural Rubber Based Fuel Rich Propellant for Ramjet Rocket

    Solomon Thomas


    Full Text Available Development of fuel rich propellants for air-breathing propulsion is one of the frontier areas of research. This carries less oxidiser than the normal propellants, uses ram air for complete combustion and is heavily metallised. Studies were conducted at VSSC for the development of hydroxyl-terminated natural rubber (HTNR-based Mg and Mg-Al alloy-filled polyurethane systems. The HTNR prepolymer was modified to make it free of acid functionality so as to avoid reaction with the metals and was then reacted with calculated amount of propylene oxide in the presence of tertiary amine catalyst at room temperature. Various levels of Mg and Mg-Al alloy were used in different compositions and the effect on propellant characteristics was studied. TMP and butane diol were used for higher crosslinking, chain extension, etc and their effects were evaluated. Propellant compositions using 30-35 per cent by weight of Mg/Mg-Al alloy with the modified prepolymer have been developed. They exhibit good processibility and mechanical properties. The feasibility of room temperature curing was attempted. Motors weighing 2 kg with 100 mm OD and 200 mm length were successfully static-tested in the primary mode for performance evaluation.

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

    Martin, Heath Thomas


    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.

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

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


    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.

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

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


    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.

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


    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.

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

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


    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

  3. Crusader solid propellant best technical approach

    Graves, V. [Oak Ridge National Lab., TN (United States); Bader, G. [Fire Support Armament Center, Picatinny Arsenal, NJ (United States); Dolecki, M. [Tank-Automotive Research, Development, and Engineering Center, Picatinny, NJ (United States); Krupski, S. [Benet Weapons Lab., Watervliet Arsenal, NY (United States); Zangrando, R. [Close Combat Armament Center, Picatinny Arsenal, NJ (United States)


    The goal of the Solid Propellant Resupply Team is to develop Crusader system concepts capable of automatically handling 155mm projectiles and Modular Artillery Charges (MACs) based on system requirements. The system encompasses all aspects of handling from initial input into a resupply vehicle (RSV) to the final loading into the breech of the self-propelled howitzer (SPH). The team, comprised of persons from military and other government organizations, developed concepts for the overall vehicles as well as their interior handling components. An intermediate review was conducted on those components, and revised concepts were completed in May 1995. A concept evaluation was conducted on the finalized concepts, from both a systems level and a component level. The team`s Best Technical Approach (BTA) concept was selected from that evaluation. Both vehicles in the BTA have a front-engine configuration with the crew situated behind the engine-low in the vehicles. The SPH concept utilizes an automated reload port at the rear of the vehicle, centered high. The RSV transfer boom will dock with this port to allow automated ammunition transfer. The SPH rearm system utilizes fully redundant dual loaders. Active magazines are used for both projectiles and MACs. The SPH also uses a nonconventional tilted ring turret configuration to maximize the available interior volume in the vehicle. This configuration can be rearmed at any elevation angle but only at 0{degree} azimuth. The RSV configuration is similar to that of the SPH. The RSV utilizes passive storage racks with a pick-and-place manipulator for handling the projectiles and active magazines for the MACs. A telescoping transfer boom extends out the front of the vehicle over the crew and engine.

  4. Crusader solid propellant best technical approach

    Graves, V. [Oak Ridge National Lab., TN (United States); Bader, G. [Fire Support Armament Center, Picatinny Arsenal, NJ (United States); Dolecki, M. [Tank-Automotive Research, Development, and Engineering Center, Picatinny, NJ (United States); Krupski, S. [Benet Weapons Lab., Watervliet Arsenal, NY (United States); Zangrando, R. [Close Combat Armament Center, Picatinny Arsenal, NJ (United States)


    The goal of the Solid Propellant Resupply Team is to develop Crusader system concepts capable of automatically handling 155mm projectiles and Modular Artillery Charges (MACs) based on system requirements. The system encompasses all aspects of handling from initial input into a resupply vehicle (RSV) to the final loading into the breech of the self-propelled howitzer (SPH). The team, comprised of persons from military and other government organizations, developed concepts for the overall vehicles as well as their interior handling components. An intermediate review was conducted on those components, and revised concepts were completed in May 1995. A concept evaluation was conducted on the finalized concepts, from both a systems level and a component level. The team`s Best Technical Approach (BTA) concept was selected from that evaluation. Both vehicles in the BTA have a front-engine configuration with the crew situated behind the engine-low in the vehicles. The SPH concept utilizes an automated reload port at the rear of the vehicle, centered high. The RSV transfer boom will dock with this port to allow automated ammunition transfer. The SPH rearm system utilizes fully redundant dual loaders. Active magazines are used for both projectiles and MACs. The SPH also uses a nonconventional tilted ring turret configuration to maximize the available interior volume in the vehicle. This configuration can be rearmed at any elevation angle but only at 0{degree} azimuth. The RSV configuration is similar to that of the SPH. The RSV utilizes passive storage racks with a pick-and-place manipulator for handling the projectiles and active magazines for the MACs. A telescoping transfer boom extends out the front of the vehicle over the crew and engine.

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

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


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

  6. The Chameleon Solid Rocket Propulsion Model

    Robertson, Glen A.


    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).

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

    Weirick, L.J.


    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.

  8. Process and quality verification controls for solid propellant manufacturing

    Rogers, C. J.


    It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.

  9. Process and quality verification controls for solid propellant manufacturing

    Rogers, C. J.


    It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.

  10. Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

    Jaime Rubio Hervas


    Full Text Available The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

  11. State-space analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine

    Zhang, Yu-Lin

    This paper states the application of state-space method to the analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine and presents a set of state equations for describing the dynamic process of the engine. An efficient numerical method for solving these system equations is developed. The theoretical solutions agree well with the experimental data. The analysis leads to the following conclusion: the set coefficient of the pulse width, the working frequency of the solenoid valves and the deviation of the critical working points of these valves are important parameters for determining the dynamic response time and the control precision of this engine. The methods developed in this paper may be used effectively in the analysis of dynamic characteristics of variable thrust liquid propellant rocket engines.

  12. Development of composite solid propellent using dicyclopentadien binder

    Bluestone, Stephen Ray

    Through the history of composite solid propellant binders new chemicals are introduced as binders to improve upon the previous generation. Sometimes this is done to improve upon the flaws or shortcomings of a previous binder. Other time it is to meet a new set of requirements desired by industry. Dicyclopentadiene (DCPD) is a hydrocarbon monomer being considered for its potential as a new binder in the composite propellant industry. The binder of a composite solid propellant is arguably the most important feature of the propellant. It is the binder that provides the majority of the structural characteristics of the propellant while also contributing itself as fuel to the combustion process. A binder in composite propellants must also be able to accept the introduction of a large quantity of solid filler; oxidizer, fuel, and other energetic and non-energetic particles. Many of the composite propellants used in industry today have over 80% of their weight composed of non-binder solid or liquid fillers. These requirements must be met by the binder in some form or fashion to produce a propellant able to compete with binders currently in use. When DCPD is polymerized it produces an extremely tough plastic with excellent tensile and impact strength. Experimentation has found that DCPD is able to support a large quantity of solid materials, over 80% weight of the mixture, while still retaining a great portion of its original strength. When compared to another similarly loaded binder currently used in industry, Hydroxyl-Terminated Polybutadiene (HTPB), it was found that DCPD composite propellant had nearly 1.5 times the stress capacity while still exhibiting over 75% of the strain capacity of HTPB based composite propellant. In addition it was also shown that DCPD composite propellant allows for tailoring of its mechanical properties with the addition of plasticizers. The DCPD based composite propellant also exhibits a burning rate nearly twice that HTPB. These factors

  13. Analytical model for liquid rocket propellant feedline dynamics

    Holster, J. L.; Astleford, W. J.


    A generalized analytical model and computer program have been developed to predict the frequency response of arbitrary liquid propellant feedline designs. The analytical model is based on an extension of an existing distributed parameter representation of a viscous fluid transmission line with laminar flow which was modified to include the effects of a turbulent mean flow. The effects of dissolved ullage gases, wall elasticity, localized gas or vapor bubbles, bellows, forced changes in length due to structural excitation, complex side branches, and structural mounting stiffness are also included. Each line component is written as a four-terminal, pressure-flow relationship in matrix form in the Laplace domain; the transfer function relating the pressure response at the line terminal (inducer inlet) to the external excitation is obtained in the computer program by sequential matrix substitution.

  14. Solid Rocket Motor Design Using Hybrid Optimization

    Kevin Albarado


    Full Text Available A particle swarm/pattern search hybrid optimizer was used to drive a solid rocket motor modeling code to an optimal solution. The solid motor code models tapered motor geometries using analytical burn back methods by slicing the grain into thin sections along the axial direction. Grains with circular perforated stars, wagon wheels, and dog bones can be considered and multiple tapered sections can be constructed. The hybrid approach to optimization is capable of exploring large areas of the solution space through particle swarming, but is also able to climb “hills” of optimality through gradient based pattern searching. A preliminary method for designing tapered internal geometry as well as tapered outer mold-line geometry is presented. A total of four optimization cases were performed. The first two case studies examines designing motors to match a given regressive-progressive-regressive burn profile. The third case study studies designing a neutrally burning right circular perforated grain (utilizing inner and external geometry tapering. The final case study studies designing a linearly regressive burning profile for right circular perforated (tapered grains.

  15. Low loss injector for liquid propellant rocket engines

    Vonpragenau, G. L. (Inventor)


    A low pressure loss injector element is disclosed for the main combustion chamber of a rocket engine which includes a lox post terminating in a cylindrical barrel. Received within the barrel is a lox plug which is threaded in the lox post and includes an interchangeable lox metering sieve which meters the lox into an annular lox passage. A second annular gas passage is coaxial with the annular lox passage. A cylindrical sleeve surrounds the annular gas passage and includes an interchangeable gas metering seive having metering orifices through which a hot gas passes into the annular passage. The jets which emerge from the annular lox passage and annular gas passage intersect in a recessed area away from the combustion area. Thus, mixing and combustion stability are enhanced.

  16. Computational Fluid Dynamics Simulation of Combustion Instability in Solid Rocket Motor : Implementation of Pressure Coupled Response Function

    S. Saha


    Full Text Available Combustion instability in solid propellant rocket motor is numerically simulated by implementing propellant response function with quasi steady homogeneous one dimensional formulation. The convolution integral of propellant response with pressure history is implemented through a user defined function in commercial computational fluid dynamics software. The methodology is validated against literature reported motor test and other simulation results. Computed amplitude of pressure fluctuations compare closely with the literarture data. The growth rate of pressure oscillations of a cylindrical grain solid rocket motor is determined for different response functions at the fundamental longitudinal frequency. It is observed that for response function more than a critical value, the motor exhibits exponential growth rate of pressure oscillations.

  17. The Effect of Solid Constituent Particle Size Distributions on TP-H1148 Propellant Slag

    May, Douglas H.; Miles, William L.; Taylor, David S.; Rackham, Jon L.


    Special aluminum and ammonium perchlorate (AP) particle size distributions were prepared for a matrix of five-inch diameter, center-perforated (CP) motor tests to measure the aluminum oxide slag response in Space Shuttle Reusable Solid Rocket Motor (RSRM) propellant. Previous tests of TP-H1148 propellant in five-inch CP spin motors have shown a correlation between aluminum particle size and generated slag. The motors for this study were cast from thirteen five-gallon propellant mixes which used five particle size levels of aluminum powder, five of unground AP and three of ground AP. Aluminum had the greatest effect on slag formation, the more coarse fractions causing greater slag quantities and larger slag particles. Unground AP had about half the effect of aluminum with the coarser fractions again producing more and larger sized slag particles. The variation in ground AP did not have a significant effect on slag formation. Quench bomb tests showed the same trends as the spin motors, that is, larger aluminum and AP particle size distributions generated larger slag particles leaving the propellant surface. Cured propellant mechanical properties were also impacted by particle size variation.

  18. Low-Cost High-Performance Non-Toxic Self-Pressurizing Storable Liquid Bi-Propellant Pressure-Fed Rocket Engine Project

    National Aeronautics and Space Administration — Exquadrum proposes a high-performance liquid bi-propellant rocket engine that uses propellants that are non-toxic, self-pressurizing, and low cost. The proposed...

  19. Studies of Fission Fragment Rocket Engine Propelled Spacecraft

    Werka, Robert O.; Clark, Rodney; Sheldon, Rob; Percy, Thomas K.


    The NASA Office of Chief Technologist has funded from FY11 through FY14 successive studies of the physics, design, and spacecraft integration of a Fission Fragment Rocket Engine (FFRE) that directly converts the momentum of fission fragments continuously into spacecraft momentum at a theoretical specific impulse above one million seconds. While others have promised future propulsion advances if only you have the patience, the FFRE requires no waiting, no advances in physics and no advances in manufacturing processes. Such an engine unequivocally can create a new era of space exploration that can change spacecraft operation. The NIAC (NASA Institute for Advanced Concepts) Program Phase 1 study of FY11 first investigated how the revolutionary FFRE technology could be integrated into an advanced spacecraft. The FFRE combines existent technologies of low density fissioning dust trapped electrostatically and high field strength superconducting magnets for beam management. By organizing the nuclear core material to permit sufficient mean free path for escape of the fission fragments and by collimating the beam, this study showed the FFRE could convert nuclear power to thrust directly and efficiently at a delivered specific impulse of 527,000 seconds. The FY13 study showed that, without increasing the reactor power, adding a neutral gas to the fission fragment beam significantly increased the FFRE thrust through in a manner analogous to a jet engine afterburner. This frictional interaction of gas and beam resulted in an engine that continuously produced 1000 pound force of thrust at a delivered impulse of 32,000 seconds, thereby reducing the currently studied DRM 5 round trip mission to Mars from 3 years to 260 days. By decreasing the gas addition, this same engine can be tailored for much lower thrust at much higher impulse to match missions to more distant destinations. These studies created host spacecraft concepts configured for manned round trip journeys. While the

  20. ASRM radiation and flowfield prediction status. [Advanced Solid Rocket Motor plume radiation prediction

    Reardon, J. E.; Everson, J.; Smith, S. D.; Sulyma, P. R.


    Existing and proposed methods for the prediction of plume radiation are discussed in terms of their application to the NASA Advanced Solid Rocket Motor (ASRM) and Space Shuttle Main Engine (SSME) projects. Extrapolations of the Solid Rocket Motor (SRM) are discussed with respect to preliminary predictions of the primary and secondary radiation environments. The methodology for radiation and initial plume property predictions are set forth, including a new code for scattering media and independent secondary source models based on flight data. The Monte Carlo code employs a reverse-evaluation approach which traces rays back to their point of absorption in the plume. The SRM sea-level plume model is modified to account for the increased radiation in the ASRM plume due to the ASRM's propellant chemistry. The ASRM cycle-1 environment predictions are shown to identify a potential reason for the shutdown spike identified with pre-SRM staging.

  1. Thermal Decomposition Behavior and Non-isothermal Decomposition Reaction of Copper(Ⅱ) Salt of 4-Hydroxy-3,5-dinitropyridine Oxide and Its Application in Solid Rocket Propellant

    CHEN Pei陈沛; ZHAO Feng-Qi赵凤起; LUO Yang罗阳; HU Rong-Zu胡荣祖; GAO Sheng-Li高胜利; ZHENG Yu-Mei郑玉梅; DENG Min-Zhi邓敏智; GAO Yin高茵


    The thermal decomposition behavior and kinetic parameters of the exothermic decomposition reactions of the title compound in a temperature-programmed mode have been investigated by means of DSC, TG-DTG and lower rate thermolysis/FTIR. The possible reaction mechanism was proposed. The critical temperature of thermal explosion was calculated. The influence of the title compound on the combustion characteristic of composite modified double base propellant containing RDX has been explored with the strand burner. The results show that the kinetic model function in differential form, apparent activation energy Ea and pre-exponential factor A of the major exothermal explosion of the compound is 312.87 C. The kinetic equation of the major exothermic decomposition process of the title compound at 0.1 MPa could be expressed as:dα/dT=1016.42(1-α)e-2.502× 104/T. As an auxiliary catalyst, the title compound can help the main catalyst lead salt of 4-hydroxy-3,5-dinitropyridine oxide to enhance the burning rate and reduce the pressure exponent of RDX-CMDB propellant.

  2. Combustion Behaviour of Advanced Solid Propellants.

    S. N. Asthana


    Full Text Available The study reports the effect of incorporation of Al and ammonium perchlorate (AP individually and in combination with each other on combustion pattern and specific impulse (Isp of minimum signature propellants. Incorporation of Al obviates the combustion instability problems; however, it has marginal effect on burning rates. The composition containing AP and zirconium silicate combination gives superior performance; however, its Isp is considerably lower than the composition incorporating 9 per cent AP. A combination of 6 per cent Al gave 20 per cent enhancement in burning rate and 12 s increase in Isp as compared to purely nitramine-based composition, cal-val results also reveal increase in energy output on incorporating AP and Al. Hot stage microscopic and propellant combustion studies indicate occurrence of intense decomposition reaction in case of AP-based compositions.

  3. Study of Solid Propellant Combustion under External Radiation

    V.E. Zarko


    Full Text Available The influence of constant and transient radiant flux on the burning rate of solid propellants is considered. The validity of the equivalence principle for the radiant flux and increase in initial temperature and also the problem of possible photochemical effect of thermal radiation are discussed. Experimental data on burning rate response to periodical perturbations of radiant flux for different types of solid propellants are reported. The problem of correlation between burning rate response to perturbations of pressure and external radiation is considered. Formulation of the problem on transient combustion in terms of the Zeldovich- Novozhilov phenomenological approach is described and the results of numerical integration are presented.

  4. Assessment of analytical techniques for predicting solid propellant exhaust plumes

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


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

  5. New Frontiers AO: Advanced Materials Bi-propellant Rocket (AMBR) Engine Information Summary

    Liou, Larry C.


    The Advanced Material Bi-propellant Rocket (AMBR) engine is a high performance (I(sub sp)), higher thrust, radiation cooled, storable bi-propellant space engine of the same physical envelope as the High Performance Apogee Thruster (HiPAT(TradeMark)). To provide further information about the AMBR engine, this document provides details on performance, development, mission implementation, key spacecraft integration considerations, project participants and approach, contact information, system specifications, and a list of references. The In-Space Propulsion Technology (ISPT) project team at NASA Glenn Research Center (GRC) leads the technology development of the AMBR engine. Their NASA partners were Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL). Aerojet leads the industrial partners selected competitively for the technology development via the NASA Research Announcement (NRA) process.

  6. Some Observations on the Ignition of Composite Solid Propellants

    K. Kishore


    Full Text Available Heat-up times derived from studies on the ignition characteristics of a few model composite solid propellants, containing polystyrene, carboxy-terminated polybutadiene, plasticised polyvinyl chloride and polyphenol formaldehyde as binders, show that they are directly proportional to the mass of the sample and inversely proportional to the heat flux. Propellant weight-loss prior to ignition and high pressure ignition temperature data on the propellants, ammonium per chlorate, and binders show that the ignition is governed by the gasification of the binder pyrolysis products. The activation energy for the gasification of the pyrolysed polymer products corresponds to their ignition behaviour suggesting that propellant ignition is controlled by the binder.

  7. C/C-SiC Composites for Nozzle of Solid Propellant Ramjet

    WANG Lingling


    Full Text Available Carbon fiber reinforced carbon and silicon carbide matrix composites for nozzle inner of solid propellant ramjet were prepared by using the hybrid process of "chemical vapor infiltration + precursor impregnation pyrolysis (CVI+PIP". The microstructure, flexural and anti-ablation properties of the C/C-SiC composites and hydraulic test and rocket motor hot firing test for nozzle inner of solid propellant ramjet were comprehensively investigated. The results show that when the flexural strength of the composite reachs 197 MPa, the fracture damage behavior of the composites presents typical toughness mode.Also the composites has excellent anti-ablative property, i.e., linear ablation rate is only 0.0063 mm·s-1 after 200 s ablation. The C/C-SiC component have excellent integral bearing performance with the hydraulic bursting pressure of 6.5 MPa, and the high temperature combination property of the C/C-SiC composite nozzle inner is verified through motor hot firing of solid propellant ramjet.

  8. Holographic Investigation of Solid Propellant Particulates.


    4~ .A*4 ~.Zwe SOUMVV Ch.&4 0IVC&TIN 0 e*9 066so. 4 evt’ o R..e High speed, high resolution motion pictures were taken to compare the cinematic data...propellant. High speed, high resolution motion pictures were taken to compare the cinematic data with that available from the holograms. TABLE employ finely powered aluminum (1-50 microns) in an attempt to capitalize on the conversion of its high heat of formation to kinetic energy for

  9. Characterisation of Materials used in Flex Bearings of Large Solid Rocket Motors

    CH.V. Ram Mohan


    Full Text Available Solid rocket motors are propulsion devices for both satellite launchers and missiles, which require guidance and steering to fly along a programmed trajectory and to compensate for flight disturbances. A typical solid rocket motor consists of motor case, solid propellant grain, motor insulation, igniter and nozzle. In most solid rocket motors, thrust vector control (TVC is required. One of the most efficient methods of TVC is by flex nozzle system. The flex nozzle consists of a flexible bearing made of an elastomeric material alternating with reinforcement rings of metallic or composite material. The material characterisation of AFNOR 15CDV6 steel and the natural rubber-based elastomer developed for use in flex nozzle are discussed. This includes testing, modelling of the material, selection of a material model suitable for analysis, and the validation of material model.Defence Science Journal, 2011, 61(3, pp.264-269, DOI:

  10. Developments on HNF based high performance and green solid propellants

    Keizers, H.L.J.; Heijden, A.E.D.M. van der; Vliet, L.D. van; Welland-Veltmans, W.H.M.; Ciucci, A.


    Worldwide developments are ongoing to develop new and more energetic composite solid propellant formulations for space transportation and military applications. Since the 90's, the use of HNF as a new high performance oxidiser is being reinvestigated. Within European development programmes, signific

  11. Advanced technologies available for future solid propellant grains

    Thépénier, Jean; Fonblanc, Gilles


    Significant advances have been made during the last decade in several fields of solid propulsion: the advances have enabled new savings in the motor development phase and recurring costs, because they help limit the number of prototypes and tests. The purpose of the paper is to describe the improvements achieved by SNPE in solid grain technologies, making these technologies available for new developments in more efficient and reliable future SRMs: new energetic molecules, new solid propellants, new processes for grain manufacturing, quick response grain design tools associated with advanced models for grain performance predictions. Using its expertise in chemical synthesis, SNPE develops new molecules to fit new energetic material requirements. Tests based on new propellant formulations have produced good results in the propellant performance/safety behavior ratio. New processes have been developed simultaneously to reduce the manufacturing costs of the new propellants. In addition, the grain design has been optimized by using the latest generation of predictive theoretical tools supported by a large data bank of experimental parameters resulting from over 30 years' experience in solid propulsion: Computer-aided method for the preliminary grain design Advanced models for SRM operating and performance predictions

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

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


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

  13. Scale Effects on Quasi-Steady Solid Rocket Internal Ballistic Behaviour

    David R. Greatrix


    Full Text Available The ability to predict with some accuracy a given solid rocket motor’s performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size. The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor’s internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor’s internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale.

  14. Characterization of Energetic Porous Silicon for a Microelectromechanical System (MEMS)-Based Solid Propellant Microthruster


    Characterization of Energetic Porous Silicon for a Microelectromechanical System (MEMS)-Based Solid Propellant Microthruster by Raghav...Energetic Porous Silicon for a Microelectromechanical System (MEMS)-Based Solid Propellant Microthruster Raghav Ramachandran, Wayne Churaman, David...Microelectromechanical System (MEMS)-Based Solid Propellant Microthruster 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6

  15. 14 CFR 420.69 - Solid and liquid propellants located together.


    ... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Solid and liquid propellants located... Licensee § 420.69 Solid and liquid propellants located together. (a) A launch site operator proposing an explosive hazard facility where solid and liquid propellants are to be located together shall determine...

  16. Solid Rocket Testing at AFRL (Briefing Charts)



  17. Viscoelastic properties of hydroxyl-terminated poly(butadiene based composite rocket propellants

    Brzić Saša J.


    Full Text Available In the present study, the viscoelastic response of three composite solid propellants based on hydroxyl-terminated poly(butadiene, ammonium perchlorate and aluminum has been investigated. The investigation was surveyed by dynamic mechanical analysis over a wide range of temperatures and frequencies. The mechanical properties of these materials are related to the macromolecular structure of the binder as well as to the content and nature of solid fillers. The storage modulus, loss modulus, loss factor and glass transition temperature for each propellant sample have been evaluated. The master curves of storage (log G' vs log ω and loss modulus (log G'' vs log ω were generated for each propellant. A comparison of logaT vs temperature curves for all propellants indicate conformance to Williams-Landel-Ferry equation. Choosing the glass transition as the reference temperature, WLF equation constants are determined. Fractional free volume at the glass transition temperature and thermal coefficient of free volume expansion values are in accordance with the consideration that Al is reinforcing filler.

  18. Consumption of stabilizers by aging in NEPE solid propellant

    Yim, Y.J.; Ryoo, B.N.; Park, Y.C.; Jeong, B.H.; Park, M.K. [Agency for Defense Development, Taejon (Korea, Republic of)


    The depletion rates of NMA and 2-NDPA were investigated by accelerated aging test in NEPE solid propellant containing BTTN and DEGDN as nitrate ester plasticizers. It was found that both NMA and 2-NDPA were depleted by the 0{sup th} order reactions and these reactions had two kinds of activation energy which showed the break-point at 60 {sup o} C for NMA and 70 {sup o} C for 2-NDPA from Arrhenius plots in temperature range of 20-80 {sup o} C. Therefore, the stabilizer content in propellant could be predicted much better by using reaction rate in the low temperature range than that in the high temperature range. The gas fissuring was not occurred under mild conditions like slow evacuation of gases from decomposition of nitrate ester plasticizers even though NMA and 2-NDPA were completely depleted in propellant.

  19. Characterization of rocket propellant combustion products: Description of sampling and analysis methods for rocket exhaust characterization studies

    Jenkins, R.A.


    A systematic approach has been developed and experimentally validated for the sampling and chemical characterization of the rocket motor exhaust generated from the firing of scaled down test motors at the US Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama. The overall strategy was to sample and analyze major exhaust constituents in near real time, while performing off-site analyses of samples collected for the determination of trace constituents of the particulate and vapor phases. Initial interference studies were performed using atmospheric pressure burns of 1 g quantities of propellants in small chambers at Oak Ridge National Laboratory. Carbon monoxide and carbon dioxide were determined using non-dispersive infrared instrumentation. Hydrogen cyanide, hydrogen chloride, and ammonia determinations were made using ion selective electrode technology. Oxides of nitrogen were determined using chemiluminescence instrumentation. Airborne particulate mass concentration was determined using infrared forward scattering measurements and a tapered element oscillating microbalance, as well as conventional gravimetry. Particulate phase metals were determined by collection on Teflon membrane filters, followed by inductively coupled plasma and atomic absorption analysis. Particulate phase polynuclear aromatic hydrocarbons (PAH) and nitro-PAH were collected using high volume sampling on a two stage filter. Target species were extracted, and quantified by gas chromatography/mass spectrometry (GC/MS). Vapor phase species were collected on multi-sorbent resin traps, and subjected to thermal desorption GC/MS for analysis. 11 refs., 1 fig., 1 tab.

  20. Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Yang, H. Q.; West, Jeff; Harris, Robert E.


    Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor

  1. Stratospheric aluminum oxide. [possibly from solid-fuel rocket exhausts

    Brownlee, D. E.; Tomandl, D.; Ferry, G. V.


    Balloons and U-2 aircraft were used to collect micrometer-sized stratospheric aerosols. It was discovered that for the past 6 years at least, aluminum oxide spheres have been the major stratospheric particulate in the size range from 3 to 8 micrometers. The most probable source of the spheres is the exhaust from solid-fuel rockets.

  2. STS-27 Atlantis, OV-104, solid rocket booster (SRB) inspection


    Engineers, kneeling inside a hollow solid rocket booster (SRB), closely inspect the SRB segments and seams in the Kennedy Space Center (KSC) rotation and processing facility. The SRB will be used on STS-27 Atlantis, Orbiter Vehicle (OV) 104. The booster segments were transported via rail car from Morton Thiokol's Utah manufacturing plant. View provided by KSC with alternate number KSC-88PC-492.

  3. [Combustion temperature measurement of solid propellant by remote sensing FTIR].

    Li, Yan; Wang, Jun-De; Sun, Xiu-Yun; Zhou, Xue-Tie


    The combustion temperature of solid propellant was measured in this paper. Emission spectra of the combustion flame were collected with remote sensing FTIR at the resolution of 4 cm(-1). The combustion temperatures with the burning time were calculated from the maximum spectral line intensity and the molecular rotation-vibration spectra of HF molecule, respectively. Combustion temperatures at each time were all 1 788.8 K from the maximum spectral line intensity method. For comparison, the temperatures calculated from the molecular rotation-vibration spectra were 1 859.7, 1 848. 3, 1 804.0 and 1 782.7 K, respectively. Results show that the two methods are all dependable in measuring combustion temperature of solid propellant. But the maximum spectral line intensity method is more convenient and rapid than the other when the combustion is relatively stable.

  4. Thermo-mechanical concepts applied to modeling liquid propellant rocket engine stability

    Kassoy, David R.; Norris, Adam


    The response of a gas to transient, spatially distributed energy addition can be quantified mathematically using thermo-mechanical concepts available in the literature. The modeling demonstrates that the ratio of the energy addition time scale to the acoustic time scale of the affected volume, and the quantity of energy added to that volume during the former determine the whether the responses to heating can be described as occurring at nearly constant volume, fully compressible or nearly constant pressure. Each of these categories is characterized by significantly different mechanical responses. Application to idealized configurations of liquid propellant rocket engines provides an opportunity to identify physical conditions compatible with gasdynamic disturbances that are sources of engine instability. Air Force Office of Scientific Research.

  5. Characterization of typical platelet injector flow configurations. [liquid propellant rocket engines

    Hickox, C. E.


    A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.

  6. Storable Hypergolic Solid Fuel for Hybrid Rocket Engines

    R. V. Singh


    Full Text Available A solid fuel was synthesised by condensing aniline with furfuraldehyde. The product was directly cast in the rocket motor casing. After curing a hard solid mass was obtained. This was found to have good hypergolicity with RFNA (Red Fuming Nitric Acid, good storability at room temperature and the mechanical properties. The paper presented the techniques of casting, ignition delay measurements and indicates the future programme for this study.

  7. AFRL Solid Propellant Laboratory Explosive Siting and Renovation Lessons Learned


    Siting and Renovation Lessons Learned 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Daniel F . Schwartz (AFRL/RZSP) 5d. PROJECT... Daniel F . Schwartz a. REPORT Unclassified b. ABSTRACT Unclassified c. THIS PAGE Unclassified SAR 35 19b. TELEPHONE NUMBER (include...1 AFRL Solid Propellant Laboratory Explosive Siting and Renovation Lessons Learned Daniel F . Schwartz Air Force Research Laboratory

  8. Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

    Kenny, Jeremy R.; Moser, Marlow D.; Hulka, James; Jones, Gregg


    Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the

  9. Interpenetrating Polymer Networks as Binders for Solid Composite Propellants

    S. Parthiban


    Full Text Available A new family of polymeric binders for solid composite propellants is proposed, based on two component interpenetrating polymer networks (IPNs. These networks comprise two different polyurethanes based on hydroxy terminated polybutadiene and ISRO polyol interpenetrated with two different vinyl polymers, viz poly methyl methacrylate and polystyrene. the networks synthesized by the simultaneous interpenetrating technique have been characterized for their properties, such as stress-strain, density, viscosity, thermal degradation, and heat of combustion. Phase morphologies have been determined using electron microscopy. Suitable explanations have been adduced to rationalize the properties of IPNs in terms of their structures and chain interactions. A study of the mechanical properties and burning rates of the ammonium perchlorate (AP-based solid propellant using the newly synthesised IPNs as binders, has been carried out. The results show that both mechanical strength and burning rate of solid propellants could be suitably modified by simply changing the nature and/or the ratio of the two interpenetrating polymer components.

  10. Applications for Solid Propellant Cool Gas Generator Technology

    van der List, M.; van Vliet, L. D.; Sanders, H. M.; Put, P. A. G.; Elst, J. W. E. C.


    In 2002 and 2003, Bradford Engineering B.V. conducted, in corporation with the Dutch research institute TNO Prins Maurits Laboratory (PML) a SME study for ESA-ESTEC for the identification of spaceflight applications and on-ground demonstration of Solid Propellant Cool Gas Generator (SPCGG) technology. This innovative technology has been developed by TNO-PML while Bradford Engineering also brought in its experience in spaceflight hardware development and manufacturing. The Solid Propellant Cool Gas Generator (SPCGG) technology allows for pure gas generation at ambient temperatures, as opposed to conventional solid propellant gas generators. This makes the SPCGG technology interesting for a wide range of terrestrial spaceflight applications. During the first part of the study, a variety of potential applications have been identified and three applications were selected for a more detailed quantitative study. In the third phase a ground demonstration was performed successfully for a cold gas propulsion system application. During the actual demonstration test, 10 cool gas generators were mounted and all operated successfully in sequence, demonstrating good repeatability of the produced amount of gas and pressure.

  11. Measurements of acoustic responses of gaseous propellant injectors. [for rocket combustion

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.


    Results are presented for an investigation intended to provide experimental data that can quantitatively describe the way in which various coaxial injector designs affect the stability of gaseous propellant rocket motors. The response factors of configurations that simulate the flow conditions in a gaseous-fuel injector element and a gaseous-oxidizer injector element are measured by using a modified impedance-tube technique and under cold-flow conditions simulating those observed in rocket motors with axial instability. The measured injector response factor data are presented and discussed. It is shown that there is reasonable agreement between the measured injector response factors and those predicted by the Feiler and Heidmann model (1967), and that the orifice length can be varied to shift the resonant frequency of the injector without any change in the magnitude of the response factor at resonance. A change in the injector open-area ratio is found to have a significant effect on the characteristics of the injector response factor.

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

    Palaszewski, Bryan; Zakany, James S.


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

  13. Contact performance analysis on the joint of thrust vector control system in a certain solid propellant rocket motor%某固体火箭发动机推力向量控制系统接头接触性能分析

    刘文芝; 戴美魁; 韦广梅; 赵永忠


    以某固体火箭发动机推力向量控制系统摆动接头的单珠承载试验模型为计算模型,采用摩擦接触问题的Lagrange乘子法与弹塑性耦合的有限元理论,计算分析了摆动接头阳球试件在不同强化层厚度下的接触应力、变形及破坏机理;为降低系统摆动力矩,同时考虑大尺寸球面的表面强化和加工工艺,提出了满足接触性能的阳球表面强化层为1~1.4mm的厚度要求.通过与单珠承载试验对比分析,检验了有限元建模及算法的合理性;确定了小变形条件下摆动接头弹塑性摩擦接触应力限额为5 384.9 MPa.%In this paper, a single ball load-bearing experiment model of the swing joint of thrust vector control system in a solid propellant rocket motor was used as computation model. By using the Lagrangian multiplier method for friction contact question, and elasto-plastic couplings finite element method, the contact stress, deformation and destruct principle under different thickness of strengthening layer of the test swing joint convex sphere were computed and analyzed. To reduce system swing moment, and consider manufacturing techniques of large size intensify sphere, 1-1.4 mm thickness of strengthening layer satisfied for contact performance was presented. Rationality of FEM modeling and algorithm was validated by single ball load-bearing experiment. The elasto-plastic friction contact stress of the swing joint was limited to 5384.9 Mpa in small deformation condition.

  14. Modeling of Nonlinear Combustion Instability in Solid Propellant Rocket Motors


    results of two solutions that tesi the ability of the LW+H+ACM to trect such problems are discussed. In both cases, the initial Pressure distur$.;nce...piezoelectric transducers. Thermal shielding of the pressure transducers was accomplished using vulcanized rubber (RTV). Based on shock i,,be calibration tests...potted with vulcanized rubber (RTV) which is utilized to provide thermal shielding 24 45. 16.0 � AFT CLOSURE WI0.45 250. for the pressure transducers

  15. A Study of Flame Physics and Solid Propellant Rocket Physics


    including suggestions for reducing the burden, to the Department of Defense, Executive Services and Communications Directorate 10704-0188...was re- placed with a 3-step kinetics model which gives improved agreement with experiental data for burning rates of a variety of heterogeneous

  16. Oscillatory burning of solid propellants including gas phase time lag.

    T'Ien, J. S.


    An analysis has been performed for oscillatory burning of solid propellants including gas phase time lag. The gaseous flame is assumed to be premixed and laminar with a one-step overall chemical reaction. The propellant is assumed to decompose according to the Arrenhius Law, with no condensed phase reaction. With this model, strong gas phase resonance has been found in certain cases at the characteristic gas-phase frequencies, but the peaking of the acoustic admittance is in the direction favoring the damping of pressure waves. At still higher frequencies, moderate wave-amplifying ability was found. The limit of low frequency response obtained previously by Denison and Baum was recovered, and the limitations of the quasi-steady theory were investigated.

  17. Oscillatory burning of solid propellants including gas phase time lag.

    T'Ien, J. S.


    An analysis has been performed for oscillatory burning of solid propellants including gas phase time lag. The gaseous flame is assumed to be premixed and laminar with a one-step overall chemical reaction. The propellant is assumed to decompose according to the Arrenhius Law, with no condensed phase reaction. With this model, strong gas phase resonance has been found in certain cases at the characteristic gas-phase frequencies, but the peaking of the acoustic admittance is in the direction favoring the damping of pressure waves. At still higher frequencies, moderate wave-amplifying ability was found. The limit of low frequency response obtained previously by Denison and Baum was recovered, and the limitations of the quasi-steady theory were investigated.

  18. A stop-restart solid propellant study with salt quench

    Kumar, R. N.


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

  19. Chemically Collapsible Mandrel for Solid Rocket Motor Processing

    Dey,Abhijit; Kumar, Arvind; Sikder,Arun K; Gupta, Manoj


    ABSTRACT: Composite propellant mainly consists of two parts, binder matrix (prepolymer, plasticizer, cross linker, antioxidant and curative etc.) and solid ingredients (oxidizer, metal fuel, burn rate modifier, combustion stabilizer etc.). Its processing involves several stages like ingredient preparation (grinding, 1.1 Hazard Division - 1.1 HD), mixing (1.1 HD), casting (1.1 HD), curing (1.3 HD) and extraction (1.3 HD). Each and every process is very hazardous. Removal of any of the mentione...

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

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


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

  1. Wind Tunnel Tests on Aerodynamic Characteristics of Advanced Solid Rocket

    Kitamura, Keiichi; Fujimoto, Keiichiro; Nonaka, Satoshi; Irikado, Tomoko; Fukuzoe, Moriyasu; Shima, Eiji

    The Advanced Solid Rocket is being developed by JAXA (Japan Aerospace Exploration Agency). Since its configuration has been changed very recently, its aerodynamic characteristics are of great interest of the JAXA Advanced Solid Rocket Team. In this study, we carried out wind tunnel tests on the aerodynamic characteristics of the present configuration for Mach 1.5. Six test cases were conducted with different body configurations, attack angles, and roll angles. A six component balance, oilflow visualization, Schlieren images were used throughout the experiments. It was found that, at zero angle-of-attack, the flow around the body were perturbed and its drag (axial force) characteristics were significantly influenced by protruding body components such as flanges, cable ducts, and attitude control units of SMSJ (Solid Motor Side Jet), while the nozzle had a minor role. With angle-of-attack of five degree, normal force of CNα = 3.50±0.03 was measured along with complex flow features observed in the full-component model; whereas no crossflow separations were induced around the no-protuberance model with CNα = 2.58±0.10. These values were almost constant with respect to the angle-of-attack in both of the cases. Furthermore, presence of roll angle made the flow more complicated, involving interactions of separation vortices. These data provide us with fundamental and important aerodynamic insights of the Advanced Solid Rocket, and they will be utilized as reference data for the corresponding numerical analysis.

  2. Effect of Experiment Environment on Calorimetric Value of Composite Solid Propellants

    Lalita S. Jawale


    Full Text Available The calorimetric value (cal-val of solid rocket propellants and explosives is determined in the presence of inert atmosphere using industrial nitrogen gas. However, due to presence of trace amount of oxygen, the cal-val is not always correct. To avoid such inaccuracy in cal-val, a systematic study has been carried out by taking different types of solid propellant samples having burning rate in the range of 5 mm/s − 30 mm/s at different pressures. The data obtained were acquired using industrial nitrogen, ultra high pure nitrogen (UHP-N2, ultra high pure argon (UHP-Ar, air and ultra high pure oxygen (UHP-O2. The data reveal that cal-val is highest in the case of UHP-O2 due to complete combustion while in the case of air and industrial nitrogen it is found to be substantially less. Moreover, the cal-val in the presence of UHP-N2 and UHP-Ar meets the standard value with reproducibility. The results, further, confirm that for authentic cal-val, the most suitable environment is UHP-N2/UHP-Ar.

  3. Effect of Experiment Environment on Calorimetric Value of Composite Solid Propellants

    Lalita S. Jawale


    Full Text Available The calorimetric value (cal-val of solid rocket propellants and explosives is determined in the presenceof inert atmosphere using industrial nitrogen gas. However, due to presence of trace amount of oxygen, the cal-val is not always correct. To avoid such inaccuracy in cal-val, a systematic study has been carried out by takingdifferent types of solid propellant samples having burning rate in the range of 5 mm/s − 30 mm/s at different pressures.The data obtained were acquired using industrial nitrogen, ultra high pure nitrogen (UHP-N2, ultra high pureargon (UHP-Ar, air and ultra high pure oxygen (UHP-O2. The data reveal that cal-val is highest in the case ofUHP-O2 due to complete combustion while in the case of air and industrial nitrogen it is found to be substantiallyless. Moreover, the cal-val in the presence of UHP-N2 and UHP-Ar meets the standard value with reproducibility.The results, further, confirm that for authentic cal-val, the most suitable environment is UHP-N2/UHP-Ar.Defence Science Journal, 2013, 63(5, pp.467-472, DOI:

  4. Quantity Distance for the Kennedy Space Center Vehicle Assembly Building for Solid Propellant Fueled Launchers

    Stover, Steven; Diebler, Corey; Frazier, Wayne


    The NASA KSC VAB was built to process Apollo launchers in the 1960's, and later adapted to process Space Shuttles. The VAB has served as a place to assemble solid rocket motors (5RM) and mate them to the vehicle's external fuel tank and Orbiter before rollout to the launch pad. As Space Shuttle is phased out, and new launchers are developed, the VAB may again be adapted to process these new launchers. Current launch vehicle designs call for continued and perhaps increased use of SRM segments; hence, the safe separation distances are in the process of being re-calculated. Cognizant NASA personnel and the solid rocket contractor have revisited the above VAB QD considerations and suggest that it may be revised to allow a greater number of motor segments within the VAB. This revision assumes that an inadvertent ignition of one SRM stack in its High Bay need not cause immediate and complete involvement of boosters that are part of a vehicle in adjacent High Bay. To support this assumption, NASA and contractor personnel proposed a strawman test approach for obtaining subscale data that may be used to develop phenomenological insight and to develop confidence in an analysis model for later use on full-scale situations. A team of subject matter experts in safety and siting of propellants and explosives were assembled to review the subscale test approach and provide options to NASA. Upon deliberations regarding the various options, the team arrived at some preliminary recommendations for NASA.

  5. [Combustion temperature measurement of solid propellant and the effect of organic compound on combustion temperature].

    Zhou, Xue-tie; Li, Yan; Chen, Zuo-ru; Wang, Jun-de


    The FTIR emission spectra in the spectral range of 4,500-300 cm-1 for the solid propellants were measured by a remote sensing FTIR system. The P-branch of fine structure of HCl fundamental band lying at 3.46 microns was used for precise combustion temperature measurement of the solid propellant. The effect of the organic compound in the solid propellant on the combustion temperature was discussed.

  6. Bonding agensi za kompozitna raketna goriva / Bonding agents for composite rocket propellants

    Mirjana Petrić


    Full Text Available U radu je analiziran uticaj bonding agensd na mehaničke karakteristike i viskozitet kompozitnih raketnih goriva na bazi hidroksiterminiranog polibutadiena i toluendiizocijanata, odnosno izoforondiizocijanata. Komparirana je efikasnost bonding agensd (BA različite strukture trietilentetramina (TET, tris-(2-metil aziridinil fosfin oksida (MAPO i N,N-di (2-hidroksietil-4,4-dimetilhidantoina (DHE. Efikasnost delovanja bonding agensa utvrdenaje na osnovu parametara dobijenih iz testa jednoosnog zatezanja uzoraka (σm - zatezna čvrstoća pri maksimalnoj sili, εm relativno izduženje pri maksimalnoj sili i εp - relativno izduženje pri prekidu. Mehaničke karakteristike merene su u temperaturskom području od -50°C do 50°C. Analiziran je uticaj BA na brzinu promene viskoziteta goriva. / The paper analyzes the effects of bonding agents on mechanical properties and viscosity of composite rocket propellants based on hydroxiterminated polybutadiene and toluendiizocyanate, i.e. izophorondiizocyanate. The efficiency of bonding agents (BA of different structure has been compared including triethylentetramine (TET, tris-(2-methyl azyrinidile phosphine oxide (MAPO and N, N-di (2-hydroxiethyl-4,4-dimethylhydantoine (DHE. The BA efficiency has been determined on the basis of parameters obtained by uniaxial tensile tests (σm - tensile strength at maximum force, εm - relative allongation at maximum force and εp - relative allongation at fracture. The mechanical properties have been measured at the temperature range from -50°C to 50°C. The BA effects on propellant viscosity change rates have been analyzed.

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

    Zimmerman, Chris J.; Litzinger, Gerald E.


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

  8. Evaluation of Solid Rocket Motor Component Data Using a Commercially Available Statistical Software Package

    Stefanski, Philip L.


    Commercially available software packages today allow users to quickly perform the routine evaluations of (1) descriptive statistics to numerically and graphically summarize both sample and population data, (2) inferential statistics that draws conclusions about a given population from samples taken of it, (3) probability determinations that can be used to generate estimates of reliability allowables, and finally (4) the setup of designed experiments and analysis of their data to identify significant material and process characteristics for application in both product manufacturing and performance enhancement. This paper presents examples of analysis and experimental design work that has been conducted using Statgraphics®(Registered Trademark) statistical software to obtain useful information with regard to solid rocket motor propellants and internal insulation material. Data were obtained from a number of programs (Shuttle, Constellation, and Space Launch System) and sources that include solid propellant burn rate strands, tensile specimens, sub-scale test motors, full-scale operational motors, rubber insulation specimens, and sub-scale rubber insulation analog samples. Besides facilitating the experimental design process to yield meaningful results, statistical software has demonstrated its ability to quickly perform complex data analyses and yield significant findings that might otherwise have gone unnoticed. One caveat to these successes is that useful results not only derive from the inherent power of the software package, but also from the skill and understanding of the data analyst.

  9. Experimental determination of the particle sizes in a subscale motor for application to the Ariane 5 solid rocket booster

    Traineau, J. C.; Kuentzmann, P.; Prevost, M.; Tarrin, P.; Delfour, A.

    The knowledge of the aluminum oxide particle size distribution inside the combustion chamber of a solid propellant rocket motor is an important factor for assessing the combustion stability or the slag mass accumulation in the motor. A representative subscale motor for the Ariane 5 P230 Solid Rocket Booster (SRB), in which helium is injected to quench the condensed phase reactions, has been designed and manufactured. Its use for combustion stability purpose has given the aluminum oxide particle size distribution in conditions representative of the actual Ariane 5 SRB. The experimental techniques, optical and particle capturing, have been found to give results in good agreement. A stretched distribution, with particles ranging from 1 micron to 120 microns and a maximum around 45 microns, has been demonstrated.

  10. Multisized Inert Particle Loading for Solid Rocket Axial Combustion Instability Suppression

    David R. Greatrix


    Full Text Available In the present investigation, various factors and trends, related to the usage of two or more sets of inert particles comprised of the same material (nominally aluminum but at different diameters for the suppression of axial shock wave development, are numerically predicted for a composite-propellant cylindrical-grain solid rocket motor. The limit pressure wave magnitudes at a later reference time in a given pulsed firing simulation run are collected for a series of runs at different particle sizes and loading distributions and mapped onto corresponding attenuation trend charts. The inert particles’ presence in the central core flow is demonstrated to be an effective means of instability symptom suppression, in correlating with past experimental successes in the usage of particles. However, the predicted results of this study suggest that one needs to be careful when selecting more than one size of particle for a given motor application.

  11. Internal Flow Analysis of Large L/D Solid Rocket Motors

    Laubacher, Brian A.


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

  12. Reusable Solid Rocket Motor - Accomplishments, Lessons, and a Culture of Success

    Moore, Dennis R.; Phelps, Willie J.


    The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solid motor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large

  13. A Review on Decomposition Deflagration of Oxidizer and Binders in Composite Solid Propellants

    K. Kishore


    Full Text Available Binder and oxidizer decomposition play very significant role during the combustion of composite solid propellants. Ammonium perchlorate (AP is the practical oxidizer in composite propellant formulations. Available information on binder decomposition in general and AP decomposition in particular have been collected and reviewed from the viewpoint of their application in propellants. This review may be useful in understanding the mechanism of propellant combustion.

  14. Rocket Propellant Ducts (Cryogenic Fuel Lines): First Cut Approximations and Design Guidance

    Brewer, William V.


    The design team has to set parameters before analysis can take place. Analysis is customarily a thorough and time consuming process which can take weeks or even months. Only when analysis is complete can the designer obtain feedback. If margins are negative, the process must be repeated to a greater or lesser degree until satisfactory results are achieved. Reduction of the number of iterations thru this loop would beneficially conserve time and resources. The task was to develop relatively simple, easy to use, guidelines and analytic tools that allow the designer to evaluate what effect various alternatives may have on performance as the design progresses. "Easy to use" is taken to mean closed form approximations and the use of graphic methods. "Simple" implies that 2-d and quasi 3-d approximations be exploited to whatever degree is useful before more resource intensive methods are applied. The objective is to avoid the grosser violation of performance margins at the outset. Initial efforts are focused on thermal expansion/contraction and rigid body kinematics as they relate to propellant duct displacements in the gimbal plane loop (GPL). The purpose of the loop is to place two flexible joints on the same two orthogonal intersecting axes as those of the rocket motor gimbals. This supposes the ducting will flex predictably with independent rotations corresponding to those of the motor gimbal actions. It can be shown that if GPL joint axes do not coincide with motor gimbal axes, displacement incompatibilities result in less predictable movement of the ducts.

  15. Experimental study of a valveless pulse detonation rocket engine using nontoxic hypergolic propellants

    Kan, Brandon K.

    A pulsed detonation rocket engine concept was explored through the use of hypergolic propellants in a fuel-centered pintle injector combustor. The combustor design yielded a simple open ended chamber with a pintle type injection element and pressure instrumentation. High-frequency pressure measurements from the first test series showed the presence of large pressure oscillations in excess of 2000 psia at frequencies between 400-600 hz during operation. High-speed video confirmed the high-frequency pulsed behavior and large amounts of after burning. Damaged hardware and instrumentation failure limited the amount of data gathered in the first test series, but the experiments met original test objectives of producing large over-pressures in an open chamber. A second test series proceeded by replacing hardware and instrumentation, and new data showed that pulsed events produced under expanded exhaust prior to pulsing, peak pressures around 8000 psi, and operating frequencies between 400-800 hz. Later hot-fires produced no pulsed behavior despite undamaged hardware. The research succeeded in producing pulsed combustion behavior using hypergolic fuels in a pintle injector setup and provided insights into design concepts that would assist future injector designs and experimental test setups.

  16. Thermal Barriers Developed for Solid Rocket Motor Nozzle Joints

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.


    Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn's braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block.

  17. Propellant Vaporization as a Criterion for Rocket-Engine Design; Experimental Performance, Vaporization and Heat-Transfer Rates with Various Propellant Combinations

    Clark, Bruce J.; Hersch, Martin; Priem, Richard J.


    Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.

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

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


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

  19. Solid propellant exhausted aluminum oxide and hydrogen chloride - Environmental considerations

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


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

  20. Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

    Messing, Wesley E.


    Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

  1. Quality assurance and control in the production and static tests of the solid rocket boosters for the Space Shuttle

    Cerny, O. F.


    The paper surveys the various aspects of design and overhaul of the solid rocket boosters. It is noted that quality control is an integral part of the design specifications. Attention is given to the production process which is optimized towards highest quality. Also discussed is the role of the DCA (Defense Contract Administration) in inspecting the products of subcontractors, noting that the USAF performs this role for prime contractors. Fabrication and construction of the booster is detailed with attention given to the lining of the booster cylinder and the mixing of the propellant and the subsequent X-ray inspection.

  2. Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical Analyzer


    P. N., Singh, P. P., and Bhattacharya, B., “Determination of Activation Energy of Relaxation Events in Composite Solid Propellants by Dynamic...Article 3. DATES COVERED (From - To) 04 August 2016 – 29 March 2017 4. TITLE AND SUBTITLE Measuring Propellant Stress Relaxation Modulus Using Dynamic...ERC 14. ABSTRACT A method for determining the stress relaxation master curve of solid rocket propellants was developed. The propellant was tested in

  3. Solid Hydrogen Experiments for Atomic Propellants: Image Analyses

    Palaszewski, Bryan


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

  4. Energy and Propulsion Optimization of Solid-Propellant Grain of a Hybrid Power Device

    Bondarchuk Sergey S.; Bondarchuk Iliya S.; Borisov Boris V.; Zhukov Alexandr S.


    A method of distribution of an additional solid-phase component (oxidizer) providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellant motor is proposed in the paper.

  5. Energy and Propulsion Optimization of Solid-Propellant Grain of a Hybrid Power Device

    Bondarchuk, Sergey S.; Bondarchuk, Iliya S.; Borisov, Boris V.; Zhukov, Alexandr S.


    A method of distribution of an additional solid-phase component (oxidizer) providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellant motor is proposed in the paper.

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

    Beckstead, M. W.


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

  7. Solid Rocket Fuel Constitutive Theory and Polymer Cure

    Ream, Robert


    Solid Rocket Fuel is a complex composite material for which no general constitutive theory, based on first principles, has been developed. One of the principles such a relation would depend on is the morphology of the binder. A theory of polymer curing is required to determine this morphology. During work on such a theory an algorithm was developed for counting the number of ways a polymer chain could assemble. The methods used to develop and check this algorithm led to an analytic solution to the problem. This solution is used in a probability distribution function which characterizes the morphology of the polymer.

  8. Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.


    Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

  9. Combustion Stability Assessments of the Black Brant Solid Rocket Motor

    Fischbach, Sean


    The Black Brant variation of the Standard Brant developed in the 1960's has been a workhorse motor of the NASA Sounding Rocket Project Office (SRPO) since the 1970's. In March 2012, the Black Brant Mk1 used on mission 36.277 experienced combustion instability during a flight at White Sands Missile Range, the third event in the last four years, the first occurring in November, 2009, the second in April 2010. After the 2010 event the program has been increasing the motor's throat diameter post-delivery with the goal of lowering the chamber pressure and increasing the margin against combustion instability. During the most recent combustion instability event, the vibrations exceeded the qualification levels for the Flight Termination System. The present study utilizes data generated from T-burner testing of multiple Black Brant propellants at the Naval Air Warfare Center at China Lake, to improve the combustion stability predictions for the Black Brant Mk1 and to generate new predictions for the Mk2. Three unique one dimensional (1-D) stability models were generated, representing distinct Black Brant flights, two of which experienced instabilities. The individual models allowed for comparison of stability characteristics between various nozzle configurations. A long standing "rule of thumb" states that increased stability margin is gained by increasing the throat diameter. In contradiction to this experience based rule, the analysis shows that little or no margin is gained from a larger throat diameter. The present analysis demonstrates competing effects resulting from an increased throat diameter accompanying a large response function. As is expected, more acoustic energy was expelled through the nozzle, but conversely more acoustic energy was generated due to larger gas velocities near the propellant surfaces.

  10. Study of plasticizer diffusion in a solid rocket motor´s bondline

    Juliano Libardi


    Full Text Available This work aims to determine the diffusion coefficient of the plasticizers dibutyl phthalate (DBP, dioctyl phthalate (DOP and dioctyl azelate (DOZ on the internal insulating layer of solid rocket motors. These plasticizers are originally present in the layers of rubber, liner and propellant, respectively. This species are not chemically bonded and tend to diffuse from propellant to insulating and vice versa. A computer program based on the mathematical model of Fick’s second Law of diffusion was developed to perform the calculus from the concentration data obtained by gas chromatographic (GC analyses. The samples were prepared with two different adhesive liners; one conventional (LHNA and the other with barrier properties (LHNT. A common feature of both liners was that they were synthesized by the reaction of hydroxyl-terminated polybutadiene (HTPB and diisocyanates. However, a bond promoter was used to increase the crosslink density of the LHNT liner and to improve its performance as barrier against the diffusion. The effects of the diffusion of the plasticizers were also investigated by hardness analyses, which were executed on samples aged at room temperature and at 80ºC. The results showed an increase trend for the samples aged at room temperature and an opposite behavior for the tests carried out at 80ºC.

  11. Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

    David Greatrix


    Full Text Available The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave development in the motor chamber flow. With the focus of the present study placed on reactive particles, a numerical internal ballistic model incorporating relevant elements, such as a transient, frequency-dependent combustion response to axial pressure wave activity above the burning propellant surface, is applied to the investigation of using aluminum particles within the central internal flow (particles whose surfaces nominally regress with time, as a function of current particle size, as they move downstream as a means of suppressing instability-related symptoms in a cylindrical-grain motor. The results of this investigation reveal that the loading percentage and starting size of the aluminum particles have a significant influence on reducing the resulting transient pressure wave magnitude.

  12. [Remote passive detection of flame temperature of solid propellant adulterating nanoparticles].

    Zhang, Li-ming; Zhang, Lin; Li, Yan; Liu, Bing-ping; Wang, Xiao-fei; Wang, Jun-de


    The flame temperature of three kinds of solid propellants was measured by passive remote sensing FTIR with the resolution of 1 cm(-1). These three kinds of solid propellants are adulterate nano-scale metal oxide particles, adulterate normal metal oxide particles, and propellant without any adulterations. The main components of the solid propellant are nitrocellulose and nitroglycerin. The metallic oxides, including 6 nm CuO, 56 nm Fe2O3, 16 nm NiO, and correspondingly the normal particles, were adulterated into the solid propellants respectively. The flame temperature was calculated through the fine structure of the emission fundamental band of H2O at 2.75 microm. The results of the flame temperature of the solid propellants adulterating nano-scale CuO, Fe2O3 and NiO are 3089, 3193 and 3183 K, respectively. The temperatures of the three kinds of solid propellants were compared, and it was shown that there is no obvious difference in the flame temperature among the three kinds of solid propellants.

  13. 固体火箭发动机预固化技术及其应用%Pre-cure Technique and Its Applications for Solid Rocket Motors

    苏昌银; 张爱科


    Based on interface properties of HTPB propellant, the cross linking level is regulated through the temperature and time of cure reaction. The chemical reaction takes place gradually in remainder functional groups of the system to form chemical bonds and hydrogen bond, so as to improve the mechanical properties of the resultant. Pre-cure technique and bonding model are described in the paper. They can be used in propellant-liner bonding, propellant loading and integral repairing of the propellant grain of the solid rocket motor (SRM). These test results have been qualified by the successful static firing tests, flight tests of motors and storage tests of specimens for ten-years. The performances of the motor meet the design requirements with good reproducibilities.

  14. Manganese oxalate nanorods as ballistic modifier for composite solid propellants

    Singh, Supriya [Department of Chemistry, DDU Gorakhpur University, Gorakhpur 273009, U.P. (India); Chawla, Mohit [School of Basic Sciences, Indian Institute of Technology Mandi, Mandi 175005, H.P. (India); Siril, Prem Felix, E-mail: [School of Basic Sciences, Indian Institute of Technology Mandi, Mandi 175005, H.P. (India); Singh, Gurdip [Department of Chemistry, DDU Gorakhpur University, Gorakhpur 273009, U.P. (India)


    Highlights: • Manganese oxalate nanorods were prepared using mild thermal precipitation and aging. • The nanorods were found to be efficient ballistic modifier for solid propellants. • The nanorods sensitized the thermolysis of ammonium perchlorate. • Controlled thermal decomposition of nanorods yielded manganese oxide nanoparticles. • MnO nanoparticles formed insitu in the condensed phase enhance the burning rates. - Abstract: Rod-shaped nanostructures of manganese oxalate (MnC{sub 2}O{sub 4}) were synthesized via mild thermal precipitation and aging process. Chemical composition of the MnC{sub 2}O{sub 4} nanorods was confirmed using Fourier transform infra-red (FTIR) spectroscopy and energy dispersive X-ray spectroscopy (EDS). X-ray diffraction (XRD) and selected area electron diffraction (SAED) studies revealed the crystal structure. Field emission scanning electron microscopy (FE-SEM) imaging and high resolution transmission electron microscopy (HR-TEM) were employed to study the structural features of the nanorods. The MnC{sub 2}O{sub 4} nanorods were found to be efficient ballistic modifier for the burning rate enhancement of composite solid propellants (CSPs). Thermal analysis using TGA-DSC showed that MnC{sub 2}O{sub 4} nanorods sensitized the thermal decomposition of ammonium perchlorate (AP) and the CSPs. Controlled thermal decomposition of the MnC{sub 2}O{sub 4} nanorods resulted in the formation of managanese oxide nanoparticles with mesoporosity. A plausible mechanism for the burning rate enhancement using MnC{sub 2}O{sub 4} nanorods was proposed.

  15. Launch Vehicle Performance with Solid Particle Feed Systems for Atomic Propellants

    Palaszewski, Bryan


    An analysis of launch vehicle Gross Liftoff Weight (GLOW) using high energy density atomic propellants with solid particle feed systems was conducted. The analyses covered several propellant combinations, including atoms of aluminum (Al), boron (B). carbon (C), and hydrogen (H) stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents (wt%) for the liquid carrier were investigated and the gross lift off weight (GLOW) of the vehicles using the solid particle feed systems were compared with a conventional 02/H2 propellant vehicle. The potential benefits and effects of feed systems using solid particles in a liquid cryogenic fluid are discussed.

  16. Environmental impact evaluation of static tests of solid propellant propellers; Evaluation de l`impact sur l`environnement des essais statiques de propulseurs a propergol solide

    Aguesse, T.; Moreau, S. [DGA, Direction des Missiles et de l`Espace, Centre d`Achevement et d`Essais des Propulseurs et Engins, 33 - Saint Medard en Jalles (France)


    The CAEPE, the French Centre of Propellers and Engines Completion and Testing is in charge of the static tests of solid propellant fuelled propellers. In order to determine the schedule of firing permissions, predictive means are used to predict the environmental impact of propellers firing. Calculation and simulation codes are used to build maps of acoustic nuisance and acid fallout. These codes, which use in-situ meteorological radio sounding data, were progressively adjusted during testing of engines with different sizes (up to the Ariane 5 P230 auxiliary propeller). In this presentation, the authors focus on a model derived from the G.A. Briggs` model for the ascension of high temperature effluents. The correct simulation of acid rains requires a good description of this phenomenon. Concerning the other aspects of the codes, the main references are given. (J.S.) 15 refs.

  17. Materials characterization of propellants using ultrasonics

    Workman, Gary L.; Jones, David


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

  18. The application of near-infrared spectroscopy for the quality control analysis of rocket propellant fuel pre-mixes.

    Judge, Michael D


    The viability of near-infrared (NIR) spectroscopy as a technique for the quality control analysis of ingredient concentrations in a rocket propellant fuel liquid pre-mix was investigated. The pre-mix analyzed consisted of a polybutadiene pre-polymer, a plasticizer and two antioxidants. It was determined that NIR spectroscopy offered a fast and convenient method of verifying the percentage level of all four ingredients while requiring no sample preparation. The NIR methodology exhibited a high level of accuracy and precision. There was also a clear indication that the technique allowed monitoring of antioxidant depletion in the pre-mix on ageing.

  19. Solid rocket booster internal flow analysis by highly accurate adaptive computational methods

    Huang, C. Y.; Tworzydlo, W.; Oden, J. T.; Bass, J. M.; Cullen, C.; Vadaketh, S.


    The primary objective of this project was to develop an adaptive finite element flow solver for simulating internal flows in the solid rocket booster. Described here is a unique flow simulator code for analyzing highly complex flow phenomena in the solid rocket booster. New methodologies and features incorporated into this analysis tool are described.

  20. Exploring the Solid Rocket Boosters and Properties of Matter

    Moffett, Amy


    I worked for the United Space Alliance, LLC (USA) with the Solid Rocket Booster (SRB) Materials and Process engineers (M&P). I was assigned a project in which I needed to research and collect chemical and physical properties information, material safety data sheets (MSDS), and other product information from the vendor's websites and existing "inhouse" files for a select group of materials used in building and refurbishing the SRBs. This information was then compiled in a report that summarized the information collected. My work site was at the Kennedy Space Center (KSC). This allowed for many opportunities to visit and tour sites operated by NASA, by USA, and by the Air Force. This included the vehicle assembly building (VAB), orbital processing facilities (OPF), the crawler with the mobile launch pad (MLP), and the SRB assembly and refurbishment facility (ARF), to name a few. In addition, the launch, of STS- 117 took place within the first week of employment allowing a day by day following of that mission including post flight operations for the SRBs. Two Delta II rockets were also launched during these 7 weeks. The sights were incredible and the operations witnessed were amazing. I learned so many things I never knew about the entire program and the shuttle itself. The entire experience, especially my work with the SRB materials, inspired my plan for implementation into the classroom.

  1. Explicit expression to predict the erosive burning rate of solid propellants

    Wang, S.


    Using the theory of gas dynamics and heat transfer from a turbulent gas flow to the burning surface of propellant along a permeable wall, an explicit expression is derived to predict the burning rate of the solid propellant with crossflow. Results of the calculation have been compared with experimental data and proved to be correct.

  2. Cathodic Protection Deployment on Space Shuttle Solid Rocket Boosters

    Zook, Lee M.


    Corrosion protection of the space shuttle solid rocket boosters incorporates the use of cathodic protection(anodes) in concert with several coatings systems. The SRB design has large carbon/carbon composites(motor nozzle) electrically connected to an aluminum alloy structure. Early in the STS program, the aluminum structures incurred tremendous corrosive attack due primarily to the galvanic couple to the carbon/carbon nozzle at coating damage locations. Also contributing to the galvanic corrosion problem were stainless steel and titanium alloy components housed within the aluminum structures and electrically connected to the aluminum structures. This paper will highlight the evolution in the protection of the aluminum structures, providing historical information and summary data from the operation of the corrosion protection systems. Also, data and information will be included regarding the evaluation and deployment of inorganic zinc rich primers as anode area on the aluminum structures.

  3. Ecological effects and environmental fate of solid rocket exhaust

    Nimmo, B.; Stout, I. J.; Mickus, J.; Vickers, D.; Madsen, B.


    Specific target processes were classified as to the chemical, chemical-physical, and biological reactions and toxic effects of solid rocket emissions within selected ecosystems at Kennedy Space Center. Exposure of Citris seedlings, English peas, and bush beans to SRM exhaust under laboratory conditions demonstrated reduced growth rates, but at very high concentrations. Field studies of natural plant populations in three diverse ecosystems failed to reveal any structural damage at the concentration levels tested. Background information on elemental composition of selected woody plants from two terrestrial ecosystems is reported. LD sub 50 for a native mouse (peromysous gossypinus) exposed to SRM exhaust was determined to be 50 ppm/g body weight. Results strongly indicate that other components of the SRM exhaust act synergically to enhance the toxic effects of HCl gas when inhaled. A brief summary is given regarding the work on SRM exhaust and its possible impact on hatchability of incubating bird eggs.

  4. Preliminary design and optimization of slotted tube grain for solid rocket motor


    In this paper,design and optimization technique of slotted tube grain for solid rocket motors has been discussed.In doing so,the design objectives and constraints have been set,geometric parameters identified,performance prediction parameters calculated,thereafter preliminary designs completed and finally optimal design reached.Geometric model for slotted tube grain configuration has been developed.Average thrust has been taken as the objective function with constraints of burning time,mass of propellant,fixed length and diameter of chamber case.Lumped parameter method has been used for calculating the performance prediction parameters.A set of preliminary designs has been completed and an analysis of these results conducted.Although all the preliminary results fulfill the design requirements in terms of objective function and constraints,however in order tO attain the optimal design,Sequen-tial quadratic programming optimization technique has been adopted.As the slotted tube grain ge-ometry is totally dependent upon various independent variables and each of these variables has a bearing on explicit characteristic of grain designing,hence affects of the independent variables on performance parameters have been examined,thus variation laws have been developed.Basing on the variation laws and the analysis of preliminary design results,upper and lower limits have been defined for the independent geometric variables and an initial guess provided for conducting optimi-zation.Resuhs attained exhibits that an optimal result has been attained and the value of objective function has been maximized.All the design constraint limits have also been met while ensuring sound values of volumetric loading fraction,web fraction and neutrality.This methodology of design and optimization of slotted tube grain for solid rocket motors can be used by engineers as a reference guide for actual design and engineering purposes.

  5. Indirect and direct methods for measuring a dynamic throat diameter in a solid rocket motor

    Colbaugh, Lauren

    In a solid rocket motor, nozzle throat erosion is dictated by propellant composition, throat material properties, and operating conditions. Throat erosion has a significant effect on motor performance, so it must be accurately characterized to produce a good motor design. In order to correlate throat erosion rate to other parameters, it is first necessary to know what the throat diameter is throughout a motor burn. Thus, an indirect method and a direct method for determining throat diameter in a solid rocket motor are investigated in this thesis. The indirect method looks at the use of pressure and thrust data to solve for throat diameter as a function of time. The indirect method's proof of concept was shown by the good agreement between the ballistics model and the test data from a static motor firing. The ballistics model was within 10% of all measured and calculated performance parameters (e.g. average pressure, specific impulse, maximum thrust, etc.) for tests with throat erosion and within 6% of all measured and calculated performance parameters for tests without throat erosion. The direct method involves the use of x-rays to directly observe a simulated nozzle throat erode in a dynamic environment; this is achieved with a dynamic calibration standard. An image processing algorithm is developed for extracting the diameter dimensions from the x-ray intensity digital images. Static and dynamic tests were conducted. The measured diameter was compared to the known diameter in the calibration standard. All dynamic test results were within +6% / -7% of the actual diameter. Part of the edge detection method consists of dividing the entire x-ray image by an average pixel value, calculated from a set of pixels in the x-ray image. It was found that the accuracy of the edge detection method depends upon the selection of the average pixel value area and subsequently the average pixel value. An average pixel value sensitivity analysis is presented. Both the indirect

  6. Advanced Multi-Phase Flow CFD Model Development for Solid Rocket Motor Flowfield Analysis

    Liaw, Paul; Chen, Y. S.; Shang, H. M.; Doran, Denise


    It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better

  7. The History of Solid-Propellant Rocketry: What We Do and Do Not Know

    Hunley, J. D.


    Contributions to the evolution of solid-propellant rocketry have come from a variety of sources. World War II research on large solids enabled one company to capitalize on work in the area of castable double-base propellants. Separate development of castable composite propellants led to production of Polaris and Minuteman powerplants. Pivotal to the development of these missiles were Edward Hall's advocacy of the Minuteman missile within the Air Force and contract funding to resolve problems. The discovery that adding large amounts of aluminum significantly increased the specific impulse of a castable composite propellant further aided large-missile technology. These separate lines of research led to the development of large solid-propellant motors and boosters. Many more discoveries went into the development of large solid-propellant motors. Ammonium perchlorate replaced potassium perchlorate as an oxidizer in the late 1940's, and binders were developed. Discoveries important in the evolution of large solid-propellant motors appear to have resulted from innovators' education and skills, an exposure to contemporary problems, an awareness of theory but a willingness not to let it dictate empirical investigations, and proper empirical techniques. Other important contributions are the adequate funding and exchange of information. However, many questions remain about these and other innovations.

  8. Numerical and experimental study of liquid breakup process in solid rocket motor nozzle

    Yen, Yi-Hsin

    Rocket propulsion is an important travel method for space exploration and national defense, rockets needs to be able to withstand wide range of operation environment and also stable and precise enough to carry sophisticated payload into orbit, those engineering requirement makes rocket becomes one of the state of the art industry. The rocket family have been classified into two major group of liquid and solid rocket based on the fuel phase of liquid or solid state. The solid rocket has the advantages of simple working mechanism, less maintenance and preparing procedure and higher storage safety, those characters of solid rocket make it becomes popular in aerospace industry. Aluminum based propellant is widely used in solid rocket motor (SRM) industry due to its avalibility, combusion performance and economical fuel option, however after aluminum react with oxidant of amonimum perchrate (AP), it will generate liquid phase alumina (Al2O3) as product in high temperature (2,700˜3,000 K) combustion chamber enviornment. The liquid phase alumina particles aggromorate inside combustion chamber into larger particle which becomes major erosion calprit on inner nozzle wall while alumina aggromorates impinge on the nozzle wall surface. The erosion mechanism result nozzle throat material removal, increase the performance optimized throat diameter and reduce nozzle exit to throat area ratio which leads to the reduction of exhaust gas velocity, Mach number and lower the propulsion thrust force. The approach to avoid particle erosion phenomenon taking place in SRM's nozzle is to reduce the alumina particle size inside combustion chamber which could be done by further breakup of the alumina droplet size in SRM's combustion chamber. The study of liquid breakup mechanism is an important means to smaller combustion chamber alumina droplet size and mitigate the erosion tack place on rocket nozzle region. In this study, a straight two phase air-water flow channel experiment is set up


    张泰华; 白以龙; 王世英; 刘培德


    In order to improve the safety of high-energy solid propellants, a study is carried out for the effects of damage on the combustion of the NEPE (Nitrate Ester Plasticized Polyether) propellant. The study includes: (1) to introduce damage into the propellants by means of a large-scale drop-weight apparatus; (2) to observe microstructural variations of the propellant with a scanning electron microscope (SEM) and then to characterize the damage with density measurements; (3) to investigate thermal decomposition; (4) to carry out closed-bomb tests. The NEPE propellant can be considered as a viscoelastic material. The matrices of damaged samples are severely degraded, but the particles are not. The results of the thermal decomposition and closed-bomb tests show that the microstructural damage in the propellant affects its decomposition and burn rate.

  10. Rocket

    K. Karmarkar


    Full Text Available The rockets of World War II represented, not the invention of a new weapon, but the modernization of a very old one. As early as 1232 A.D, the Chinese launched rockets against the Mongols. About a hundred years later the knowledge of ledge of rockets was quite widespread and they were used to set fire to buildings and to terrorize the enemy. But as cannon developed, rockets declined in warfare. However rockets were used occasionally as weapons till about 1530 A.D. About this time improvements in artillery-rifled gun barrel and mechanism to absorb recoil-established a standard of efficiency with which rockets could not compare until World War II brought pew conditions

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

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


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

  12. A Rate-Dependent Viscoelastic Damage Model for Simulation of Solid Propellant Impacts

    Matheson, E. R.; Nguyen, D. Q.


    A viscoelastic deformation and damage model (VED) for solid rocket propellants has been developed based on an extensive set of mechanical properties experiments. Monotonic tensile tests performed at several strain rates showed rate and dilatation effects. During cyclic tensile tests, hysteresis and a rate-dependent shear modulus were observed. A tensile relaxation experiment showed significant stress decay in the sample. Taylor impact tests exhibited large dilatations without significant crack growth. Extensive modifications to a viscoelastic-viscoplastic model (VEP) necessary to capture these experimental results have led to development of the VED model. In particular, plasticity has been eliminated in the model, and the multiple Maxwell viscoelastic formulation has been replaced with a time-dependent shear modulus. Furthermore, the loading and unloading behaviors of the material are modeled independently. To characterize the damage and dilatation behavior, the Tensile Damage and Distention (TDD) model is run in conjunction with VED. The VED model is connected to a single-cell driver as well as to the CTH shock physics code. Simulations of tests show good comparisons with tensile tests and some aspects of the Taylor tests.

  13. DURACON - Variable Emissivity Broadband Coatings for Liquid Propellant Rocket Nozzles Project

    National Aeronautics and Space Administration — The need exists for a fast drying, robust, low gloss, black, high emissivity coating that can be applied easily on aircraft rocket nozzles and nozzle extensions....

  14. Proposal for a Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft Project

    National Aeronautics and Space Administration — A new technology, the Fission Fragment Rocket Engine (FFRE), requires small amounts of readily available, energy dense, long lasting fuel, significant thrust at...

  15. Analytical model of the combustion of multicomponent solid propellants

    Cohen, N. S.; Price, C. F.; Strand, L. D.


    Multiple flame models derived for simple composite propellants are extended to describe the combustion of propellants containing multimodal particle sizes, mixed oxidizers and monopropellant binders. Models combining the component contributions to propellant surface structure, flame structure and energy distribution are based in part upon experimental observations and in part upon hypotheses constrained to provide reasonable agreement with measured burning rate characteristics. The methods employed consist of superposition, interaction and iteration. The computerized model is applied to explain the effects of multiple ingredients and to discuss burning rate tailoring problems of current interest.

  16. Investigation of Propellant and Explosive Solid Solution Systems II X-Ray Studies


    A\\Yj* ^\\C/*^ ^ 1 tatf AD 7t ott w AD-E400 125 TECHNICAL REPORT ARLCD-TR-77066 INVESTIGATION OF PROPELLANT AND EXPLOSIVE SOLID SOLUTION SYSTEMS...Report ARLCD-TR-77066 2. GOVT ACCESSION NO. *. TITLE (and Subtitle) INVESTIGATION OF PROPELLANT AND EXPLOSIVE SOLID SOLUTION SYSTEMS II X-RAY...Interplanar spacings and x-ray diffraction 9 intensities of AP, KP and their physical mixtures and solid solutions 4 X-ray data of 3 AN: KP solid solution and

  17. High-speed schlieren imaging of rocket exhaust plumes

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


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

  18. Solid Rocket Booster (SRB) Flight System Integration at Its Best

    Wood, T. David; Kanner, Howard S.; Freeland, Donna M.; Olson, Derek T.


    The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads, environments and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. However, the in-flight data and postflight assessment process revealed the hardware was affected much more strongly than originally anticipated. Assembly and integration of the booster subsystems required acceptance testing of reused hardware components for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed

  19. Viscoelastic Behaviour of Solid Propellants based on Various Polymeric Binders

    N. Prabhakaran


    Full Text Available The dynamic mechanical properties of different binders and corresponding propellants are studied in terms of storage modulus and loss tangent. The binders investigated are HTPB, CTPB, PBAN, HEF-20 and ISRO polyol. The viscoelastic behaviour is investigated using Rheovibron viscoelastometer at 35 Hz covering a wide temperature range (-100 degree centigrade to 100 degree centigrade. The properties of the binder and corresponding propellant are compared in terms of parameters, tan delta/sub max/, T/sub g/ and the trend of their master relaxation modulus curves. It is found that polybutadiene binders exhibit lowest T/sub g/ (around -60 degree centigrade and ISRO polyol the highest (near -20 degree centigrade. The propellants have higher moduli than the binders at any temperature. The master relaxation modulus curve is influenced by the type of propellant.

  20. A Multiconstrained Ascent Guidance Method for Solid Rocket-Powered Launch Vehicles

    Si-Yuan Chen


    Full Text Available This study proposes a multiconstrained ascent guidance method for a solid rocket-powered launch vehicle, which uses a hypersonic glide vehicle (HGV as payload and shuts off by fuel exhaustion. First, pseudospectral method is used to analyze the two-stage launch vehicle ascent trajectory with different rocket ignition modes. Then, constraints, such as terminal height, velocity, flight path angle, and angle of attack, are converted into the constraints within height-time profile according to the second-stage rocket flight characteristics. The closed-loop guidance method is inferred by different spline curves given the different terminal constraints. Afterwards, a thrust bias energy management strategy is proposed to waste the excess energy of the solid rocket. Finally, the proposed method is verified through nominal and dispersion simulations. The simulation results show excellent applicability and robustness of this method, which can provide a valuable reference for the ascent guidance of solid rocket-powered launch vehicles.

  1. Design and optimization of solid rocket motor Finocyl grain using simulated annealing

    Ali Kamran; LIANG Guo-zhu


    The research effort outlined the application of a computer aided design (CAD)-centric technique to the design and optimization of solid rocket motor Finocyl (fin in cylinder) grain using simulated annealing.The proper method for constructing the grain configuration model, ballistic performance and optimizer integration for analysis was presented. Finoeyl is a complex grain configuration, requiring thirteen variables to define the geometry. The large number of variables not only complicates the geometrical construction but also optimization process. CAD representation encapsulates all of the geometric entities pertinent to the grain design in a parametric way, allowing manipulation of grain entity (web), performing regression and automating geometrical data calculations. Robustness to avoid local minima and efficient capacity to explore design space makes simulated annealing an attractive choice as optimizer. It is demonstrated with a constrained optimization of Finocyl grain geometry for homogeneous, isotropic propellant, uniform regression, and a quasi-steady, bulk mode internal ballistics model that maximizes average thrust for required deviations from neutrality.

  2. Analysis of pressure blips in aft-finocyl solid rocket motor

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


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

  3. Studies on Decomposition and Combustion Mechanism of Solid Fuel Rich Propellants


    energetic plasticizers in place of conventional inert binders and inert plasticizers like organic phthalates and acetates . FRP formulations based on poly...9), 1994, pp. 269-278. 16) Singh, H; Shekhar, H. Science and Technology of rocket propellants, Printed by Print well, Tow -Chowk, Darbhanga, (India...diazido 3,3 nitroazo pentane) -- -80 554 -- Diethylene glycol bis- azido acetate 1 -100 -392 -- Trimethylol nitro methane 1.4 -72 -230

  4. Solid Rocket Booster Hydraulic Pump Port Cap Joint Load Testing

    Gamwell, W. R.; Murphy, N. C.


    The solid rocket booster uses hydraulic pumps fabricated from cast C355 aluminum alloy, with 17-4 PH stainless steel pump port caps. Corrosion-resistant steel, MS51830 CA204L self-locking screw thread inserts are installed into C355 pump housings, with A286 stainless steel fasteners installed into the insert to secure the pump port cap to the housing. In the past, pump port cap fasteners were installed to a torque of 33 Nm (300 in-lb). However, the structural analyses used a significantly higher nut factor than indicated during tests conducted by Boeing Space Systems. When the torque values were reassessed using Boeing's nut factor, the fastener preload had a factor of safety of less than 1, with potential for overloading the joint. This paper describes how behavior was determined for a preloaded joint with a steel bolt threaded into steel inserts in aluminum parts. Finite element models were compared with test results. For all initial bolt preloads, bolt loads increased as external applied loads increased. For higher initial bolt preloads, less load was transferred into the bolt, due to external applied loading. Lower torque limits were established for pump port cap fasteners and additional limits were placed on insert axial deformation under operating conditions after seating the insert with an initial preload.

  5. NDE of Space Shuttle Solid Rocket Motor field joint

    Johnston, Patrick H.

    One of the most critical areas for inspection in the Space Shuttle Solid Rocket Motors is the bond between the steel case and rubber insulation in the region of the field joints. The tang-and-clevis geometry of the field joints is sufficiently complex to prohibit the use of resonance-based techniques. One approach we are investigating is to interrogate the steel-insulation bondline in the tang and clevis regions using surface-travelling waves. A low-frequency contact surface wave transmitting array transducer is under development at our laboratory for this purpose. The array is placed in acoustic contact with the steel and surface waves are launched on the inside surface or the clevis leg which propagate along the steel-insulation interface. As these surface waves propagate along the bonded surface, the magnitude of the ultrasonic energy leaking into the steel is monitored on the outer surface of the case. Our working hypothesis is that the magnitude of energy received at the outer surface of the case is dependent upon the integrity of the case-insulation bond, with less attenuation for propagation along a disbond due to imperfect acoustic coupling between the steel and rubber. Measurements on test specimens indicate a linear relationship between received signal amplitude and the length of good bend between the transmitter and receiver, suggesting the validity of this working hypothesis.

  6. Filled Ethylene-propylene Diene Terpolymer Elastomer as ThermalInsulator for Case-bonded Solid Rocket Motors

    C. M. Bhuvaneswari


    Full Text Available Ethylene-propylene diene terpolymer (EPDM-based insulation system is being globallyused for case-bonded solid rocket motors. A study was undertaken using EPDM as base polymer,blended with hypalon and liquid EPDM and filled with fibrous and non-fibrous fillers. Theseformulations were evaluated as rocket motor insulation system. The basic objective of the studywas to develop an insulation system based on EPDM for case-bonded applications. A series ofrocket motor insulator compositions based on EPDM, filled with particulate and fibrous fillerslike precipitated silica, fumed silica, aramid, and carbon fibres have been studied for mechanical,rheological, thermal, and interface properties. Compositions based on particulate fillers wereoptimised for the filler content. Comparatively, fumed silica was found to be superior as fillerin terms of mechanical and interface properties. Addition of fibrous filler (5 parts improved thepeel strength, and reduced the thermal conductivity and erosion rate. All the compositions wereevaluated for sulphur and peroxide curing. Superior mechanical properties were achieved forsulphur-cured products, whereas peroxide-cured products exhibited an excellent ageing resistance.Rocket motors were insulated with optimised composition and propellant cast, and the motorswere evaluated by conducting static test in end-burning mode.Defence Science Journal, 2008, 58(1, pp.94-102, DOI :

  7. Evaluation of Energetic Plasticisers for Solid Gun Propellant

    R.S. Damse


    Full Text Available This paper reports the evaluation of four different energetic plasticisers, viz., glycidyl azidepolymer (GAP, MW = 390, 1,5-diazido-3-nitrazapentane (DANPE, ethylene-glycol-bis-azido-acetate (EGBAA and N-n-butyl-N-(2 nitroxyethyl nitramine (n-Bu-NENA separately into highenergy gun propellant containing 28 per cent NC (13.1 N %, 65 per cent RDX, 6 per cent di-octyl-phthahate (DOP and 1 per cent carbamite.  Four different propellant compositions based on theenergetic plasticiser have been formulated separately with the replacement of non-energeticplasticiser, DOP. The propellants were processed by standard solvent method and evaluatedexperimentally along with the control composition to determine the ballistic parameters, cal-val,sensitivity, thermal characterisation, thermal stability and mechanical properties. The performanceof the propellants containing the energetic plasticiser has  been compared with that of thecontrol composition containing the non-energetic plasticiser, DOP so as to assess the suitabilityof the energetic plasticiser for the futuristic gun propellant formulations. It has been found outthat n-Bu-NENA is the superior plasticiser among the four energetic plasticisers evaluated inthis study.Defence Science Journal, 2008, 58(1, pp.86-93, DOI:

  8. Maturation of Structural Health Management Systems for Solid Rocket Motors Project

    National Aeronautics and Space Administration — Solid rocket motor cases are subject to a variety of external environmental and loading conditions from cradle-to-grave. These conditions can significantly impact...

  9. Polarographic study on the evolution of the diphenylamine as stabiliser of the solid propellants.

    de Diego Martínez, A Pérez; Tascón, M L; Vázquez, M D; Sánchez Batanero, P


    Differential pulse polarography (DPP) and square wave voltammetry (SWV) were investigated, in order to know the stability of solid propellants which contain diphenylamine. The simultaneous determination of N-nitrousdiphenylamine (NnDPA), 2-nitrodiphenylamine (2nDPA), 4-nitrodiphenylamine (4nDPA) and 2,4-dinitrodiphenylamine (2,4dnDPA) by DPP and SWV was proposed due to these nitro-derivatives appear during the stabilisation process from degradation of diphenylamine (DPA) used as stabiliser in propellant compositions. The proposed methods were successfully applied to the simple base solid propellant (with nitrocellulose as the only active component), with its stabiliser DPA. In all cases and with both the techniques, detection limits propellants accurately, faster and cheaper than LC methods.

  10. Hybrid Rocket Technology

    Sankaran Venugopal


    Full Text Available With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nontoxic, and not hazardous to store and transport. Solid fuel grains are not highly susceptible to cracks, imperfections, and environmental temperature and are therefore safer to manufacture, store, transport, and use for launch. The status of development based on the experience of the last few decades indicating the maturity of the hybrid rocket technology is given in brief.Defence Science Journal, 2011, 61(3, pp.193-200, DOI:

  11. Energy and Propulsion Optimization of Solid-Propellant Grain of a Hybrid Power Device

    Bondarchuk Sergey S.


    Full Text Available A method of distribution of an additional solid-phase component (oxidizer providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellant motor is proposed in the paper.

  12. Detailed analysis of a quench bomb for the study of aluminum agglomeration in solid propellants

    Gallier, S.; Kratz, J.-G.; Quaglia, N.; Fouin, G.


    A standard quench bomb (QB) - widely used to characterize condensed phase from metalized solid propellant combustion - is studied in detail. Experimental and numerical investigations proved that collected particles are mostly unburned aluminum (Al) agglomerates despite large quenching distances. Particles are actually found to quench early as propellant surface is swept by inert pressurant. Further improvements of the QB are proposed which allow measuring both Al agglomerates and alumina residue with the same setup. Finally, the results obtained on a typical aluminized ammonium perchlorate (AP) / hydroxyl-terminated polybutadiene (HTPB) propellant are briefly discussed.

  13. Application of Ultrasonic Technique for Measurement of Instantaneous Burn Rate of Solid Propellants .

    Desh Deepak


    Full Text Available The ultrasonic pulse-echo technique has been applied for the measurement of instantaneous burnrate of aluminised composite solid propellants. The tests have been carried out on end-burning 30 mmthick propellant specimens at nearly constant pressure of about 1.9 MPa. Necessary software forpost-test data processing and instantaneous burn rate computations have been developed. The burnrates measured by the ultrasonic technique have been compared with those obtained from ballisticevaluation motor tests on propellant from the same mix. An accuracy of about +- 1 per cent ininstantaneous burn rate measurements and reproducibility of results have been demonstrated byapplying ultrasonic technique.

  14. Studies on the Pyrolysis of Composite Solid Propellants leading to Ignition

    K. Kishore


    Full Text Available Studies on the pyrolysis leading to ignition of polystyrene (PS/ammonium perchlorate (AP, polyvinyl chloride (PVC/AP and polyphenol formaldehyde (PPF/AP  propellants revealed that the activation energy for the ignition strongly depends upon the binder. Double base propellants exhibit an empirical relationship between the ignition delay (~ and the oxidiser concentration; when the same equation is applied to composite solid propellants, the plot of In ~ vs In C (C=Concentration, % yields a straight line with a knee corresponding to 65-70 per cent AP above which the dependence on ~ becomes less sensitive.

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

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


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

  16. Storage of solid propellants in a dry environment. [mechanical properties resulting from long term exposure to aerospace environments

    Udlock, D. E.


    Storage of solid propellants in either a dry or a vacuum environment causes a significantly greater increase in the propellants' modulus and maximum tensile strength than does ambient storage. It is postulated that these physical property changes can be attributed to the effect trace amount of moisture has on the bond between the propellants' binder and oxidizer.

  17. Propellants

    Lt. Col. B. N Mitra


    Full Text Available There is little doubt that explosives had their origin in warfare. In the armed conflict between groups of individuals or of states, where each sought and still seeks to impose its will upon the other by force, it was inevitable that arms should grow and flourish. The sling, the bow and arrow, the sword and firearm typify evolution in warfare weapons. Bs a means of propelling missiles, the gun and gun powder were thought of. The history of explosives, therefore, may be said to begin with black powder.

  18. Electrets used in measuring rocket exhaust effluents from the space shuttle's solid rocket booster during static test firing, DM-3

    Susko, M.


    The purpose of this experimental research was to compare Marshall Space Flight Center's electrets with Thiokol's fixed flow air samplers during the Space Shuttle Solid Rocket Booster Demonstration Model-3 static test firing on October 19, 1978. The measurement of rocket exhaust effluents by Thiokol's samplers and MSFC's electrets indicated that the firing of the Solid Rocket Booster had no significant effect on the quality of the air sampled. The highest measurement by Thiokol's samplers was obtained at Plant 3 (site 11) approximately 8 km at a 113 degree heading from the static test stand. At sites 11, 12, and 5, Thiokol's fixed flow air samplers measured 0.0048, 0.00016, and 0.00012 mg/m3 of CI. Alongside the fixed flow measurements, the electret counts from X-ray spectroscopy were 685, 894, and 719 counts. After background corrections, the counts were 334, 543, and 368, or an average of 415 counts. An additional electred, E20, which was the only measurement device at a site approximately 20 km northeast from the test site where no power was available, obtained 901 counts. After background correction, the count was 550. Again this data indicate there was no measurement of significant rocket exhaust effluents at the test site.

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

    Styborski, Jeremy A.

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

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


    quanti- tative statistical analysis would be of tremendous help to the modelers . This type of Information is available in high-speed movies of flows...fivJ,’t(,rriiilation to be completely consumed, movies sholld * mad- av-r extended time periods. Mu, !I of those cuantitative data could be organized...The emphasis in this program was to deliver to AFRPL a data base on the microscopic and transient combustion of propellants which can be used in models

  1. Monitor for physical property changes in solid propellants

    Black, R. E., Jr.


    Specially designed sensor is attached to or imbedded in propellant. When sensor is driven into vibration, it moves with a phase lag directly proportional to internal friction or loss coefficent. Resonance frequency of the system is related to Young's modulus. Modulus or internal friction can be monitored over long period of time.

  2. Ignition models and simulation of solid propellant of thermodynamic undersea vehicle

    ZHANG Jin-jun; QIAN Zhi-bo; YANG Jie; YAN Ping


    The starting characteristics of thermodynamic undersea vehicle systems are determined by the geometry, size and combustion area of solid propellants, which directly effect liquid propellant pipeline design. It is necessary to establish accurate burning models for solid propellants. Based on combustion models using powder tings and two different solid ignition grains, namely star-shaped ignition grains and stuffed ignition grains, a mathematic model of the ignition process of the propulsion system was built.With the help of Matlah, a series of calculations were made to determine the effects of different grains on ignition characteristics. The results show that stuffed ignition grain is best suited to be the ignition grain of a thermodynamic undersea vehicle system.

  3. Infrared radiation signature of exhaust plume from solid propellants with different energy characteristics

    Wang Weichen; Li Shipeng; Zhang Qiao; Wang Ningfei


    The infrared radiation signature of the plume from solid propellants with different energy characteristics is not the same.Three kinds of double-base propellants of different energy characteristics are chosen to measure the infrared spectral radiance from 1000cm-1 to 4500 cm-1 of their plumes.The radiative spectrum is obtained in the tests.The experimental results indicate that the infrared radiation of the plume is determined by the energy characteristics of the propellant.The radiative transfer calculation models of the exhaust plume for the solid propellants are established.By including the chemical reaction source term and the radiation source term into the energy equation,the plume field and the radiative transfer are solved in a coupled way.The calculated results are consistent with the experimental data,so the reliability of the models is confirmed.The temperature distribution and the extent of the afterburning of the plume are distinct for the propellants of different energy characteristics,therefore the plume radiation varies for different propellants.The temperature of the fluid cell in the plume will increase or decrease to some extent by the influence of the radiation term.

  4. Homogenization Issues in the Combustion of Heterogeneous Solid Propellants

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


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

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

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


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

  6. Solid propellant combustion response to oscillatory radiant heat flux

    Strand, L. D.; Weil, M. T.; Cohen, N. S.


    A progress report is given on a research project to use the microwave Doppler velocimeter technique to measure the combustion response to an oscillating thermal radiation source (CO2 laser). The test technique and supporting analyses are described, and the results are presented for an initial test series on the nonmetallized, composite propellant, Naval Weapons Center formulation A-13. It is concluded that in-depth transmission of radiant heat flux is not a factor at the CO2 laser wave length.

  7. Flow visualization of a rocket injector spray using gelled propellant simulants

    Green, James M.; Rapp, Douglas C.; Roncace, James


    A study was conducted at NASA-Lewis to compare the atomization characteristics of gelled and nongelled propellant simulants. A gelled propellant simulant composed of water, sodium hydroxide, and an acrylic acid polymer resin (as the gelling agent) was used to simulate the viscosity of an aluminum/PR-1 metallized fuel gel. Water was used as a comparison fluid to isolate the rheological effects of the water-gel and to simulate nongelled RP-1. The water-gel was injected through the central orifice of a triplet injector element and the central post of a coaxial injector element. Nitrogen gas flowed through the outer orifices of the triplet injector element and through the annulus of the coaxial injector element and atomized the gelled and nongelled liquids. Photographs of the water-gel spray patterns at different operating conditions were compared with images obtained using water and nitrogen. A laser light was used for illumination of the sprays. The results of the testing showed that the water sprays produced a finer and more uniform atomization than the water-gel sprays. Rheological analysis of the water-gel showed poor atomization caused by high viscosity of water-gel delaying the transition to turbulence.

  8. Application of transient burning rate model of solid propellant in electrothermal-chemical launch simulation

    Yan-jie Ni


    Full Text Available A 30 mm electrothermal-chemical (ETC gun experimental system is employed to research the burning rate characteristics of 4/7 high-nitrogen solid propellant. Enhanced gas generation rates (EGGR of propellants during and after electrical discharges are verified in the experiments. A modified 0D internal ballistic model is established to simulate the ETC launch. According to the measured pressure and electrical parameters, a transient burning rate law including the influence of EGGR coefficient by electric power and pressure gradient (dp/dt is added into the model. The EGGR coefficient of 4/7 high-nitrogen solid propellant is equal to 0.005 MW−1. Both simulated breech pressure and projectile muzzle velocity accord with the experimental results well. Compared with Woodley's modified burning rate law, the breech pressure curves acquired by the transient burning rate law are more consistent with test results. Based on the parameters calculated in the model, the relationship among propellant burning rate, pressure gradient (dp/dt and electric power is analyzed. Depending on the transient burning rate law and experimental data, the burning of solid propellant under the condition of plasma is described more accurately.

  9. Application of transient burning rate model of solid propellant in electrothermal-chemical launch simulation

    Yan-jie NI; Yong JIN; Gang WAN; Chun-xia YANG; Hai-yuan LI; Bao-ming LI


    A 30 mm electrothermal-chemical (ETC) gun experimental system is employed to research the burning rate characteristics of 4/7 high-nitrogen solid propellant. Enhanced gas generation rates (EGGR) of propellants during and after electrical discharges are verified in the experiments. A modified 0D internal ballistic model is established to simulate the ETC launch. According to the measured pressure and electrical parameters, a transient burning rate law including the influence of EGGR coefficient by electric power and pressure gradient (dp/dt) is added into the model. The EGGR coefficient of 4/7 high-nitrogen solid propellant is equal to 0.005 MW−1. Both simulated breech pressure and projectile muzzle velocity accord with the experimental results well. Compared with Woodley’s modified burning rate law, the breech pressure curves acquired by the transient burning rate law are more consistent with test results. Based on the parameters calculated in the model, the relationship among propellant burning rate, pressure gradient (dp/dt) and electric power is analyzed. Depending on the transient burning rate law and experimental data, the burning of solid propellant under the condition of plasma is described more accurately.

  10. Determination of Activation Energy of Relaxation Events in Composite Solid Propellants by Dynamic Mechanical Analysis

    B. K Bihari


    Full Text Available The shelf life of a composite solid propellant is one of the critical aspects for the usage of solid propellants. To assess the ageing behavior of the composite solid propellant, the activation energy is a key parameter. The activation energy is determined by analysis of visco-elastic response of the composite solid propellant when subjected to sinusoidal excitation. In the present study, dynamic mechanical analyzer was used to characterize six different types of propellants based on hydroxyl terminated polybutadiene, aluminium, ammonium perchlorate cured with toluene diisocyanate having burning rates varying from 5 mm/s to 25 mm/s at 7000 kPa. Each propellant sample was given a multi-frequency strain of 0.01 percent at three discrete frequencies (3.5 Hz, 11 Hz, 35 Hz in the temperature range -80 °C to + 80 °C. It was observed that all the propellants have shown two relaxation events (α- and β- transition in the temperature range -80 °C to +80 °C. The α-transition was observed between -66 °C and -51 °C and β-transition between 7 °C and 44 °C for the propellants studied. The activation energy for both transitions was determined by Arrhenius plot from dynamic properties measured at different frequencies and also by time temperature superposition principle using Williams-Landel-Ferry and Arrhenius temperature dependence equations. The data reveal that the activation energy corresponding to α-transition varies from 90 kJ/mol to 125 kJ/mol for R-value between 0.7 to 0.9 while for β-transition the values are from 75 kJ/mol to 92 kJ/mol. The activation energy corresponding to β-transition may be used to predict the useful life of solid propellant.Defence Science Journal, 2014, 64(2, pp. 173-178. DOI:

  11. Liquid propellant rocket engine combustion simulation with a time-accurate CFD method

    Chen, Y. S.; Shang, H. M.; Liaw, Paul; Hutt, J.


    Time-accurate computational fluid dynamics (CFD) algorithms are among the basic requirements as an engineering or research tool for realistic simulations of transient combustion phenomena, such as combustion instability, transient start-up, etc., inside the rocket engine combustion chamber. A time-accurate pressure based method is employed in the FDNS code for combustion model development. This is in connection with other program development activities such as spray combustion model development and efficient finite-rate chemistry solution method implementation. In the present study, a second-order time-accurate time-marching scheme is employed. For better spatial resolutions near discontinuities (e.g., shocks, contact discontinuities), a 3rd-order accurate TVD scheme for modeling the convection terms is implemented in the FDNS code. Necessary modification to the predictor/multi-corrector solution algorithm in order to maintain time-accurate wave propagation is also investigated. Benchmark 1-D and multidimensional test cases, which include the classical shock tube wave propagation problems, resonant pipe test case, unsteady flow development of a blast tube test case, and H2/O2 rocket engine chamber combustion start-up transient simulation, etc., are investigated to validate and demonstrate the accuracy and robustness of the present numerical scheme and solution algorithm.

  12. Large Solid Rocket Motor Safety Analyses: Thermal Effects Issues


    aluminium droplet combustion and aluminium oxide condensation complete. Flame true temperature drops to 2235 ±7 °K and 2206 ±7 °K respectively at 26...of burning Al droplets and aluminium oxide particles is significant on flame radiation to the surroundings and toward the burning propellant

  13. Alternatives to Chlorinated Solvents in the Solid Rocket Motor Industry


    cleaners, such as a terpene /surfactant mixture. Organic solvents generally are effective in removing heavy oils and greases; however, they may pose...occur. Small decreases in DTA exotherms were observed indicating that some solvent was absorbed into the propellant; however, no significant safety

  14. Influence of lead Inorganic Compounds on Combustion Rate of Double Base Rocket Propellants

    V. B. Pillai


    Full Text Available The influence of lead nitrate, red lead, lead chromate, lead floride and lead carbonate on the combustion behaviour of double base propellants in the pressure range-35-140kg/cm /sup 2/ was studied. While all these compounds increased burning rates in lower pressure range (35-60 kg/cm/sup 2/ and higher pressure range (120-140 kg/cm/sup 2/, only lead chromate and lead fluoride were effective in the intermediate pressure range of 60-105 kg/cm/sup 2/. None of these compounds were effective as platonizer, except lead fluoride, which lowered n value to 0.34 in the lower pressure range. Addition of carbon black along with lead compounds raised burning rates further and reduced n values significantly in the higher pressure regins. A probable mechanism on the role of lead compounds studied has been suggested based on burning rate and DTA results.

  15. Aluminum flame temperature measurements in solid propellant combustion.

    Parigger, Christian G; Woods, Alexander C; Surmick, David M; Donaldson, A B; Height, Jonathan L


    The temperature in an aluminized propellant is determined as a function of height and plume depth from diatomic AlO and thermal emission spectra. Higher in the plume, 305 and 508 mm from the burning surface, measured AlO emission spectra show an average temperature with 1σ errors of 2980 ± 80 K. Lower in the plume, 152 mm from the burning surface, an average AlO emission temperature of 2450 ± 100 K is inferred. The thermal emission analysis yields higher temperatures when using constant emissivity. Particle size effects along the plume are investigated using wavelength-dependent emissivity models.

  16. Flame spectra of solid propellants during unstable combustion.

    Eisel, J. L.; Ryan, N. W.; Baer, A. D.


    The spectral and temporal details of the flames of a series of ammonium perchlorate-polyurethane propellants during both unstable and stable combustion were observed experimentally. A 400-scan per second optical spectrometer operating in the middle infrared region was used. During unstable combustion at low ratios of chamber free volume to nozzle throat area, three different frequencies were observed simultaneously. These were attributable to at least two mechanisms. During stable combustion periodic fluctuations in flame temperature and composition were also observed. Some aspects of theory of bulk mode instability were confirmed, but the assumptions of constant flame temperature and constant composition were found to be inaccurate.

  17. Cu–Co–O nano-catalysts as a burn rate modifier for composite solid propellants

    D. Chaitanya Kumar Rao


    Full Text Available Nano-catalysts containing copper–cobalt oxides (Cu–Co–O have been synthesized by the citric acid (CA complexing method. Copper (II nitrate and Cobalt (II nitrate were employed in different molar ratios as the starting reactants to prepare three types of nano-catalysts. Well crystalline nano-catalysts were produced after a period of 3 hours by the calcination of CA–Cu–Co–O precursors at 550 °C. The phase morphologies and crystal composition of synthesized nano-catalysts were examined using Scanning Electron Microscope (SEM, Energy Dispersive Spectroscopy (EDS and Fourier Transform Infrared Spectroscopy (FTIR methods. The particle size of nano-catalysts was observed in the range of 90 nm–200 nm. The prepared nano-catalysts were used to formulate propellant samples of various compositions which showed high reactivity toward the combustion of HTPB/AP-based composite solid propellants. The catalytic effects on the decomposition of propellant samples were found to be significant at higher temperatures. The combustion characteristics of composite solid propellants were significantly improved by the incorporation of nano-catalysts. Out of the three catalysts studied in the present work, CuCo-I was found to be the better catalyst in regard to thermal decomposition and burning nature of composite solid propellants. The improved performance of composite solid propellant can be attributed to the high crystallinity, low agglomeration and lowering the decomposition temperature of oxidizer by the addition of CuCo-I nano-catalyst.

  18. Development of the platelet micro-orifice injector. [for liquid propellant rocket engines

    La Botz, R. J.


    For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.

  19. Review of Solid Propellant Ignition Models Relative to the Interior Ballistic Modelling of Gun Systems


    the point of phase change is reached. Solid-phase exothermic reactions may occur in some propellant ingredients (such as AP [9] or ADN [4]) leading...M.L. Gross. Two-dimensional modeling of AP/HTPB utilizing a vorticity formula- tion and one-dimensional modeling of AP and ADN . PhD thesis, Brigham

  20. Mutagenic Potential of DIGL-RO Solid Propellant in the Ames Salmonella/ Mammalian Microsome Mutagenicity Test


    303913 11. TITLE (Include Security Clasification ) Mutagenic Potential of DIGL-RP Solid Propellant in the Ames Salmonella/Mammalian Microsome protein /plate. After all the ingredients were added, the top agar was mixed, then overlaid on MGA plates. These plates contained 2% glucose and

  1. A MEMS-based solid propellant microthruster array for space and military applications

    Chaalane, A.; Chemam, R.; Houabes, M.; Yahiaoui, R.; Metatla, A.; Ouari, B.; Metatla, N.; Mahi, D.; Dkhissi, A.; Esteve, D.


    Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based solid propellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO2/SiNx thin membrane to ensure a good heat transfer to the propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x%BP) propellant composed of a DoubleBase and Black-Powder. A temperature of 250 °C, enough to propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20%BP propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 μm of throat wide.

  2. Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

    Popov, Pavel; Sideris, Athanasios; Sirignano, William


    We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

  3. Inconsistent Definitions of the Pressure-Coupled Response and the Admittance of Solid Propellants

    Cardiff, Eric H.


    When an acoustic wave is present in a solid propellant combustion environment, the mass flux from the combustion zone oscillates at the same frequency as the acoustics. The acoustic wave is either amplified or attenuated by the response of the combustion to the acoustic disturbance. When the acoustic wave is amplified, this process is called combustion instability. The amplification is quantitatively measured by a response function. The ability to predict combustion stability for a solid propellant formulation is essential to the formulator to prevent or minimize the effects of instabilities, such as an oscillatory thrust. Unfortunately, the prediction of response values for a particular propellant remains a technical challenge. Most predictions of the response of propellants are based on test data, but there are a number of questions about the reliability of the standard test method, the T-burner. Alternate methods have been developed to measure the response of a propellant, including the ultrasound burner, the magnetic flowmeter and the rotating valve burner, but there are still inconsistencies between the results obtained by these different methods. Aside from the experimental differences, the values of the pressure-coupled responses obtained by different researchers are often compared erroneously, for the simple reason that inconsistencies in the definitions of the responses and admittances are not considered. The use of different definitions has led to substantial confusion since the first theoretical treatments of the problem by Hart and McClure in 1959. The definitions and relations derived here seek to alleviate this problem.

  4. Energetics and Compatibility of Plasticizers in Composite Solid Propellants

    Rm. Muthiah


    Full Text Available In this paper a comparative analysis on the energetics of ester type plasticizers such as dioctyl adipate (DOA, dioctyl phthalate (DOP,dibutyl sebacate (DBS, isodecyl pelargonate (IDP, trioctylphosphate (TOF, diethyl phthalate (DEP, tricresyl phosphate (TCPand dibutyl phthalate (DBP and hydrocarbon type plasticizers such as polybutene (PB, spindle oil, naphthenic oil, polymer extender oil(PEO and poly isobutylene (PIB and the impact of some of the plasticizers on the work ability, pot life and mechanical properties of propellants based on two selected polymeric binders namely polybutadiene-acrylic acid-acrylonitrile (PBAN ter polymer andhydroxyl terminated polybutadiene (HTPB have been reported. The compatibility of all the plasticizers on HTPB binder was also studied at different concentration levels and temperatures using Brookfield viscometer and reported. The mechanism of plasticization is also reviewed.

  5. A detailed numerical simulation of a liquid-propellant rocket engine ground test experiment

    Lankford, D. W.; Simmons, M. A.; Heikkinen, B. D.


    A computational simulation of a Liquid Rocket Engine (LRE) ground test experiment was performed using two modeling approaches. The results of the models were compared with selected data to assess the validity of state-of-the-art computational tools for predicting the flowfield and radiative transfer in complex flow environments. The data used for comparison consisted of in-band station radiation measurements obtained in the near-field portion of the plume exhaust. The test article was a subscale LRE with an afterbody, resulting in a large base region. The flight conditions were such that afterburning regions were observed in the plume flowfield. A conventional standard modeling approach underpredicted the extent of afterburning and the associated radiation levels. These results were attributed to the absence of the base flow region which is not accounted for in this model. To assess the effects of the base region a Navier-Stokes model was applied. The results of this calculation indicate that the base recirculation effects are dominant features in the immediate expansion region and resulted in a much improved comparison. However, the downstream in-band station radiation data remained underpredicted by this model.

  6. Prediction of Maximum Strain in Finocyl Port Case-bonded Solid Propellants under Pressure Loading

    Himanshu Shekhar


    Full Text Available Finite element analysis of case-bonded solid propellants in finocyl port configuration hasbeen carried out using finite element method. The parametric studies have also been conductedfor loading conditions, material properties, and geometrical configurations. The results arepresented in the form of a universal power law, which can be utilised for primary assessmentof peak strain in any finocyl port propellant configuration without using finite element software.This eliminates dependence on finite element software for structural integrity analysis of solidpropellants in finocyl port configuration under port pressurisation. The results obtained by finiteelement analysis and power law are in close agreement.

  7. Energetic Properties of Rocket Propellants Evaluated through the Computational Determination of Heats of Formation of Nitrogen-Rich Compounds.

    Forquet, Valérian; Miró Sabaté, Carles; Chermette, Henry; Jacob, Guy; Labarthe, Émilie; Delalu, Henri; Darwich, Chaza


    The use of ab initio and DFT methods to calculate the enthalpies of formation of solid ionic compounds is described. The results obtained from the calculations are then compared with those from experimental measurements on nitrogen-rich salts of the 2,2-dimethyltriazanium cation (DMTZ) synthesized in our laboratory and on other nitrogen-rich ionic compounds. The importance of calculating accurate volumes and lattice enthalpies for the determination of heats of formation is also discussed. Furthermore, the crystal structure and hydrogen-bonding networks of the nitroformate salt of the DMTZ cation is described in detail. Lastly, the theoretical heats of formation were used to calculate the specific impulses (Isp ) of the salts of the DMTZ cation in view of a prospective application in propellant formulations.

  8. Hierarchical Theoretical Methods for Understanding and Predicting Anisotropic Thermal Transport Release in Rocket Propellant Formulations


    solid curve) and continuous indentation (purple curve for loading and green curve for unloading from 30 Å). The Hertzian prediction is shown as the flash method," Report No. UCRL-52565, Lawrence Livermore Laboratory, October 1978. DISTRIBUTION A: Distribution approved for public release...Mass fraction: liquid RDX(blue), intermediate( red), gas phase( green ) In this case we tested the concept of using mass fractions, defined by binning

  9. Single-pulse CARS spectra in solid propellant combustion at atmosphere pressure

    Zhiyun Hu(胡志云); Jingru Liu(刘晶儒); Zhenrong Zhang(张振荣); Xiaowei Guan(关小伟); Meisheng Huang(黄梅生); Xisheng Ye(叶锡生)


    The unstable-resonator spatially enhanced detection (USED) coherent anti-stokes Raman spectroscopy (CARS) measurements of temperature and N2 concentration in the combustion of solid propellant at atmosphere pressure are reported. The USED CARS measurement system has a high spatial solution of ~ 0.1 mm in diameter and 3 mm in length, and permits instantaneous measurement at 10-Hz rate. The single-pulse N2 Q-branch CARS spectra have been obtained from the propellant combustion. The temperatures and N2 concentrations of the propellant flame at different height have been achieved by fitting the experimental data to theoretical spectra. The results indicate that the temperature is up to~2500 K with N2 concentration in a range from 10% to 26%.

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

    Goldberg, Ben E.; Wiley, Dan R.


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

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

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


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

  12. Overpressure wave generated by the pneumatic explosion of a solid propellant-fuelled propeller; Onde de surpression generee par l`explosion pneumatique d`un propulseur a propergol solide

    Rolland, C.; Roussille, A. [DGA, Direction des Missiles et de l`Espace, Centre d`Achevement et d`Essais des Propulseurs et Engins, 33 - Saint Medard en Jalles (France)


    The CAEPE, the French Centre of Propellers and Engines Completion and Testing is in charge of the static tests of solid propellant fuelled propellers. In order to evaluate the risks linked with the explosion of such propellers, an analytical method is proposed which allows to determine the energy, impulses and overpressure due to an explosion occurring at any time of the firing process. This method is based on the exploitation of the shock tube equations. Concerning the overpressure wave propagation, several laws are compared to full scale tests performed at the CAEPE. (J.S.) 14 refs.

  13. Thermographic inspection of solid-fuel rocket booster field joint components

    Thompson, Karen G.; Crisman, Elton M.


    Thermographic nondestructive evaluation techniques were investigated for possible application on Space Shuttle solid rocket booster field joint hardware. This investigation included evaluation of the clevis and tang mating surfaces for scratches and measurement of grease film thickness. The field joint insulation system was inspected for voids and disbonds.

  14. Transient Burning Rate Model for Solid Rocket Motor Internal Ballistic Simulations

    David R. Greatrix


    Full Text Available A general numerical model based on the Zeldovich-Novozhilov solid-phase energy conservation result for unsteady solid-propellant burning is presented in this paper. Unlike past models, the integrated temperature distribution in the solid phase is utilized directly for estimating instantaneous burning rate (rather than the thermal gradient at the burning surface. The burning model is general in the sense that the model may be incorporated for various propellant burning-rate mechanisms. Given the availability of pressure-related experimental data in the open literature, varying static pressure is the principal mechanism of interest in this study. The example predicted results presented in this paper are to a substantial extent consistent with the corresponding experimental firing response data.

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

    Marsh, H. E., Jr.


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

  16. The Research on Transient Burning Rate of Solid Propellant by Digital Image Processing

    Xin Peng


    Full Text Available In order to obtain the burn rate of the solid propellant that is the important parameter of transient burning, the new method named digital image processing is presented. In the article , the principle of digital image processing is analysed; The burning face of the sample in the each time is located according the image and the coordinates of the burning face is obtained. In experiment the transient burn rate is measured by digital image processing and the accuracy is acceptable.

  17. Application of Al/B/Fe2O3 Nano Thermite in Composite Solid Propellant

    Jingke Deng


    Full Text Available Hydroxyl-terminated polybutadiene (HTPB propellant were prepared with different content of Al/B/Fe2O3 nano thermite, and the mechanical, thermal and energetic performances were studied. Al/B/Fe2O3 nano thermite exhibited good compatibility with HTPB and dioctyl sebacate (DOS through differential scanning calorimetry (DSC tests. Mechanical experiments show that the mechanical properties of HTPB propellant could be improved by the addition of a small quantity of Al/B/Fe2O3 nano thermite, compared with the absence of Al/B/Fe2O3 nano thermite. For example, with the addition of 3% Al/B/Fe2O3 nano thermite, the tensile strength and elongation of propellant had the increase of 15.3% and 32.1%, respectively. Thermal analysis indicated that the decomposition of ammonium perchlorate (AP in HTPB propellant could be catalyzed by Al/B/Fe2O3 nano thermite, the high-temperature exothermic peak of AP was shifted to lower temperature by 70.8 °C when the content of Al/B/Fe2O3 nano thermite was 5%, and the heat released was enhanced by 70%. At the same time, the heat of explosion of HTPB propellant could also be enhanced by the addition of Al/B/Fe2O3 nano thermite. Copyright © 2016 BCREC GROUP. All rights reserved  Received: 5th November 2015; Revised: 4th December 2015; Accepted: 30th December 2015 How to Cite: Deng, J., Li, G., Shen, L., Luo, Y. (2016. Application of Al/B/Fe2O3 Nano Thermite in Composite Solid Propellant. Bulletin of Chemical Reaction Engineering & Catalysis, 11 (1: 109-114. (doi:10.9767/bcrec.11.1.432.109-114 Permalink/DOI:

  18. Experimental investigation of the combustion products in an aluminised solid propellant

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


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

  19. [Studies on spectral characteristics of solid propellant by remote sensing FTIR].

    Li, Yan; Huang, Zhong-hua; Zhou, Xue-tie; Wang, Jun-de


    The infrared spectral characteristics of high-intensity IR solid propellant were measured in this paper using a Bruker EQUINOX55 remote sensing FTIR spectrometer. The emission spectra of the combustion flame were recorded in the range of 4700-1800 cm(-1) with a spectral resolution of 4 cm(-1). The combustion temperatures of solid propellant at the burning time of 0, 9, 18, 27 and 36 s calculated from the molecular rotation-vibration spectra were 1992.5, 1610.9, 2294.4, 2361.1 and 1916.9 K, respectively. Moreover, the spectral radiance distributions of the high IR flare material at different times were given, and the combustion product concentrations of HCl, HF, CO2 and CO were determined quantitatively. Results showed that remote sensing FTIR is a potential technology that can be applied to the measurement of IR spectral characteristics, especially to the identification of the IR objects, guidance and anti-guidance in the military, and the modification of the make-up of solid propellant.

  20. Experimental study on composite solid propellant material burning rate using algorithm MATLAB

    Thunaipragasam Selvakumaran


    Full Text Available In rocketry application, now-a-days instead of monopropellants slowly composite propellants are introduced. Burning rate of a solid state composite propellant depends on many factors like oxidizer-binder ratio, oxidizer particle size and distribution, particle size and its distribution, pressure, temperature, etc. Several researchers had taken the mass varied composite propellant. In that, the ammonium perchlorate mainly varied from 85 to 90%. This paper deals with the oxidizer rich propellant by allowing small variation of fuel cum binder ranging from 2%, 4%, 6%, and 8% by mass. Since the percent of the binder is very less compared to the oxidizer, the mixture remains in a powder form. The powder samples are used to make a pressed pellet. Experiments were conducted in closed window bomb set-up at pressures of 2, 3.5, and 7 MN/m2. The burning rates are calculated from the combustion photography (images taken by a high-speed camera. These images were processed frame by frame in MATLAB, detecting the edges in the images of the frames. The burning rate is obtained as the slope of the linear fit from MATLAB and observed that the burn rate increases with the mass variation of constituents present in solid state composite propellant. The result indicates a remarkable increase in burn rate of 26.66%, 20%, 16.66%, and 3.33% for Mix 1, 2, 3, 4 compared with Mix 5 at 7 MN/m2. The percentage variations in burn rate between Mix 1 and Mix 5 at 2, 3.5, and 7 MN/m2 are 25.833%, 32.322%, and 26.185%, respectively.

  1. Research on Instantaneous Thrust Measurement for Attitude-control Solid Rocket Motor

    OUYANG Hua-bing; WANG Jian-ping; LIN Feng; XU Wen-gan


    In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.

  2. The development of a solid-state hydrogen sensor for rocket engine leakage detection

    Liu, Chung-Chiun

    Hydrogen propellant leakage poses significant operational problems in the rocket propulsion industry as well as for space exploratory applications. Vigorous efforts have been devoted to minimizing hydrogen leakage in assembly, test, and launch operations related to hydrogen propellant. The objective has been to reduce the operational cost of assembling and maintaining hydrogen delivery systems. Specifically, efforts have been made to develop a hydrogen leak detection system for point-contact measurement. Under the auspices of Lewis Research Center, the Electronics Design Center at Case Western Reserve University, Cleveland, Ohio, has undertaken the development of a point-contact hydrogen gas sensor with potential applications to the hydrogen propellant industry. We envision a sensor array consisting of numbers of discrete hydrogen sensors that can be located in potential leak sites. Silicon-based microfabrication and micromachining techniques are used in the fabrication of these sensor prototypes. Evaluations of the sensor are carried out in-house at Case Western Reserve University as well as at Lewis Research Center and GenCorp Aerojet, Sacramento, California. The hydrogen gas sensor is not only applicable in a hydrogen propulsion system, but also usable in many other civilian and industrial settings. This includes vehicles or facility use, or in the production of hydrogen gas. Dual space and commercial uses of these point-contacted hydrogen sensors are feasible and will directly meet the needs and objectives of NASA as well as various industrial segments.

  3. Optimization of Construction of the rocket-assisted projectile

    Arkhipov Vladimir


    Full Text Available New scheme of the rocket motor of rocket-assisted projectile providing the increase in distance of flight due to controlled and optimal delay time of ignition of the solid-propellant charge of the SRM and increase in reliability of initiation of the SRM by means of the autonomous system of ignition excluding the influence of high pressure gases of the propellant charge in the gun barrel has been considered. Results of the analysis of effectiveness of using of the ignition delay device on motion characteristics of the rocket-assisted projectile has been presented.

  4. Measurement and Characterization of Space Shuttle Solid Rocket Motor Plume Acoustics

    Kenny, Jeremy; Hobbs, Chris; Plotkin, Ken; Pilkey, Debbie


    Lift-off acoustic environments generated by the future Ares I launch vehicle are assessed by the NASA Marshall Space Flight Center (MSFC) acoustics team using several prediction tools. This acoustic environment is directly caused by the Ares I First Stage booster, powered by the five-segment Reusable Solid Rocket Motor (RSRMV). The RSRMV is a larger-thrust derivative design from the currently used Space Shuttle solid rocket motor, the Reusable Solid Rocket Motor (RSRM). Lift-off acoustics is an integral part of the composite launch vibration environment affecting the Ares launch vehicle and must be assessed to help generate hardware qualification levels and ensure structural integrity of the vehicle during launch and lift-off. Available prediction tools that use free field noise source spectrums as a starting point for generation of lift-off acoustic environments are described in the monograph NASA SP-8072: "Acoustic Loads Generated by the Propulsion System." This monograph uses a reference database for free field noise source spectrums which consist of subscale rocket motor firings, oriented in horizontal static configurations. The phrase "subscale" is appropriate, since the thrust levels of rockets in the reference database are orders of magnitude lower than the current design thrust for the Ares launch family. Thus, extrapolation is needed to extend the various reference curves to match Ares-scale acoustic levels. This extrapolation process yields a subsequent amount of uncertainty added upon the acoustic environment predictions. As the Ares launch vehicle design schedule progresses, it is important to take every opportunity to lower prediction uncertainty and subsequently increase prediction accuracy. Never before in NASA s history has plume acoustics been measured for large scale solid rocket motors. Approximately twice a year, the RSRM prime vendor, ATK Launch Systems, static fires an assembled RSRM motor in a horizontal configuration at their test facility

  5. Migration of a Moonlet in a Ring of Solid Particles: Theory and Application to Saturn's Propellers

    Crida, Aurélien; Papaloizou, John C. B.; Rein, Hanno; Charnoz, Sébastien; Salmon, Julien


    Hundred-meter-sized objects have been identified by the Cassini spacecraft in Saturn's A ring through the so-called propeller features they create in the ring. These moonlets should migrate due to their gravitational interaction with the ring; in fact, some orbital variations have been detected. The standard theory of type I migration of planets in protoplanetary disks cannot be applied to the ring system as it is pressureless. Thus, we compute the differential torque felt by a moonlet embedded in a two-dimensional disk of solid particles, with a flat surface density profile, both analytically and numerically. We find that the corresponding migration rate is too small to explain the observed variations of the propeller's orbit in Saturn's A ring. However, local density fluctuations (due to gravity wakes in the marginally gravitationally stable A ring) may exert a stochastic torque on a moonlet. Our simulations show that this torque can be large enough to account for the observations depending on the parameters of the rings. We find that on timescales of several years the migration of propellers is likely to be dominated by stochastic effects (while the former, non-stochastic migration dominates after ~104-105 years). In that case, the migration rates provided by observations so far suggest that the surface density of the A ring should be on the order of 700 kg m-2. The age of the propellers should not exceed 1-100 million years depending on the dominant migration regime.


    Oladipupo Olaosebikan Ogunleye


    Full Text Available This study was designed to evaluate the effect of propellant formulation and geometry on the solid propellant grains internal ballistic performance using core, bates, rod and tubular and end-burn geometries. Response Surface Methodology (RSM was used to analyze and optimize the effect of sucrose, potassium nitrate and carbon on the chamber pressure, temperature, thrust and specific impulse of the solid propellant grains through Central Composite Design (CCD of the experiment. An increase in potassium nitrate increased the specific impulse while an increase in sucrose and carbon decreased specific impulse. The coefficient of determination (R2 for models of chamber pressure, temperature, thrust and specific impulse in terms of composition and geometry were 0.9737, 0.9984, 0.9745 and 0.9589, respectively. The optimum specific impulse of 127.89 s, pressure (462201 Pa, temperature (1618.3 K and thrust (834.83 N were obtained using 0.584 kg of sucrose, 1.364 kg of potassium nitrate and 0.052 kg of carbon as well as bate geometry. There was no significant difference between the calculated and experimented ballistic properties at p < 0.05. The bate grain geometry is more efficient for minimizing the oscillatory pressure in the combustion chamber.

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

    Benreuven, M.; Summerfield, M.


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

  8. Impact of the Surface Morphology on the Combustion of Simulated Solid Rocket Motor

    Abdelkarim M. Hegab


    Full Text Available An advanced and intensive computational solution development is integrated with an asymptotic technique, to examine the impact of the combustion surface morphology on the generated rotational flow field in a solid rocket chamber with wide ranges of forcing frequencies. The simulated rectangular chamber is closed at one end and is open at the aft end. The upper and lower walls are permeable to allow steady and unsteady injected air to generate internal flow mimicking the flow field of the combustion gases in real rocket chamber. The frequencies of the unsteady injected flow are chosen to be very close or away from the resonance frequencies of the adapted chamber. The current study accounts for a wide range of wave numbers that reflect the complexity of real burning processes. Detailed derivation for Navier-Stokes equations at the four boundaries of the chamber is introduced in the current study. Qualitative comparison is performed with recent experimental work carried out on a two-inch hybrid rocket motor using a mixture of polyethylene and aluminum powder. The higher the percentage of aluminum powder in the mixture, the more the corrugations of the combustion surface. This trend is almost similar to the computational and analytical results of a simulated solid rocket chamber.

  9. The Use of Fractal for Prediction of Burning Rate of Composite Solid Propellants

    ManouchehrNikazar; MohammadB.Bagherpour; 等


    By using fractal geometry is is possible to calculate the actual AP(Ammonium Perchlorate)surface area and oxidizer-binder interface fractal dimension in the prediction of burning rate of commposite solid propellants.In this investigation,the fractal dimension was determined by a procedure known as the "Box counting Method".using this dimensio,surface area relations were developed for the rough particles.This method was implemented in the PEM(Petite Ensemble Model) burning rate model,The comparison of burning rates for a typical propellant by the PEM and fractal model shows that the burning rates botained by using the fractal geometry are slightly less than those obtained by the PEM model.

  10. The direct effects of strain on burning rates of composite solid propellants

    Langhenry, M. T.


    A mathematical model is developed to predict burn rate augmentation due to strain in a composite solid propellant. The model assumes the effect is due to the ability of the flame to penetrate the small fissures and voids that form when a propellant is strained. The number and size of these fissures is obtained by applying a flaw propagation analysis to randomly distributed flaws that form when the binder-oxidizer particle bonds break under stress. A flame height is calculated with Summerfield's burn rate equation and is used to compute the burn rate augmentation based upon the additional burn area created when the flame penetrates the fissures. Comparisons are made with data obtained from published sources. The existence of threshold pressure and strains, above which augmentation occurs, is verified although the model predicts a lower threshold pressure and higher threshold strain than expected. Further results and applications of the model are discussed.

  11. Failure mode and effects analysis (FMEA) for the Space Shuttle solid rocket motor

    Russell, D. L.; Blacklock, K.; Langhenry, M. T.


    The recertification of the Space Shuttle Solid Rocket Booster (SRB) and Solid Rocket Motor (SRM) has included an extensive rewriting of the Failure Mode and Effects Analysis (FMEA) and Critical Items List (CIL). The evolution of the groundrules and methodology used in the analysis is discussed and compared to standard FMEA techniques. Especially highlighted are aspects of the FMEA/CIL which are unique to the analysis of an SRM. The criticality category definitions are presented and the rationale for assigning criticality is presented. The various data required by the CIL and contribution of this data to the retention rationale is also presented. As an example, the FMEA and CIL for the SRM nozzle assembly is discussed in detail. This highlights some of the difficulties associated with the analysis of a system with the unique mission requirements of the Space Shuttle.

  12. An improved heat transfer configuration for a solid-core nuclear thermal rocket engine

    Clark, John S.; Walton, James T.; Mcguire, Melissa L.


    Interrupted flow, impingement cooling, and axial power distribution are employed to enhance the heat-transfer configuration of a solid-core nuclear thermal rocket engine. Impingement cooling is introduced to increase the local heat-transfer coefficients between the reactor material and the coolants. Increased fuel loading is used at the inlet end of the reactor to enhance heat-transfer capability where the temperature differences are the greatest. A thermal-hydraulics computer program for an unfueled NERVA reactor core is employed to analyze the proposed configuration with attention given to uniform fuel loading, number of channels through the impingement wafers, fuel-element length, mass-flow rate, and wafer gap. The impingement wafer concept (IWC) is shown to have heat-transfer characteristics that are better than those of the NERVA-derived reactor at 2500 K. The IWC concept is argued to be an effective heat-transfer configuration for solid-core nuclear thermal rocket engines.

  13. Probabilistic Fracture Mechanics and Optimum Fracture Control Analytical Procedures for a Reusable Solid Rocket Motor Case

    Hanagud, S.; Uppaluri, B.


    A methodology for the reliability analysis of a reusable solid rocket motor case is discussed. The analysis is based on probabilistic fracture mechanics and probability distribution for initial flaw sizes. The developed reliability analysis is used to select the structural design variables of the solid rocket motor case on the basis of minimum expected cost and specified reliability bounds during the projected design life of the case. Effects of failure prevention plans such as nondestructive inspection and the material erosion between missions are also considered in the developed procedure for selection of design variables. The reliability-based procedure can be modified to consider other similar structures of reusable space vehicle systems with different failure prevention plans.

  14. Palynological Investigation of Post-Flight Solid Rocket Booster Foreign Material

    Nelson, Linda; Jarzen, David


    Investigations of foreign material in a drain tube, from the Solid Rocket Booster (SRB) of a recent Space Shuttle mission, was identified as pollen. The source of the pollen is from deposits made by bees, collecting pollen from plants found at the Kennedy Space Center, Cape Canaveral, Florida. The pollen is determined to have been present in the frustum drain tubes before the shuttle flight. During the flight the pollen did not undergo thermal maturation.

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

    Palaszewski, Bryan


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

  16. Numerical study on similarity of plume infrared radiation between reduced-scale solid rocket motors

    Zhang Xiaoying; Chen Huandong


    This study seeks to determine the similarities in plume radiation between reduced and full-scale solid rocket models in ground test conditions through investigation of flow and radiation for a series of scale ratios ranging from 0.1 to 1. The radiative transfer equation (RTE) considering gas and particle radiation in a non-uniform plume has been adopted and solved by the finite volume method (FVM) to compute the three dimensional, spectral and directional radiation of a plume in the infrared waveband 2–6μm. Conditions at wavelengths 2.7μm and 4.3μm are discussed in detail, and ratios of plume radiation for reduced-scale through full-scale models are examined. This work shows that, with increasing scale ratio of a computed rocket motor, area of the high-temperature core increases as a 2 power function of the scale ratio, and the radiation intensity of the plume increases with 2–2.5 power of the scale ratio. The infrared radiation of plume gases shows a strong spectral dependency, while that of Al2O3 particles shows spectral continuity of gray media. Spectral radiation intensity of a computed solid rocket plume’s high temperature core increases sig-nificantly in peak radiation spectra of plume gases CO and CO2. Al2O3 particles are the major radi-ation component in a rocket plume. There is good similarity between contours of plume spectral radiance from different scale models of computed rockets, and there are two peak spectra of radi-ation intensity at wavebands 2.7–3.0μm and 4.2–4.6μm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle.

  17. Test and Analysis of Solid Rocket Motor Nozzle Ablative Materials

    Clayton, J. Louie


    Asbestos free solid motor internal insulation samples were tested at the MSFC Hyperthermal Facility. Objectives of the test were to gather data for analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Instrumentation included use of total calorimeters, thermocouples, and a surface pyrometer for surface temperature measurement. Post-test sample forensics involved measurement of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero-thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

  18. Chemical reactions identified in the Titan 2, Titan 4, and Delta 2 propellant systems and their application to source modeling

    Prince, S. P.; Banning, D. W.; Wiseman, F. L.


    A series of tests involving the combustion of solid and liquid propellants used to fuel the Titan 2, Titan 4, and Delta 2 launch vehicles was performed. The purpose of these tests was to evaluate the nature and amounts of combustion gases from reacting these propellants in various proportions, and to apply the derived data to predicting toxic chemical emissions arising from a launch vehicle explosion. Propellants tested in this study included Aerozine-50 and nitrogen tetroxide (liquid propellants used in the Titan 2 and Titan 4 launch vehicles), PBAN solid propellant (used on the Titan 4 solid rocket motor), RP-1 and liquid oxygen (liquid propellants used to fuel the Delta 2 launch vehicle), and the Castor IVA solid rocket propellant used on the Delta 2 first stage engine. Tests were conducted in a 150-liter stainless steel combustion chamber in air at nominal pressure (0.8 atmospheres at Denver barometric conditions). Measurements of the chamber gas temperature and internal pressure were taken and gas samples were withdrawn and analyzed for expected combustion gases, unreacted propellants, organic vapors, and oxygen reacted from the air. A stainless steel witness plate was used to collect condensates which formed during the course of the propellant combustion tests. Results of this study suggest significantly different chemical fates for some of the rocket propellants than those predicted by chemical theory only. A description of the test parameters, results, and application to source predictions is presented.

  19. Real-Time X-ray Radiography Diagnostics of Components in Solid Rocket Motors

    Cortopassi, A. C.; Martin, H. T.; Boyer, E.; Kuo, K. K.


    Solid rocket motors (SRMs) typically use nozzle materials which are required to maintain their shape as well as insulate the underlying support structure during the motor operation. In addition, SRMs need internal insulation materials to protect the motor case from the harsh environment resulting from the combustion of solid propellant. In the nozzle, typical materials consist of high density graphite, carbon-carbon composites and carbon phenolic composites. Internal insulation of the motor cases is typically a composite material with carbon, asbestos, Kevlar, or silica fibers in an ablative matrix such as EPDM or NBR. For both nozzle and internal insulation materials, the charring process occurs when the hot combustion products heat the material intensely. The pyrolysis of the matrix material takes away a portion of the thermal energy near the wall surface and leaves behind a char layer. The fiber reinforcement retains the porous char layer which provides continued thermal protection from the hot combustion products. It is of great interest to characterize both the total erosion rates of the material and the char layer thickness. By better understanding of the erosion process for a particular ablative material in a specific flow environment, the required insulation material thickness can be properly selected. The recession rates of internal insulation and nozzle materials of SRMs are typically determined by testing in some sort of simulated environment; either arc-jet testing, flame torch testing, or subscale SRMs of different size. Material recession rates are deduced by comparison of pre- and post-test measurements and then averaging over the duration of the test. However, these averaging techniques cannot be used to determine the instantaneous recession rates of the material. Knowledge of the variation in recession rates in response to the instantaneous flow conditions during the motor operation is of great importance. For example, in many SRM configurations

  20. Application of fractional calculus to modeling transient combustion of solid propellants

    Kulish, Vladimir; Horák, Vladimír; Duc, Linh Do; Lukáč, Tomáš


    It was Zel'dovich, who first considered the transient combustion problem of solid propellants. Some more detailed models of that process have been developed afterwards. However, until today, numerical methods remain the prevailing tool for modeling unsteady combustion processes. In this work, it has been demonstrated that at least one of the problems of the unsteady combustion theory, which previously investigated numerically, can be treated analytically by means of fractional calculus. The solution for the unsteady speed of combustion thus derived is then compared with the solution obtained by numerical means in previous studies. The comparison shows a good agreement between those results, especially for small values of time.

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

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


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

  2. Feasibility demonstration of a variable frequency driver-microwave transient regression rate measurement system. [for solid propellant combustion response

    Strand, L. D.; Mcnamara, R. P.


    The feasibility of a system capable of rapidly and directly measuring the low-frequency (motor characteristics length bulk mode) combustion response characteristics of solid propellants has been investigated. The system consists of a variable frequency oscillatory driver device coupled with an improved version of the JPL microwave propellant regression rate measurement system. The ratio of the normalized regression rate and pressure amplitudes and their relative phase are measured as a function of varying pressure level and frequency. Test results with a well-characterized PBAN-AP propellant formulation were found to compare favorably with the results of more conventional stability measurement techniques.

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

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


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

  4. Advanced Tactical Booster Technologies: Applications for Long-Range Rocket Systems


    System HIMARS [3] which can employ the MGM- 140 Army Tactical Missile System (ATacMS) solid propellant missile [4] to achieve the required range...launcher. 15. SUBJECT TERMS solid rocket; optimisation; artillery 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT 18. NUMBER OF PAGES 19a...the volumetrically constrained environment of a land-based launcher. Keywords— solid rocket; optimisation; artillery I. INTRODUCTION The Australian

  5. State Machine Modeling of the Space Launch System Solid Rocket Boosters

    Harris, Joshua A.; Patterson-Hine, Ann


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

  6. Thermo-Structural Response Caused by Structure Gap and Gap Design for Solid Rocket Motor Nozzles

    Lin Sun


    Full Text Available The thermo-structural response of solid rocket motor nozzles is widely investigated in the design of modern rockets, and many factors related to the material properties have been considered. However, little work has been done to evaluate the effects of structure gaps on the generation of flame leaks. In this paper, a numerical simulation was performed by the finite element method to study the thermo-structural response of a typical nozzle with consideration of the structure gap. Initial boundary conditions for thermo-structural simulation were defined by a quasi-1D model, and then coupled simulations of different gap size matching modes were conducted. It was found that frictional interface treatment could efficiently reduce the stress level. Based on the defined flame leak criteria, gap size optimization was carried out, and the best gap matching mode was determined for designing the nozzle. Testing experiment indicated that the simulation results from the proposed method agreed well with the experimental results. It is believed that the simulation method is effective for investigating thermo-structural responses, as well as designing proper gaps for solid rocket motor nozzles.

  7. Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

    Nielsen, D.; Townsend, J.; Kappus, K.; Driskill, T.; Torres, I.; Parks, R.


    The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.

  8. Response of selected plant and insect species to simulated solid rocket exhaust mixtures and to exhaust components from solid rocket fuels

    Heck, W. W.; Knott, W. M.; Stahel, E. P.; Ambrose, J. T.; Mccrimmon, J. N.; Engle, M.; Romanow, L. A.; Sawyer, A. G.; Tyson, J. D.


    The effects of solid rocket fuel (SRF) exhaust on selected plant and and insect species in the Merritt Island, Florida area was investigated in order to determine if the exhaust clouds generated by shuttle launches would adversely affect the native, plants of the Merritt Island Wildlife Refuge, the citrus production, or the beekeeping industry of the island. Conditions were simulated in greenhouse exposure chambers and field chambers constructed to model the ideal continuous stirred tank reactor. A plant exposure system was developed for dispensing and monitoring the two major chemicals in SRF exhaust, HCl and Al203, and for dispensing and monitoring SRF exhaust (controlled fuel burns). Plants native to Merritt Island, Florida were grown and used as test species. Dose-response relationships were determined for short term exposure of selected plant species to HCl, Al203, and mixtures of the two to SRF exhaust.

  9. Reusable Solid Rocket Motor - Accomplishment, Lessons, and a Culture of Success

    Moore, D. R.; Phelps, W. J.


    The Reusable Solid Rocket Motor (RSRM) represents the largest solid rocket motor (SRM) ever flown and the only human-rated solid motor. High reliability of the RSRM has been the result of challenges addressed and lessons learned. Advancements have resulted by applying attention to process control, testing, and postflight through timely and thorough communication in dealing with all issues. A structured and disciplined approach was taken to identify and disposition all concerns. Careful consideration and application of alternate opinions was embraced. Focus was placed on process control, ground test programs, and postflight assessment. Process control is mandatory for an SRM, because an acceptance test of the delivered product is not feasible. The RSRM maintained both full-scale and subscale test articles, which enabled continuous improvement of design and evaluation of process control and material behavior. Additionally RSRM reliability was achieved through attention to detail in post flight assessment to observe any shift in performance. The postflight analysis and inspections provided invaluable reliability data as it enables observation of actual flight performance, most of which would not be available if the motors were not recovered. RSRM reusability offered unique opportunities to learn about the hardware. NASA is moving forward with the Space Launch System that incorporates propulsion systems that takes advantage of the heritage Shuttle and Ares solid motor programs. These unique challenges, features of the RSRM, materials and manufacturing issues, and design improvements will be discussed in the paper.

  10. Structural and mechanical design challenges of space shuttle solid rocket boosters separation and recovery subsystems

    Woodis, W. R.; Runkle, R. E.


    The design of the space shuttle solid rocket booster (SRB) subsystems for reuse posed some unique and challenging design considerations. The separation of the SRBs from the cluster (orbiter and external tank) at 150,000 ft when the orbiter engines are running at full thrust meant the two SRBs had to have positive separation forces pushing them away. At the same instant, the large attachments that had reacted launch loads of 7.5 million pounds thrust had to be servered. These design considerations dictated the design requirements for the pyrotechnics and separation rocket motors. The recovery and reuse of the two SRBs meant they had to be safely lowered to the ocean, remain afloat, and be owed back to shore. In general, both the pyrotechnic and recovery subsystems have met or exceeded design requirements. In twelve vehicles, there has only been one instance where the pyrotechnic system has failed to function properly.

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

    Palaszewski, Bryan


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

  12. Development of a Polyuretbane Binder System Giving a 'Knottable' Composite Solid Propellant

    E. Devadoss


    Full Text Available Polyurethane propellants, which constitute one of the 'work-horse' binder systems in modern solid rocketry are easily amenable for tailoring the mechanical properties in terms of variations in the molecular structure of the backbone polyols, the isocyanates and stoichiometry of the reactants. The paper deals with studies in developing an advanced binder system based on poly (oxy propylene glycol and toluene-di-isocyanate, which is capable of accommodating high solids loading and conceding elongatlon at maximum stress, of more than 125 per cent. The gum-stock properties of the binder are related with those of a low molecular weight version and the results are explained based on the network theory of condensation polymers.

  13. Rocket noise - A review

    McInerny, S. A.


    This paper reviews what is known about far-field rocket noise from the controlled studies of the late 1950s and 1960s and from launch data. The peak dimensionless frequency, the dependence of overall sound power on exhaust parameters, and the directivity of the overall sound power of rockets are compared to those of subsonic jets and turbo-jets. The location of the dominant sound source in the rocket exhaust plume and the mean flow velocity in this region are discussed and shown to provide a qualitative explanation for the low peak Strouhal number, fD(e)/V(e), and large angle of maximum directivity. Lastly, two empirical prediction methods are compared with data from launches of a Titan family vehicle (two, solid rocket motors of 5.7 x 10 to the 6th N thrust each) and the Saturn V (five, liquid oxygen/rocket propellant engines of 6.7 x 10 to the 6th N thrust, each). The agreement is favorable. In contrast, these methods appear to overpredict the far-field sound pressure levels generated by the Space Shuttle.

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

    Kapp, J. R.


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

  15. Propelling Ariane. The Vulcain engines and the solid propellant engines; Propulser Ariane. Les moteurs Vulcain et les moteurs a propergol solide



    The development of the Vulcain program was ensured thanks to a European cooperation with an ESA (European Space Agency) financing. The CNES (European Centre for Space Studies) has ensured the technical and financial direction of the program and gave the control of the development to the SEP. The manufacturing of the Vulcain engine is managed under the Arianespace contract, in charge of the marketing of the Ariane 5 launcher. The overall engineering of the engine and its tests were carried out by the SEP in Vernon (France) and Lampoldshausen (Germany) test facilities. SEP has also developed and built the hydrogen turbo-pump, the gas generator and its feeding valves. Several companies are involved in the development of this engine: DASA for the combustion chamber, Fiat Avio for the oxygen turbo-pump, Volvo Aero Corp. for the divergent and the hydrogen and oxygen turbines, Techspace Aero for the chamber injection valves and the drain and hot gases valves, Microtechnica for the electro-valves and check valves, SPE for the firing and start-up equipments, Avica for the feeding lines, Devtec for the supports, and MAN for the cardan and the thermal protection. This paper describes the functioning principle of the Vulcain engine, and of the two solid propellant fuel acceleration stages of the Ariane 5 launcher. Some future projects of the SEP are also described: the dual liquid-fuels engine and the plasma engine. (J.S.)

  16. A Flight Investigation of the Damping in Roll and Rolling Effectiveness Including Aeroelastic Effects of Rocket Propelled Missile Models Having Cruciform, Triangular, Interdigitated Wings and Tails

    Hopko, R. N.


    The damping in roll and rolling effectiveness of two models of a missile having cruciform, triangular, interdigitated wings and tails have been determined through a Mach number range of 0.8 to 1.8 by utilizing rocket-propelled test vehicles. Results indicate that the damping in roll was relatively constant over the Mach umber range investigated. The rolling effectiveness was essentially constant at low supersonic speeds and increased with increasing mach numbers in excess of 1.4 over the Mach number range investigated. Aeroelastic effects increase the rolling-effectiveness parameters pb/2V divided by delta and decrease both the rolling-moment coefficient due to wing deflection and the damping-in-roll coefficient.

  17. Process Study on Disposal Solid Composite Propellant in Applying Critical Fluid Extraction Technology%用临界液萃取技术处理复合固体推进剂的工艺研究

    高兴勇; 巩永校; 杜仕国; 刘国庆


    Foreign country has been successful in applying critical fluid extraction technology for demilitarization of solid propellant and explosive. This paper introduced the process of critical fluid extraction technology through demilitarization of rocket motors containing ammonium percholrate (AP) composite propellants. The demilitarization process consists of four-step. Step one involves removing the AP propellant from the rocket motor. Step two extracts AP and separates the AP/liquid ammonia solution from binder residue. Step three recovers the AP by evaporating the ammonia. Step four condenses the ammonia vapor and recycles the liquid ammonia for a continuous removal/extraction operation.%通过对火箭发动机含高氯酸铵(AP)的复合推进剂用临界液氨销毁来介绍这一工艺方法。销毁工艺由四个连续的步骤构成:第一步采用临界液氨把含AP的复合推进剂从火箭发动机中取出;第二步萃取AP并将含AP的液氨和粘结剂残渣分离开;第三步从氨水中蒸发出AP;第四步将氨蒸气冷凝并重新循环使用。

  18. The effects of various cure cycles upon the viability of Bacillus subtillis var. niger spores within solid propellant.

    Brewer, W. A.; Paik, W. W.; Robillard, C. L.; Green, R. H.; Smith, C. D.


    Saturethane solid propellant was used in all tests. The spore inoculum was evenly distributed in the propellant. Samples weighing approximately 5 g were aseptically removed, placed into curing ovens, and exposed to cure temperatures. Initial tests were conducted at 82 and 93 C. Analysis of the obtained data indicated that the survivor curves were not linear. Exposure of the inoculated propellant samples to 93 C reduced the initial population to less than 0.01% in about 20 hours. At 82 C, approximately 168 hours were required for a similar reduction. Tests involving curing temperatures of 105 and 115 C were also conducted. It is pointed out that changes in the mechanism of spore inactivation due to chemical and physical changes in the propellant could account for the nonlinear survivor curves obtained.

  19. A Computer Program for the Prediction of Solid Propellant Rocket Motor Performance. Volume 3


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  20. Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method

    Hueser, J. E.; Brock, F. J.; Melfi, L. T., Jr.; Bird, G. A.


    A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).

  1. The Fire Environment of a Solid Rocket Propellant Burning in Air


    Radiant flux from agglomrates per unit area per unit tim8 reaching the cawra pps Ptctures per secoed Q Film exposure per untL area from smoke end an...clear how the temperature there could be related to the free strem temperature, even if the thermocouples survived. D-3 Lmmrsion Calor mter In m

  2. Extended temperature range rocket injector

    Schneider, Steven J. (Inventor)


    A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connnected to valves between these manifolds and propellant storage tanks. When cryogenic propellant temperatures are sensed, only a portion of the valves are opened to furnish propellants to some of the manifolds. When lower temperatures are sensed, additional valves are opened to furnish propellants to more of the manifolds.

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


    Badakhshan A1 , Danczyk S. A.2, Wirth D.3 and Pilon L. 3 Abstract We have studied the ignition of fuel sprays and simulated solid rocket fuels (SRF...photoignition of solid oxidizer/CNT mixtures exposed to a flash of light. The flash source was a commercial studio flash lamp with a rated maximum

  4. Characterization of rocket propellant combustion products. Chemical characterization and computer modeling of the exhaust products from four propellant formulations: Final report, September 23, 1987--April 1, 1990

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


    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.

  5. Migration of a moonlet in a ring of solid particles : Theory and application to Saturn's propellers

    Crida, A; Rein, H; Charnoz, S; Salmon, J


    Hundred meter sized objects have been identified by the Cassini spacecraft in Saturn's A ring through the so-called "propeller" features they create in the ring. These moonlets should migrate, due to their gravitational interaction with the ring ; in fact, some orbital variation have been detected. The standard theory of type I migration of planets in protoplanetary disks can't be applied to the ring system, as it is pressureless. Thus, we compute the differential torque felt by a moonlet embedded in a two-dimensional disk of solid particles, with flat surface density profile, both analytically and numerically. We find that the corresponding migration rate is too small to explain the observed variations of the propeller's orbit in Saturn's A-ring. However, local density fluctuations (due to gravity wakes in the marginally gravitationally stable A-ring) may exert a stochastic torque on a moonlet. Our simulations show that this torque can be large enough to account for the observations, depending on the paramet...

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

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


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

  7. Diffusion of plasticizer in a solid propellant based on hydroxyl-terminated polybutadiene

    Juliano Libardi


    Full Text Available The aim of this work was to determine the dioctyl azelate (DOZ plasticizer diffusion coefficient (D for samples containing the interfaces of rubber, liner and solid composite propellant based on hydroxyl-terminated polybutadiene (HTPB. The samples used in the diffusion study were aged up to 31 days after the cure period at 80 °C. A computer program implementing a mathematical model of Fick's second Law of diffusion was developed to calculate the diffusion coefficient based on concentration data obtained from gas chromatographic analyses. The effects of the diffusion phenomenon were also investigated by Shore A hardness and scanning electron microscope (SEM techniques. These analyses were carried out using samples aged at room temperature and at 80 °C. The hardness results showed an increasing trend for the samples aged at room temperature; however in the tests carried out at 80 °C they showed the opposite trend. The SEM analyses detected meaningful changes in the surface morphology of the propellant for both aging temperatures.

  8. Modeling and simulation of the debonding process of composite solid propellants

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


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

  9. Propellants: the feasibility of their manufacture in India

    R. K. Srivastava


    Full Text Available In the light of the recent Chinese aggression, the modernization of our defence forces has become a vital necessity. In this paper, one aspect of this problem has been considered in detail viz. the development of an indigenous rocket and missile force. While it is true that many factors e.g. rocket motors, propellants, guidance systems etc. are involved, it is also true that a start in one area will act as an impetus to developments in the other fields. Solid/liquid propellant and oxidizer systems have been considered, the properties of solid and liquid propellants evaluated and on the basis of such comparison, it has been concluded that effort concentrated on the development of liquid propellants will be well expended. Liquid propellant/oxidizer systems have been compared amongst themselves and it has been concluded that the hydrazine fuels oxidized by RFNA/WFNA/IRFNA would represent systems fulfilling the country's immediate military needs best. The availability of raw materials for the manufacture of hydrazine fuels (and also of some solid propellants has been considered and it is shown that the necessary raw materials are available in sufficient quantities to support an indigenous propellants industry.

  10. Application of solid hydrocarbon propellant on turbocharged solid propellant ramjet%固体碳氢推进剂在涡轮增压固冲发动机中的应用

    李江; 刘凯; 王伟; 刘洋; 田园


    提出了固体碳氢推进剂作为涡轮增压固冲发动机( TSPR)驱涡推进剂的方案,分析了适用于TSPR推进剂的热力参数和一次燃烧产物成分,完成了驱涡推进剂的选择;进行了备选推进剂( CH04)对TSPR性能的影响性评估,证明该推进剂能够满足TSPR的性能要求;对所选推进剂了进行了一、二次燃烧试验,试验结果表明,CH04推进剂在补燃室点火较困难,但其一次、二次燃烧稳定性好,燃气参数基本满足TSPR对推进剂性能要求.%Solid hydrocarbon propellant was proposed to be used to drive turbine of Turbocharged Solid Propellant Ramjet (TSPR).The thermodynamic parameters and the first combustion production of propellant used for TSPR were analyzed to select the suitable propellant(CH04) to drive the turbine. An evaluation for propellant' s influence on the performance of TSPR was per-formed,which proved that the propellant(CH04)meets the requirements of TSPR.The first and second combustion experiments were carried out.The results show that,the first and second combustion have high stability,despite of the difficulty in ignition in the after-burner,and the feasibility of the selected propellant used for TSPR was verified.

  11. Characterization and Fate of Gun and Rocket Propellant Residues on Testing and Training Ranges: Interim Report 1


    the gun barrel liners. Examples include wax , talc, and titanium dioxide. Grain size and shape The properties of the propellant are greatly...phthalate 2,4-dinitrotoluene nitroglycerin ethyl centralite triacetin 2,6-dinitrotoluene nitroguanidine akardite wax ethyl centralite diethylene...The extracts were filtered by passing each through a Millex-FH PTFE 0.45 syringe filter (Millipore Corp.). For reverse-phase high- performance liquid

  12. Energetic Abilities of Solid Composite Propellants Based on 3,4,5-Trinitropyrazole and Ammonium Dinitramide

    LEMPERT David B; CHUKANOV Nikita; SHU Yuan-jie


    The investigation aims at the expansion of the basis of formulations of solid composite propellants by introducing new compositions with lower sensitivity to mechanic impact and improved thermal stability .The formulations based on trinitropyrazole (TNP) contains a binder (a hydrocarbon or active one) ,aluminum and inorganic oxidizer ADN .The results show that a binary for‐mulation TNP + active binder (18% -19% )(volume fraction) with no metal is well designed which would achieve high specific im‐pulse (at Pc∶ Pa=40∶1) of 248 s ,high density of 1 .80 g/cm3 and combustion temperature Tc about 3450 K .In terms of energy , metal‐free compositions with TNP lose a bit to those with HMX ,only if HMX fraction in formulation is higher than 45% -50% .

  13. A functional on-chip pressure generator using solid chemical propellant for disposable lab-on-a-chip.

    Hong, Chien-Chong; Murugesan, Suresh; Kim, Sanghyo; Beaucage, Gregory; Choi, Jin-Woo; Ahn, Chong H


    This paper presents a functional on-chip pressure generator that utilizes chemical energy from a solid chemical propellant to perform fluidic delivery in applications of plastic-based disposable biochips or lab-on-a-chip systems. In this functional on-chip pressure generator, azobis-isobutyronitrile (AIBN) as the solid chemical propellant is deposited on a microheater using a screen-printing technique, which can heat the AIBN at 70 degrees C to produce nitrogen gas. The output pressure of nitrogen gas, generated from the solid chemical propellant, is adjustable to a desired pressure by controlling the input power of the heater. Using this chemical energy source, the generated pressure depends on the deposited amount of the solid chemical propellant and the temperature of the microheater. Experimental measurements show that this functional on-chip pressure generator can achieve around 3 000 Pa pressure when 189 mJ of energy is applied to heat the 100 microg of AIBN. This pressure can drive 50 nl of water through a microfluidic channel of 70 mm and cross-sectional area of 100 microm x 50 microm. Due to its compact size, ease of fabrication and integration, high reliability (no moving parts), biologically inert gas output along with functionality of gas generation, this pressure generator will be an excellent pressure source for handling the fluids of disposable lab-on-a-chip, biochemical analysis systems or drug delivery systems.

  14. Effect of Temperature Sensitivity and Plasticizer Diffusive Transport on Performance of Layered Solid Propellants under Electrothermal Plasma Injection

    Bourham, Mohamed


    .... The plasma jet velocity 2 inches from the source exit was found to be about 1300 m/s. Following characterization of the plasma-flow field, a set of experiments was conducted on JA-2 solid propellant with controlled bed temperature...

  15. Internal Ballistic Code for Solid Rocket Motors using Minimum Distance Function for Grain Burnback

    Afroz Javed


    Full Text Available A computer code has been developed for internal ballistic performance evaluation of solid rocket motors, using minimum distance function (MDF approach for prediction of geometry evolution. This method can handle any complex geometry without the need to define different geometrical shapes and their evolution as used in several existing analytical geometry evolution-based methodologies. The code is validated with both experimental results published in literature, as well as for solid rocket motors of tactical and strategic missiles and a very good match is obtained with static test results. The output of the code gives p-t (pressure-time curve as well as the detailed parameters of the flow along the axial direction, and geometries in the form of mesh file, which can be further used as input to codes for CFD analysis.Defence Science Journal, Vol. 65, No. 3, May 2015, pp.181-188, DOI:

  16. A Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Yang, H. Q.; West, Jeff


    A capability to couple NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This paper summarizes the efforts in applying the installed coupling software to demonstrate/investigate fluid-structure interaction (FSI) between pressure wave and flexible inhibitor inside reusable solid rocket motor (RSRM). First a unified governing equation for both fluid and structure is presented, then an Eulerian-Lagrangian framework is described to satisfy the interfacial continuity requirements. The features of fluid solver, Loci/CHEM and structural solver, CoBi, are discussed before the coupling methodology of the solvers is described. The simulation uses production level CFD LES turbulence model with a grid resolution of 80 million cells. The flexible inhibitor is modeled with full 3D shell elements. Verifications against analytical solutions of structural model under steady uniform pressure condition and under dynamic condition of modal analysis show excellent agreements in terms of displacement distribution and eigen modal frequencies. The preliminary coupled result shows that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor.

  17. Simulation of reactive polydisperse sprays strongly coupled to unsteady flows in solid rocket motors: Efficient strategy using Eulerian Multi-Fluid methods

    Sibra, A.; Dupays, J.; Murrone, A.; Laurent, F.; Massot, M.


    In this paper, we tackle the issue of the accurate simulation of evaporating and reactive polydisperse sprays strongly coupled to unsteady gaseous flows. In solid propulsion, aluminum particles are included in the propellant to improve the global performances but the distributed combustion of these droplets in the chamber is suspected to be a driving mechanism of hydrodynamic and acoustic instabilities. The faithful prediction of two-phase interactions is a determining step for future solid rocket motor optimization. When looking at saving computational ressources as required for industrial applications, performing reliable simulations of two-phase flow instabilities appears as a challenge for both modeling and scientific computing. The size polydispersity, which conditions the droplet dynamics, is a key parameter that has to be accounted for. For moderately dense sprays, a kinetic approach based on a statistical point of view is particularly appropriate. The spray is described by a number density function and its evolution follows a Williams-Boltzmann transport equation. To solve it, we use Eulerian Multi-Fluid methods, based on a continuous discretization of the size phase space into sections, which offer an accurate treatment of the polydispersion. The objective of this paper is threefold: first to derive a new Two Size Moment Multi-Fluid model that is able to tackle evaporating polydisperse sprays at low cost while accurately describing the main driving mechanisms, second to develop a dedicated evaporation scheme to treat simultaneously mass, moment and energy exchanges with the gas and between the sections. Finally, to design a time splitting operator strategy respecting both reactive two-phase flow physics and cost/accuracy ratio required for industrial computations. Using a research code, we provide 0D validations of the new scheme before assessing the splitting technique's ability on a reference two-phase flow acoustic case. Implemented in the industrial

  18. Coupled Solid Rocket Motor Ballistics and Trajectory Modeling for Higher Fidelity Launch Vehicle Design

    Ables, Brett


    Multi-stage launch vehicles with solid rocket motors (SRMs) face design optimization challenges, especially when the mission scope changes frequently. Significant performance benefits can be realized if the solid rocket motors are optimized to the changing requirements. While SRMs represent a fixed performance at launch, rapid design iterations enable flexibility at design time, yielding significant performance gains. The streamlining and integration of SRM design and analysis can be achieved with improved analysis tools. While powerful and versatile, the Solid Performance Program (SPP) is not conducive to rapid design iteration. Performing a design iteration with SPP and a trajectory solver is a labor intensive process. To enable a better workflow, SPP, the Program to Optimize Simulated Trajectories (POST), and the interfaces between them have been improved and automated, and a graphical user interface (GUI) has been developed. The GUI enables real-time visual feedback of grain and nozzle design inputs, enforces parameter dependencies, removes redundancies, and simplifies manipulation of SPP and POST's numerous options. Automating the analysis also simplifies batch analyses and trade studies. Finally, the GUI provides post-processing, visualization, and comparison of results. Wrapping legacy high-fidelity analysis codes with modern software provides the improved interface necessary to enable rapid coupled SRM ballistics and vehicle trajectory analysis. Low cost trade studies demonstrate the sensitivities of flight performance metrics to propulsion characteristics. Incorporating high fidelity analysis from SPP into vehicle design reduces performance margins and improves reliability. By flying an SRM designed with the same assumptions as the rest of the vehicle, accurate comparisons can be made between competing architectures. In summary, this flexible workflow is a critical component to designing a versatile launch vehicle model that can accommodate a volatile

  19. Formulation and Testing of Paraffin-Based Solid Fuels Containing Energetic Additives for Hybrid Rockets

    Larson, Daniel B.; Boyer, Eric; Wachs,Trevor; Kuo, Kenneth K.; Story, George


    Many approaches have been considered in an effort to improve the regression rate of solid fuels for hybrid rocket applications. One promising method is to use a fuel with a fast burning rate such as paraffin wax; however, additional performance increases to the fuel regression rate are necessary to make the fuel a viable candidate to replace current launch propulsion systems. The addition of energetic and/or nano-sized particles is one way to increase mass-burning rates of the solid fuels and increase the overall performance of the hybrid rocket motor.1,2 Several paraffin-based fuel grains with various energetic additives (e.g., lithium aluminum hydride (LiAlH4) have been cast in an attempt to improve regression rates. There are two major advantages to introducing LiAlH4 additive into the solid fuel matrix: 1) the increased characteristic velocity, 2) decreased dependency of Isp on oxidizer-to-fuel ratio. The testing and characterization of these solid-fuel grains have shown that continued work is necessary to eliminate unburned/unreacted fuel in downstream sections of the test apparatus.3 Changes to the fuel matrix include higher melting point wax and smaller energetic additive particles. The reduction in particle size through various methods can result in more homogeneous grain structure. The higher melting point wax can serve to reduce the melt-layer thickness, allowing the LiAlH4 particles to react closer to the burning surface, thus increasing the heat feedback rate and fuel regression rate. In addition to the formulation of LiAlH4 and paraffin wax solid-fuel grains, liquid additives of triethylaluminum and diisobutylaluminum hydride will be included in this study. Another promising fuel formulation consideration is to incorporate a small percentage of RDX as an additive to paraffin. A novel casting technique will be used by dissolving RDX in a solvent to crystallize the energetic additive. After dissolving the RDX in a solvent chosen for its compatibility

  20. Raman Spectroscopy for Instantaneous Multipoint, Multispecies Gas Concentration and Temperature Measurements in Rocket Engine Propellant Injector Flows

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc


    Propellant injector development at MSFC includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellant mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

  1. Program ELM: A tool for rapid thermal-hydraulic analysis of solid-core nuclear rocket fuel elements

    Walton, James T.


    This report reviews the state of the art of thermal-hydraulic analysis codes and presents a new code, Program ELM, for analysis of fuel elements. ELM is a concise computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in a nuclear thermal rocket reactor with axial coolant passages. The program was developed as a tool to swiftly evaluate various heat transfer coefficient and friction factor correlations generated for turbulent pipe flow with heat addition which have been used in previous programs. Thus, a consistent comparison of these correlations was performed, as well as a comparison with data from the NRX reactor experiments from the Nuclear Engine for Rocket Vehicle Applications (NERVA) project. This report describes the ELM Program algorithm, input/output, and validation efforts and provides a listing of the code.

  2. Investigation on plume interference effect of solid propellant micro-thruster

    ZHANG Bin; MAO Gen-wang; HU Song-qi; CHEN Mao-lin


    The three-dimensional numerical simulation of two-phase plume flow of solid propellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Monte Carlo(DSMC) method for multi-component gas flow with the two-way coupling model for two-phase rarefied flow.At different space between the two micro-thrusters and different wall temperature,the plume interference effect was analyzed specifically.The results show that under the plume interference effect the gas is compressed and the flow direction is changed,which resulted in the increasing of gas pressure and temperature;solid phase made no significant effect on the flow parameters of gas phase;with the rising of the space between the two micro-thrusters,the maximum pressure decreased and the maximum temperature increased in the domain under the plume interference effect;the wall temperature could influence the temperature of the gas which is extremely close to the wall,but not the gas pressure.

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



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

  4. Research on Mechanical Property Prediction of Solid Propellant Base on GA-BP Neural Network%固体推进剂力学性能预估研究

    李进贤; 莫文宾; 唐金兰


    固体火箭发动机中,药柱的结构完整性直接关系到发动机的结构完整性和可靠性,而推进剂的力学性能对保持药柱结构完整性起着重要作用,也是决定推进剂寿命的重要指标.为了预估固体推进剂的力学性能,提高系统的可靠性,将遗传算法和神经网络相结合,建立了预估固体推进剂力学性能的遗传神经网络(GA-BP)模型.利用模型预测了某固体推进剂在不同温度、湿度和时间下的抗拉强度、延伸率、弹性模量变化情况,并与试验结果进行了比较.结果表明,模型预估精度高,泛化能力强,仿真计算与试验在结果上有很好的一致性.从而为固体火箭发动机的结构完整性研究提供可靠依据.%Solid propellant grain structural integrity influences the structural integrity and reliability of solid rocket motor (SRM). Mechanical property of solid propellant plays an important role in grain structural integrity, which is critical criterion of solid propellant life. In order to predict mechanical property of solid propellant, a new mechanical property prediction model for solid propollant was established by means of combination of 8enetic algorithm with neural network (GA-BP). Using above model, the mechanical proporty of a solid propellant in conditions of different ternperature, humidity and time was predicted and compared with experiment results. The comparison results show high precision of the model and strong ability of generalization and with good consistency between prediction of model and experiment. The investigation provides reliable assistance for structural integrity research of SRM.

  5. Convective Heat Transfer in the Reusable Solid Rocket Motor of the Space Transportation System

    Ahmad, Rashid A.; Cash, Stephen F. (Technical Monitor)


    This simulation involved a two-dimensional axisymmetric model of a full motor initial grain of the Reusable Solid Rocket Motor (RSRM) of the Space Transportation System (STS). It was conducted with CFD (computational fluid dynamics) commercial code FLUENT. This analysis was performed to: a) maintain continuity with most related previous analyses, b) serve as a non-vectored baseline for any three-dimensional vectored nozzles, c) provide a relatively simple application and checkout for various CFD solution schemes, grid sensitivity studies, turbulence modeling and heat transfer, and d) calculate nozzle convective heat transfer coefficients. The accuracy of the present results and the selection of the numerical schemes and turbulence models were based on matching the rocket ballistic predictions of mass flow rate, head end pressure, vacuum thrust and specific impulse, and measured chamber pressure drop. Matching these ballistic predictions was found to be good. This study was limited to convective heat transfer and the results compared favorably with existing theory. On the other hand, qualitative comparison with backed-out data of the ratio of the convective heat transfer coefficient to the specific heat at constant pressure was made in a relative manner. This backed-out data was devised to match nozzle erosion that was a result of heat transfer (convective, radiative and conductive), chemical (transpirating), and mechanical (shear and particle impingement forces) effects combined.

  6. Ozone depletion in the plume of a solid-fuelled rocket

    B. C. Krüger

    Full Text Available The local effects of the emission of a solid-fuelled rocket on the stratospheric ozone concentration have been investigated by photochemical model calculations. A one-dimensional horizontal model has been applied which calculates the trace gas composition at a single atmospheric altitude spatially resolved around the exhaust plume. Different cases were tested for the emissions of the Space Shuttle concerning the composition of the exhaust and the effects of heterogeneous reactions on atmospheric background aerosol.

    The strongest depletion of ozone is achieved when a high amount of the emitted chlorine is Cl2. If it is purely HCl, the effect is smallest, though in this case the heterogeneous reactions show their largest influence. From the results it may be estimated whether ozone depletion caused by rocket launches can be detected by satellite instruments. It appears that the chance of coincidental detection of such an event is rather small.

  7. 3D flash lidar performance in flight testing on the Morpheus autonomous, rocket-propelled lander to a lunar-like hazard field

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.


    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-σ. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  8. 3-D Flash Lidar Performance in Flight Testing on the Morpheus Autonomous, Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.


    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-delta. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  9. Pressure oscillations and instability of working processes in the combustion chambers of solid rocket motors

    Emelyanov, V. N.; Teterina, I. V.; Volkov, K. N.; Garkushev, A. U.


    Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

  10. Optimization of the stand for test of hybrid rocket engines of solid fuel

    Zolotorev Nikolay


    Full Text Available In the paper the laboratory experimental stand of the hybrid rocket engine of solid fuel to study ballistic parameters of the engine at burning of high-energy materials in flow of hot gas is presented. Mixture of air with nitrogen with a specified content of active oxygen is used as a gaseous oxidizer. The experimental stand has modular design and consists of system of gas supply, system of heating of gas, system for monitoring gas parameters, to which a load cell with a model engine was connected. The modular design of the stand allows to change its configuration under specific objective. This experimental stand allows to conduct a wide range of the pilot studies at interaction of a hot stream of gas with samples high-energy materials.

  11. Numerical techniques for solving nonlinear instability problems in smokeless tactical solid rocket motors. [finite difference technique

    Baum, J. D.; Levine, J. N.


    The selection of a satisfactory numerical method for calculating the propagation of steep fronted shock life waveforms in a solid rocket motor combustion chamber is discussed. A number of different numerical schemes were evaluated by comparing the results obtained for three problems: the shock tube problems; the linear wave equation, and nonlinear wave propagation in a closed tube. The most promising method--a combination of the Lax-Wendroff, Hybrid and Artificial Compression techniques, was incorporated into an existing nonlinear instability program. The capability of the modified program to treat steep fronted wave instabilities in low smoke tactical motors was verified by solving a number of motor test cases with disturbance amplitudes as high as 80% of the mean pressure.

  12. Numerical Simulation of a Dual Pulse Solid Rocket Motor Flow Field

    Afroz Javed


    Full Text Available Numerical simulations are carried out for the internal flow field of a dual pulse solid rocket motor port to understand the flow behaviour. Three dimensional Reynolds Averaged Navier Stokes equations are solved alongwith shear stress transport turbulence model using commercial code. The combustion gas is assumed as a mixture of alumina and gases and single phase flow calculations are done with the thermo chemical properties provided for the mixture. The simulation captures all the essential features of the flow field. The flow accelerates through the pulse separation device (PSD port and high temperature and high velocity gas is seen to impinge the motor wall near the PSD port. The overall total pressure drop through motor port and through PSD is found to be moderate.Defence Science Journal, 2012, 62(6, pp.369-374, DOI:

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

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


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

  14. Effect of Oxidizer Particle Size on Burning Rate and Thermal Decomposition of Composite Solid Propellants

    K. Kishore


    Full Text Available Studies on Thermal decomposition of ammonium perchlorarte(AP- polystyrene(PS propellant and burning rate of PS/AP propellant have been carried out as a function of oxidizer particle size. Thermal decomposition of AP and AP/PS propellant as function of AP particle size shows a maximum rate around 100 micro particle size which has been explained on the basis of Mample's theory. No such maximum is observed in the case of PS/AP propellant burning rate.

  15. Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

    Bellman, Donald R; Humphrey, Jack C


    Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

  16. Design and optimization of micro-semiconductor bridge used for solid propellant microthrusters array

    Ru, Chengbo; Dai, Ji; Xu, Jianbing; Ye, Yinghua; Zhu, Peng; Shen, Ruiqi


    Igniter is the basic component of MEMS-based solid propellant microthrusters (SPM) array, which should response fast with low dissipated energy. To satisfy the requirements, micro-semiconductor bridge (MSCB) with dual V-angles contributing to the reduction of dissipated energy is introduced into the array. The electrical explosion characteristics of MSCB investigated under constant voltage, in the considerations of the limit of power supply in the micro/nanosatellite, are similar to those under capacitive discharge. The bridge was completely vaporized and burst into the hot plasma under high-level voltage. While under low-level voltage, the bridge was partly vaporized without detected plasma. The 3#SCB (90° V-angle) with the smallest power density resulting in the lowest applied voltage (12.0 V) and the 2#SCB with the smallest narrow width (Wmin) leading to the lowest average power (11.3 W) were chosen to the further optimization. The ignition tests were implemented to examine the ignition reliability of the contractible igniters, which were optimized to satisfy the voltage limitation with sufficient ignition capacity. The results show that the effective SCB with lowest dissipated power is the 3#-3SCB, which function within 100 μ under 7.0 V, and the average power below 5.0 W.

  17. Characterization of aluminum/RP-1 gel propellant properties

    Rapp, Douglas C.; Zurawski, Robert L.


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

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


    76 mm ID, 95.2 mm OD) with welded end-plugs 450 mm apart. The volume of the vessel is approximately 2 liters. Nickel- chrome heating wire is wound...representation of the NWC ESD RC discharge test apparatus. VOLTAGE PROBE GC SHEATH ON O OAAG , * E W. O’ 5RASS ELECTRODE CONO U E PALA IN OR -’-’" i

  19. Characterization of nal powders for rocket propulsion

    Merotto, L.; Galfetti, L.; Colombo, G.; DeLuca, L. T.


    Nanosized metal powders are known to significantly improve both solid and hybrid rocket performance, but have some drawbacks in terms of cost, safety, and possible influence on propellant mechanical properties. Performance enhancement through nanosized metal or metal hydride addition to solid fuels is currently under investigation also for hybrid propulsion. Therefore, a preburning characterization of the powders used in solid propellant or fuel manufacturing is useful to assess their effects on the ballistic properties and engine performance. An investigation concerning the comparative characterization of several aluminum powders having different particle size, age, and coating is presented. Surface area, morphology, chemical species concentration and characteristics, surface passivation layers, surface and subsurface chemical composition, ignition temperature and ignition delay are investigated. The aim of this characterization is to experimentally assess the effect of the nAl powder properties on ballistic characteristics of solid fuels and solidrocket composite-propellant performance, showing an increase in terms of Is caused by the decrease of two-phase losses in solid and a possible significant rf increase in hybrid rockets.

  20. Method of Comparative Analysis of Highly Dispersed Condensed Fuel Combustion Efficiency in Arbitrary Geometry Solid Propellant Ramjet Burners

    A. V. Voroneckii


    Full Text Available The paper deals with various theoretical approaches to the mathematical modeling of the operating process in solid propellant ramjets (SPRJ that use highly metalized solid propellant. It introduces a new method (combustion operating law method that allows us to carry out comparative analysis of combustion efficiency in SPRJ arbitrary geometry ram-burners (RB when there is no accurate information on the combustion law of condensed fuel particles. To illustrate an application of the proposed method, mathematical modeling of the operating process was conducted for three SPRJ ram-burners with three different air intakes (AI, for which distribution fields of main parameters of gas and fuel particles have been obtained. Most complete combustion of fuel particles and the lowest level of particles buildup are registered for RB180 (180 degree angle between AIs. The results of a comparative analysis show that the relative (compared to RB180 efficiency of the particle burning process equals 0.64 and 0.6, respectively, for RB90 (90 degree angle between AIs and RB60 (60 degree angle between AIs. The proposed method may be applied to solve the most difficult problems of mathematical modeling when the optimization development of the solid propellant and ramjet structure are fulfilled simultaneously, i.e. when designers do not have the complete information about the combustion law of the condensed fuel particles.

  1. Arc Jet Test and Analysis of Asbestos Free Solid Rocket Motor Nozzle Dome Ablative Materials

    Clayton, J. Louie


    Asbestos free solid motor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.


    李卓; 徐秉业; 蒲建军


    In this paper ,viscoelastic property of solid rocket propellant is described by fractional derivative model.Frequency response function and impulse response function of one dimensional pole is deduced by means of Laplace transformation and its inverse transformation.The proportions of pole part and branchcut part in the response are analyzed.The transitional and steady response of the pole is calculated under harmonic excitation and the steady acceleration response is compared with expeimental results.%本文用三参数分数阶导数模型描述了固体推进剂的粘弹特性,用Laplace变换和反变换求解一维杆件的频响函数和脉冲响应函数,分析了极点部分和截断部分在响应中所占比例,计算了在简谐激励下杆件的瞬态和稳态响应,并就加速度稳态响应部分与试验做了对比。

  3. High temperature reformation of aluminum and chlorine compounds behind the Mach disk of a solid-fuel rocket exhaust

    Park, C.


    Chemical reactions expected to occur among the constituents of solid-fuel rocket engine effluents in the hot region behind a Mach disk are analyzed theoretically. With the use of a rocket plume model that assumes the flow to be separated in the base region, and a chemical reaction scheme that includes evaporation of alumina and the associated reactions of 17 gas species, the reformation of the effluent is calculated. It is shown that AlClO and AlOH are produced in exchange for a corresponding reduction in the amounts of HCl and Al2O3. For the case of the space shuttle booster engines, up to 2% of the original mass of the rocket fuel can possibly be converted to these two new species and deposited in the atmosphere between the altitudes of 10 and 40 km. No adverse effects on the atmospheric environment are anticipated with the addition of these two new species.

  4. Laser transmittance measurement technology on plume from solid-propellant engine%固体发动机羽烟的激光透过率测试

    陈胜石; 宁子立; 欧秋晔; 陈文胜; 纪明; 高教波; 薛常佳


    为获得羽烟对激光透过率的影响,用烟箱法对2种配方的缩比发动机羽烟在1.06 μm、10.6 μm激光波段的透过率进行测试.采用1.064 μm激光调制发射、接收、数据采集系统对1.06 μm激光波段烟雾透过率测试;用黑体、光谱辐射计、数据采集系统可测出2 μm ~13 μm 的光学透过率,从中提出10.6 μm激光波段烟雾透过率,得到不同推进剂配方、不同烟雾浓度情况下10.6 μm光波和1.06 μm光波的烟雾透过率测试数据.烟箱1.8 m烟道上的测试数据表明:配方2推进剂优于配方1推进剂,10.6 μm光波的烟雾透过率96%~97%大于1.06 μm光波的烟雾透过率92%~93%.%The main performance of hyper-velocity kinetic missile and its solid propellants is described. The mechanism of laser attenuation caused by the plume of solid rocket motor is analyzed. The plume transmittance of two solid-propellant formulations was tested by the aid of a smoke-box when the laser wavebands were 1.06 μm and 10.6 μm. The plume transmittance in 1.06 μm laser was tested with 1.064 μm laser modulation emitting, receiving and data acqui-sition system. The plume transmittance of 10.6 μm laser was derived after the optical transmittance of 2 μm ~13 μm laser was tested with the blackbody radiation source, spectroradiometer and data acquisition system. The tested data indicates that formulation 2 is better than formulation 1, and the plume transmittance 96%~97% of 10.6 μm laser is higher than that 92%~93% of 1.06 μm laser.

  5. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    O. A. Vorozheeva


    Full Text Available Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence

  6. Mars Integrated Propellant Production System Project

    National Aeronautics and Space Administration — The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian...

  7. Mars Integrated Propellant Production System Project

    National Aeronautics and Space Administration — The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian...

  8. ISRO Polyol - The Versatile Binder for Composite Solid Propellants for Launch Vehicles and Missiles

    V. N. Krishnamurthy


    Full Text Available A family of propellants based on a low cost hydroxy terminated binder has been developed and proved in large size motors. It can meet the requirements of Apogee motors as well as large boosters. The system offers advantages comparable with HTPB propellants in terms of high ballistic performance, stringent mechanical properties, ease and reliability of cure even at ambient conditions and high storage stability. The near-Newtonian flow behaviour, simplicity and processing characteristics of this saturated binder propellant are particularly note-worthy.

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


    Knockdown Factors for Common Propellant Families used in the United States. Propellant FamjiyT 1 5__ yr 1 0l s 1 15 ys 2rs . 1.. 25 yrs- CMDB a 0.96 0.92...3] Meili, G.;Thepenie,J.;Pasquier, M.; Dubroca, G.; Mechanical Design of Case-bonded CMDB Grains by a Non-linear Viscoelastic Method, AIAA-80-1177

  10. Ballistic properties and burning behaviour of an ammonium perchlorate/guanidine nitrate/sodium nitrate airbag solid propellant

    A. Ulas; G.A. Risha; K.K. Kuo [Middle East Technical University, Ankara (Turkey). Department of Mechanical Engineering


    An experimental investigation on the determination of ballistic properties and burning behavior of a composite solid propellant for airbag application was conducted. The experimental results were obtained using a high-pressure optical strand burner. Steady-state burning rates were determined for a pressure range of 20.8-41.5 MPa and initial propellant temperatures of -30 to +80{sup o}C. For the pressure and temperature ranges tested, the temperature sensitivity was on the order of 1 x 10{sup -3} K{sup -1}. The pressure exponent was found to be a function of the initial propellant temperature and was 0.75 at 25{sup o}C. The activation energy and the pre-exponential factor of the Arrhenius equation are 2.735 kcal/mol and 15.06 cm/s, respectively. The pressure deflagration limit for this propellant was found to be in the range of 8.37-8.72 MPa. During combustion, small condensed-phase spherical particles were ejected from the burning surface. The size of the particles decreased with either increasing the pressure or the initial propellant temperature. For pressures below 41.4 MPa, average particle size was on the order of 900 {mu}m, and at 84.4 MPa, the bead size was much smaller, on the order of 300 {mu}m. A chemical analysis on these particles using both the ESEM and the X-ray diffraction method indicated that the material of the beads was mostly sodium chloride with a small amount of silicon-containing compounds. 6 refs., 17 figs.



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


    molecular mass, g/mole 7K temperature sensitivity of steady chamber pressure, K1( PC combustion chamber gas density, g/cm 3 pp propellant density, g...chamber volume (V), the molecular mass of the gas (M), the gas constant (R), the flame temperature (T.) and the total propellant weight (WP) M m = (42...TNO PML G.Gadiot1 NORWAY, NORDIC AMMUNITION COMPANY St. Haugen SPAIN INSTITUTO NACIONAL DE TECNICA M. Mulero AEROSPACIAL (INTA) UNITED KINGDOM ROYAL

  13. Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

    Thorpe, Douglas G.


    An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

  14. Flow Simulation of Solid Rocket Motors. 2; Sub-Scale Air Flow Simulation of Port Flows

    Yeh, Y. P.; Ramandran, N.; Smith, A. W.; Heaman, J. P.


    The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

  15. Reusable Solid Rocket Motor - V(RSRMV)Nozzle Forward Nose Ring Thermo-Structural Modeling

    Clayton, J. Louie


    During the developmental static fire program for NASAs Reusable Solid Rocket Motor-V (RSRMV), an anomalous erosion condition appeared on the nozzle Carbon Cloth Phenolic nose ring that had not been observed in the space shuttle RSRM program. There were regions of augmented erosion located on the bottom of the forward nose ring (FNR) that measured nine tenths of an inch deeper than the surrounding material. Estimates of heating conditions for the RSRMV nozzle based on limited char and erosion data indicate that the total heat loading into the FNR, for the new five segment motor, is about 40-50% higher than the baseline shuttle RSRM nozzle FNR. Fault tree analysis of the augmented erosion condition has lead to a focus on a thermomechanical response of the material that is outside the existing experience base of shuttle CCP materials for this application. This paper provides a sensitivity study of the CCP material thermo-structural response subject to the design constraints and heating conditions unique to the RSRMV Forward Nose Ring application. Modeling techniques are based on 1-D thermal and porous media calculations where in-depth interlaminar loading conditions are calculated and compared to known capabilities at elevated temperatures. Parameters such as heat rate, in-depth pressures and temperature, degree of char, associated with initiation of the mechanical removal process are quantified and compared to a baseline thermo-chemical material removal mode. Conclusions regarding postulated material loss mechanisms are offered.

  16. Vibration, acoustic, and shock design and test criteria for components on the Solid Rocket Boosters (SRB), Lightweight External Tank (LWT), and Space Shuttle Main Engines (SSME)


    The vibration, acoustics, and shock design and test criteria for components and subassemblies on the space shuttle solid rocket booster (SRB), lightweight tank (LWT), and main engines (SSME) are presented. Specifications for transportation, handling, and acceptance testing are also provided.

  17. Lidar Sensor Performance in Closed-Loop Flight Testing of the Morpheus Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Roback, V. Eric; Pierrottet, Diego F.; Amzajerdian, Farzin; Barnes, Bruce W.; Bulyshev, Alexander E.; Hines, Glenn D.; Petway, Larry B.; Brewster, Paul F.; Kempton, Kevin S.


    For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging Flash Lidar is a second generation, compact, real-time, aircooled instrument developed from a number of components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The Flash Lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision (1-s). The Flash Lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Navigation Doppler Lidar (NDL) system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The NDLâ€"TM"s measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter (LA), also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the Flash Lidar, can provide range along a separate vector. The LA measurements are also fed

  18. Effect of Burning Rate Modifiers on Subatmospheric Flame Temperatures of AP/HTPB Composite Solid Propellants

    S. Krishnan


    Full Text Available Using 30 um. pt and Pt 13 percent Rh thermocouples, flame temperatures of uncatalysed andcatalysed ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/lffPB composite solidpropellants were measured under subatmospheric conditions. Ferric oxide F e 2 and copper chromite(CC were the catalysts used. The study demonstrates that Fe2O3 catalysed propellant, notwithstandingits least combustion efficiency undt;r subatmospheric conditions and weak gas-phase flame, has themaximum burning rate enhancement. This is argued to be due to the increased surface and subsurfacereactions caused by Fe2O3. CC-catalysed propellant burns to the least subatmospheric pressure withminimum loss in combustion efficiency indicating that this class of propellant may be more suitablefor base-bleed applications.

  19. The use of image analysis for the study of interfacial bonding in solid composite propellant



    Full Text Available In the framework of this research, the program Image Pro Plus was applied for determining the polymer–oxidizer interactions in HTPB-based composite propellants. In order to improve the interactions, different bonding agents were used, and their efficiency was analyzed. The determination of the quantity, area and radius of non-bonded oxidizer crystals is presented. The position of formed cracks in the specimen and their area has a great influence on the mechanical properties of composite propellant. The preparation of the composite propellant in order to enable the photographing of their structure by means of stereoscopic and metallographic microscopes with the digital camera is also described as well.

  20. Effects of the acceleration vector on transient burning rate of an aluminized solid propellant.

    Northam, G. B.


    Experimental results concerning the transient burning-rate augmentation of a 16% aluminum polybutadiene acrylic acid (PBAA) propellant burned in a 2-in. web motor at pressure levels from 300 to 1200 psia with centrifugal accelerations from 0 to 140 g. The orientation of the acceleration vector was varied to determine its effect on the transient burning rate. The burning-rate augmentation was strongly dependent on (1) acceleration level, (2) propellant distance burned (or burn time), and (3) orientation of the acceleration vector with respect to the burning surface. This transient rate augmentation resulted from the retention of molten metallic residue on the burning surface by the normal acceleration loading. The presence of the residue altered the combustion zone heat transfer and caused increased localized burning rates, as evidenced by the pitted propellant surfaces that were observed from extinction tests conducted at various acceleration levels.

  1. Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

    Watanabe, Yoichi; Appelbaum, Jacob; Diaz, Nils; Maya, Isaac


    Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

  2. Probabilistic assessment for the design of high reliability objects (greater than 0.999) such as solid propellant grains

    Nahon, S.; Silvestrini, P.; Thepenier, J.


    A Bayesian method is proposed which uses the safety margin to justify the design of high-reliability objects characterized by one-shot destructive functioning and by a high manufacturing cost. To validate the Bayesian reliability model, a program of overtests has been conducted with seven solid propellant grains of six tons each. Each trial has been performed with a reduced safety margin leading to a design reliability of 0.5 at 60 percent confidence level. The reliability model gives reasonably good agreement with experimental results as the confidence level stabilizes, leading to the conclusion that it can be applied to estimate higher reliability levels (greater than 0.999).

  3. Solid Rocket Motor for Ultralow Temperature Operation During the Mars Sample Return Mission Project

    National Aeronautics and Space Administration — A small Mars (or other celestial body) ascent vehicle is unlikely to achieve the necessary propellant fraction required to achieve orbit. Scaling down of liquid...

  4. The sky is falling: chemical characterization and corrosion evaluation of deposition produced during the static testing of solid rocket motors.

    Doucette, William J; McNeill, Laurie S; Mendenhall, Scout; Hancock, Paul V; Wells, Jason E; Thackeray, Kevin J; Gosen, David P


    Static tests of horizontally restrained rocket motors at the ATK facility in Promontory UT, USA result in the deposition of entrained soil and fuel combustion products, referred to as Test Fire Soil (TFS), over areas as large as 30-50 mile (80-130 km) and at distances up to 10-12 miles (16-20 km) from the test site. Chloride is the main combustion product generated from the ammonium perchlorate-aluminum based composite propellant. Deposition sampling/characterization and a 6-month field corrosivity study using mild steel coupons were conducted in conjunction with the February 25th 2010 FSM-17 static test. The TFS deposition rates at the three study sites ranged from 1 to 5 g/min/m. TFS contained significantly more chloride than the surface soil collected from the test site. The TFS collected during two subsequent tests had similarly elevated chloride, suggesting that the results obtained in this study are applicable to other tests assuming that the rocket fuel composition remains similar. The field-deployed coupons exposed to the TFS had higher corrosion rates (3.6-5.0 mpy) than paired non-exposed coupons (1.6-1.8 mpy). Corrosion rates for all coupons decreased over time, but coupons exposed to the TFS always had a higher rate than the non-exposed. Differences in corrosion rates between the three study sites were also observed, with sites receiving more TFS deposition having higher corrosion rates.

  5. Time-Accurate Computational Fluid Dynamics Simulation of a Pair of Moving Solid Rocket Boosters

    Strutzenberg, Louise L.; Williams, Brandon R.


    Since the Columbia accident, the threat to the Shuttle launch vehicle from debris during the liftoff timeframe has been assessed by the Liftoff Debris Team at NASA/MSFC. In addition to engineering methods of analysis, CFD-generated flow fields during the liftoff timeframe have been used in conjunction with 3-DOF debris transport methods to predict the motion of liftoff debris. Early models made use of a quasi-steady flow field approximation with the vehicle positioned at a fixed location relative to the ground; however, a moving overset mesh capability has recently been developed for the Loci/CHEM CFD software which enables higher-fidelity simulation of the Shuttle transient plume startup and liftoff environment. The present work details the simulation of the launch pad and mobile launch platform (MLP) with truncated solid rocket boosters (SRBs) moving in a prescribed liftoff trajectory derived from Shuttle flight measurements. Using Loci/CHEM, time-accurate RANS and hybrid RANS/LES simulations were performed for the timeframe T0+0 to T0+3.5 seconds, which consists of SRB startup to a vehicle altitude of approximately 90 feet above the MLP. Analysis of the transient flowfield focuses on the evolution of the SRB plumes in the MLP plume holes and the flame trench, impingement on the flame deflector, and especially impingment on the MLP deck resulting in upward flow which is a transport mechanism for debris. The results show excellent qualitative agreement with the visual record from past Shuttle flights, and comparisons to pressure measurements in the flame trench and on the MLP provide confidence in these simulation capabilities.

  6. Modeling Plasma Flow in Solid Propellant Charges Using the NGEN Multiphase CFD Code


    using these equations derived by a formal averaging technique applied to the microscopic flow. These equations require a number of constitutive laws...disk (dimensions shown are from Chang and Howard [32]). acrylic, that allows cinematography of plasma flows and ignition events along the propellant

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

    Northam, G. B.


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

  8. Experimental determination of convective heat transfer coefficients in the separated flow region of the Space Shuttle Solid Rocket Motor

    Whitesides, R. Harold; Majumdar, Alok K.; Jenkins, Susan L.; Bacchus, David L.


    A series of cold flow heat transfer tests was conducted with a 7.5-percent scale model of the Space Shuttle Rocket Motor (SRM) to measure the heat transfer coefficients in the separated flow region around the nose of the submerged nozzle. Modifications were made to an existing 7.5 percent scale model of the internal geometry of the aft end of the SRM, including the gimballed nozzle in order to accomplish the measurements. The model nozzle nose was fitted with a stainless steel shell with numerous thermocouples welded to the backside of the thin wall. A transient 'thin skin' experimental technique was used to measure the local heat transfer coefficients. The effects of Reynolds number, nozzle gimbal angle, and model location were correlated with a Stanton number versus Reynolds number correlation which may be used to determine the convective heating rates for the full scale Space Shuttle Solid Rocket Motor nozzle.

  9. Standard Molded Composite Rocket Pyrogen Igniter - A progress report

    Lucy, M. H.


    The pyrogen igniter has the function to furnish a controlled, high temperature, high pressure gas to ignite solid propellant surfaces in a rocket motor. Present pyrogens consist of numerous inert components. The Standard Molded Pyrogen Igniter (SMPI) consists of three basic parts, a cap with several integrally molded features, an ignition pellet retainer plate, and a tube with additional integrally molded features. A description is presented of an investigation which indicates that the SMPI concept is a viable approach to the design and manufacture of pyrogen igniters for solid propellant rocket motors. For some applications, combining the structural and thermal properties of molded composites can result in the manufacture of lighter assemblies at considerable cost reduction. It is demonstrated that high strength, thin walled tubes with high length to diameter ratios can be fabricated from reinforced plastic molding compound using the displacement compression process.

  10. 复合固体推进剂老化研究%Summary of Study on Ageing of Composite Solid Propellant

    于亮; 陈燕; 李智; 李树娟


    针对复合固体推进剂的老化问题,探讨了复合固体推进剂老化的主要因素;从复合固体推进剂组分的影响、环境湿度、储存温度等方面分析了其影响复合固体推进剂老化的机理;综述了复合固体推进剂贮存寿命的几种预估方法:力学性能法、阿累尼乌斯方程式法、凝胶含量法、傅里叶红外光谱法和动态粘弹法,并对这些研究方法的内容和结果可信度进行了分析。最后对复合固体推进剂老化研究的发展趋势进行了展望。%In the view of the composite solid propellant ageing,the main factors of the composite solid propellant aging were discussed and aging mechanism of the composite solid propellant was analyzed from the influence of the component of solid propellant,environment humidity and storing temperature,several forecast methods of composite solid propellant storage life was overviewed:the methods of mechanical properties, activation energy, gel content, Fourier infrared spectrum analysis and dynamic mechanical analysis. And simultaneously the main content of these methods and the reliability of these results were analyzed. Finally,aging research trends of the composite solid pro-pellant were discussed.

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

    Zibit, Benjamin Seth

    This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the

  12. Study of Cavitation/Vaporization in Liquid Rocket Thruster Injectors


    Caveny, L. H., and Summerfield, M., Aluminized Solid Propellants Burning in a Rocket Motor Flowfield, AIAA Journal, Vol. 16, No. 7, 1978, pp. 736-739. [2...the swirl chamber, and the pulsator and manifold are made of 304 stainless steel . Figure 1: Nomenclature defined for swirl injector. Table 1...Wayne, NJ, 2009 [17] MATLAB, Matrix Laboratory, Software Package, R2009a, The MathWorks, Natick, MA, 2009. [18] Coleman, H.W. and Steele , W.G

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


    CMDB propellants decreased 40 percent and CTPB propellants decreased about 3 percent. The AC volume resistivity decreased with increasing frequency for...each sample. CMDB propel- lants are somewhat conductive even at the low frequencies. Both HTPB and CTPB propellants have a relatively high resistivity

  14. The Internal Flow Modeling of a Simulated Solid Propellant-Liner Debond Using Loci-CHEM


    other non-destructive evaluation ( NDE ) methods is a routine part of many motor programs. Those evaluations do not measure propellant mechanical...6 , was used in the current effort. The model combines the Wilcox k−ω model in the near-wall region with the k− model in the outer part of the...flaw geometry to a far greater extent, since NDE equipment generally do not have high resolution. Therefore, several sensitivity analyses were

  15. Enhanced Large Solid Rocket Motor Understanding Through Performance Margin Testing: RSRM Five-Segment Engineering Test Motor (ETM-3)

    Huppi, Hal; Tobias, Mark; Seiler, James


    The Five-Segment Engineering Test Motor (ETM-3) is an extended length reusable solid rocket motor (RSRM) intended to increase motor performance and internal environments above the current four-segment RSRM flight motor. The principal purpose of ETM-3 is to provide a test article for RSRM component margin testing. As the RSRM and Space Shuttle in general continue to age, replacing obsolete materials becomes an ever-increasing issue. Having a five-segment motor that provides environments in excess of normal opera- tion allows a mechanism to subject replacement materials to a more severe environment than experienced in flight. Additionally, ETM-3 offers a second design data point from which to develop and/or validate analytical models that currently have some level of empiricism associated with them. These enhanced models have the potential to further the understanding of RSRM motor performance and solid rocket motor (SRM) propulsion in general. Furthermore, these data could be leveraged to support a five-segment booster (FSB) development program should the Space Shuttle program choose to pursue this option for abort mode enhancements during the ascent phase. A tertiary goal of ETM-3 is to challenge both the ATK Thiokol Propulsion and NASA MSFC technical personnel through the design and analysis of a large solid rocket motor without the benefit of a well-established performance database such as the RSRM. The end result of this undertaking will be a more competent and experienced workforce for both organizations. Of particular interest are the motor design characteristics and the systems engineering approach used to conduct a complex yet successful large motor static test. These aspects of ETM-3 and more will be summarized.

  16. 固体火箭发动机药柱加压固化仿真%Simulation on pressure cure of solid rocket motor grain

    宗路航; 杜聪; 卢山; 姚东; 郜婕; 沙宝林


    For those case-bonded casting solid rocket motors ( SRMs) with large outside/inside radius ratio of the grain, high thermal strain will be generated in the propellant grain subjected to thermal loading, which significantly limit the performance of SRM. Pressure cure is an effective method to reduce the thermal strain in the grain. In this paper, the theory of pressure cure was analyzed and the relationship between the desired pressure and the parameters of the SRM was deduced. Then, a finite element method ( FEM) of pressure cure named two step method was proposed. Theoretical calculation and FEM simulation were carried out on a tube motor with four different material cases. The recommended pressure of pressure cure of four cases were given out.%对于药柱外/内径比(m数)很大的贴壁浇注式固体火箭发动机,在固化降温后,推进剂药柱内会产生显著的热应变,这严重限制了发动机的进一步高性能化. 加压固化是一种降低推进剂药柱内热应变的有效方法. 文中分析了加压固化的原理,推导出了加压固化所需压强与发动机参数之间的关系式,提出了一种两步分析法的加压固化有限元分析方法. 针对4种不同壳体材料的圆管发动机,进行了加压固化理论计算与有限元仿真分析,给出了4种壳体加压固化时的推荐压强.

  17. Liquid Rocket Engine Testing


    booster rocket engines • 6000-10000 psia capabilities – Can use gaseous nitrogen, helium, or hydrogen to pressurize propellant tanks 9Distribution A...Approved for Public Release; Distribution Unlimited. PA Clearance 16493 Simplified Test Stand Layout Oxidizer  TankFuel  Tank High  Pressure   Gas (GN2...requires large, complex facilities to deliver propellant at the proper pressure , temperature, and flow rates • The enormous energies involved

  18. Improved hybrid rocket fuel

    Dean, David L.


    McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

  19. The evaluation of solid phase micro-extraction fibre types for the analysis of organic components in unburned propellant powders.

    Dalby, Oliver; Birkett, Jason W


    This work describes the evaluation of various solid phase micro-extraction (SPME) fibre types for the detection of compounds originating from particles of unburned propellant powders. These compounds may also be found in association with organic gunshot residues (OGSR). Seven SPME fibres were assessed based on their ability to extract the compounds of interest (diphenylamine (DPA), 4-nitrodiphenylamine (4-NDPA), ethyl centralite (EC), nitroglycerin (NG) and dibutyl phthalate (DBP)) from four ammunition types across three calibres (9 mm, 5.56 mm and 7.62 mm). Extracts were analysed by gas chromatography/mass spectrometry (GC/MS). Results indicated that the 65 μm polydimethylsiloxane/divinylbenzene (PDMS/DVB) was the most suitable fibre type for the extraction of these compounds across the ammunition types tested. Optimal extraction time parameters were also assessed with a 35-min period determined to be suitable. A number of previously unreported considerations for extracting propellant powders and potentially OGSR related materials are discussed.

  20. Evaluation of Additives to Reduce Solid Propellant Flammability in Ambient Air.


    WP 4gw«’WUK_u.,HU!«J»B^!PB!*-pi!^p5ggB5,^.IJll.l II Hill JMflH» !,S--*"JU’j*llfgg V. D. Organometallic Compounds E. Enthalpy and Gas near the burning surface of the propellent where the reactive species can be neutralised more effectively. B. Char Formers Several phisphorus... Enthalpy and Gas Composition Measurements of Coolants and Binders* Two experimental techniques were used to measure the enthalpy changes that are

  1. Liquid Propellants for Advanced Gun Ammunitions

    K. P. Rao


    Full Text Available With constant improvements, the conventional solid propellants for guns have almost reached their limit in performance. Liquid gun propellants are promising new comers capable of surpassing these performance limits and have numerous advantages over solid propellants. A method has been worked out to predict the internal ballistics of a liquid propellant gun and illustrated in a typical application.

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

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


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

  3. Metallic hydrogen: The most powerful rocket fuel yet to exist

    Silvera, Isaac F [Lyman Laboratory of Physics, Harvard University, Cambridge MA 02138 (United States); Cole, John W, E-mail: silvera@physics.harvard.ed [NASA MSFC, Huntsville, AL 35801 (United States)


    Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, I{sub sp}. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of {approx}460s; metallic hydrogen has a theoretical I{sub sp} of 1700s. Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

  4. A new method for detecting the aging property of HTPB solid propellant%HTPB固体推进剂老化性能检测新方法

    任宁莉; 赵新强; 邓凯; 余志刚


    引入了光致正电子湮灭分析方法 (PIPA,Photon Induced Positron Analysis),对HTPB固体推进剂老化性能进行检测研究。介绍了PIPA的原理、PIPA试验平台的搭建以及PIPA的数值处理方法,并用511KeV能谱的FWHM值ε(半高宽,Full Width Half Maximum)表征HTPB固体推进剂老化引起的微观变化,所得结论与固体推进剂的常规测试一致,证明了PIPA用于固体推进剂老化性能无损检测的可行性。%To study the aging character of HTPB solid propellant,PIPA(Photon Induced Positron Analysis) was presented.The positrons created by the PIPA process are formed throughout the solid propellant.The principle of PIPA was introduced along with the built-up of the PIPA system and the arithmetic method.The information of HTPB solid propellant micro change due to aging was extracted from the FWHM of 511keV spectra.There is positive correlation between the results of experiments and regular tests,indicating that PIPA is a feasible nondestructive detection method for solid propellant aging evaluation.

  5. The materials and elements production practice of counter-erosional and thermal protection system of the SPR-solid-propellant sustainer nozzle

    Shkurenko, V. M.


    This paper presents the production scheme for heat- and erosion-protective carbon plastic materials for heat shield elements of solid-propellant nozzles. Attention is also given the method of manufacturing adhesive joint assemblies, and the production scheme is included.

  6. Assessment of Various Flow Solvers Used to Predict the Thermal Environment inside Space Shuttle Solid Rocket Motor Joints

    Wang, Qun-Zhen; Cash, Steve (Technical Monitor)


    It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.

  7. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    Kalomba Mboyi


    Full Text Available A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the emitting material with the radioisotope decay heat and by powering the different valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

  8. Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs

    Keba, John E.


    This paper presents viewgraphs of The Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs. The topics include: 1) Rocket Turbomachinery Shaft Seals (Inter-Propellant-Seal (IPS) Systems, Lift-off Seal Systems, and Technology Development Needs); 2) Rocket Engine Characteristics (Engine cycles, propellants, missions, etc., Influence on shaft sealing requirements); and 3) Conclusions.

  9. Changes in the Calorimetric Value and Ignition Temperature of Composite Solid Propellants During Ageing-A Note

    K. Kishore


    Full Text Available Calorimetric value and ignition temperature of carboxy terminated polybutadiene/ammonium perchlorate propellant decrease during accelerated ageing. The behaviour has been explained on account of binder loss and condensed phase reactions in the propellant matrix.

  10. Stimuli-responsive cylindrical hydrogels mimic intestinal peristalsis to propel a solid object.

    Nistor, V; Cannell, J; Gregory, J; Yeghiazarian, L


    The emerging field of soft robotics relies on soft, stimuli-responsive materials to enable load transport, manipulation, and mobility in complex unconstrained environments. These materials often need to replicate biological functionality such as muscle contractions and flexibility. Here we demonstrate a soft actuator prototype based on thermosensitive PNIPAAM hydrogels that can transport and manipulate objects. A hollow cylindrical hydrogel was selectively heated and cooled with Peltier devices to yield a traveling wave of shrinking and swelling akin to intestinal peristalsis. A 4 mm diameter bead was placed inside the cylinder and propelled 19.5 mm, equal to distance traveled by the peristaltic wave. We derived conditions that enable peristaltic transport as a function of transporter-cargo design parameters. We conclude that hydrogel-based peristaltic manipulators covering 2 orders of magnitude in stiffness (1-10(2) kPa) could transport cargo spanning 4 orders of magnitude in size (μm-m).

  11. The effect of solid phase reactions on the ballistic properties of propellants

    Schmidt, W. G.


    The combustion of NHClO4 composite propellants was studied between 15 and 3000 psi. The emphasis was on determining the mechanisms by which the fuel components influence the burning rate of the composites. All combustion experiments were performed with pressed powder strands. The fuels affected the combustion mechanism of the composite through their effect on the oxidizer decomposition mechanism, the composite surface temperature, and by reacting heterogeneously with the oxidizer on the surface of the composite. The results indicate that the fuel component has a significant effect on the surface temperature in both catalyzed and uncatalyzed composites. The data suggest that the contribution, to the overall combustion reaction, of heterogeneous reactions increases with increasing pressure. There is also evidence that even in steady state combustion the ignition of individual oxidizer particles is an important parameter in determining the combustion characteristics of the composite.

  12. Microwave measurement of the solid propellant pressure-coupled response function

    Strand, L. D.; Mcnamara, R. P.; Magiawala, K. R.


    The results of an investigation are presented on the applicability of a microwave Doppler shift technique for directly determining propellant response functions over the desired frequency range. The investigation consisted of three phases. In Phase 1 the validity of the technique was established by comparing measured pressure-coupled response function data to existing data from T-burners and rotating valve tests. In Phase 2 a new microwave burner-pressure modulation system capable of achieving frequencies and mean chamber pressures of at least 1500 Hz and 10.5 MPa (1500 psia), respectively, was developed. During Phase 3 test firings are being carried out to define the frequency limit, response function resolution, and precision of the new design.

  13. Thermal-Flow Code for Modeling Gas Dynamics and Heat Transfer in Space Shuttle Solid Rocket Motor Joints

    Wang, Qunzhen; Mathias, Edward C.; Heman, Joe R.; Smith, Cory W.


    A new, thermal-flow simulation code, called SFLOW. has been developed to model the gas dynamics, heat transfer, as well as O-ring and flow path erosion inside the space shuttle solid rocket motor joints by combining SINDA/Glo, a commercial thermal analyzer. and SHARPO, a general-purpose CFD code developed at Thiokol Propulsion. SHARP was modified so that friction, heat transfer, mass addition, as well as minor losses in one-dimensional flow can be taken into account. The pressure, temperature and velocity of the combustion gas in the leak paths are calculated in SHARP by solving the time-dependent Navier-Stokes equations while the heat conduction in the solid is modeled by SINDA/G. The two codes are coupled by the heat flux at the solid-gas interface. A few test cases are presented and the results from SFLOW agree very well with the exact solutions or experimental data. These cases include Fanno flow where friction is important, Rayleigh flow where heat transfer between gas and solid is important, flow with mass addition due to the erosion of the solid wall, a transient volume venting process, as well as some transient one-dimensional flows with analytical solutions. In addition, SFLOW is applied to model the RSRM nozzle joint 4 subscale hot-flow tests and the predicted pressures, temperatures (both gas and solid), and O-ring erosions agree well with the experimental data. It was also found that the heat transfer between gas and solid has a major effect on the pressures and temperatures of the fill bottles in the RSRM nozzle joint 4 configuration No. 8 test.

  14. Innovative Swirl Injector for LOX and Hydrocarbon Propellants Project

    National Aeronautics and Space Administration — Gases trapped in the propellant feed lines of space-based rocket engines due to cryogenic propellant boil-off or pressurant ingestion can result in poor combustion...

  15. Scaled Rocket Testing in Hypersonic Flow

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


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

  16. Erosive Augmentation of Solid Propellant Burning Rate: Motor Size Scaling Effect

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


    Two different independent variable forms, a difference form and a ratio form, were investigated for correlating the normalized magnitude of the measured erosive burning rate augmentation above the threshold in terms of the amount that the driving parameter (mass flux or Reynolds number) exceeds the threshold value for erosive augmentation at the test condition. The latter was calculated from the previously determined threshold correlation. Either variable form provided a correlation for each of the two motor size data bases individually. However, the data showed a motor size effect, supporting the general observation that the magnitude of erosive burning rate augmentation is reduced for larger rocket motors. For both independent variable forms, the required motor size scaling was attained by including the motor port radius raised to a power in the independent parameter. A boundary layer theory analysis confirmed the experimental finding, but showed that the magnitude of the scale effect is itself dependent upon scale, tending to diminish with increasing motor size.

  17. Development of high temperature materials for solid propellant rocket nozzle applications. [tantalum carbides-tungsten fiber composites

    Manning, C. R., Jr.; Honeycutt, L., III


    Evaluation of tantalum carbide-tungsten fiber composites has been completed as far as weight percent carbon additions and weight percent additions of tungsten fiber. Extensive studies were undertaken concerning Young's Modulus and fracture strength of this material. Also, in-depth analysis of the embrittling effects of the extra carbon additions on the tungsten fibers has been completed. The complete fabrication procedure for the tantalum carbide-tungsten fiber composites with extra carbon additions is given. Microprobe and metallographic studies showed the effect of extra carbon on the tungsten fibers, and evaluation of the thermal shock parameter fracture strength/Young's Modulus is included.

  18. Rocket Engine Innovations Advance Clean Energy


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

  19. Inter-Batch Variation and the Effect of Casting Vacuum on Ballistic and Mechanical Properties of a High Performing Cast Composite Rocket Propellant


    13 8 Scrape down - - - 9 Mix 30 0 45 10 Scrape down - - - 11 Add IPDI 30 1 4 12 Scrape down - - - 13 Mix 30 0 20 The dry clumps of material...http:// web Booster, HTPB, high performance composite propellant 19

  20. Studies on crack extension in solid propellant%固体推进剂裂纹扩展研究综述

    胡松启; 周宴星; 刘迎吉; 陈静


    The research status of the crack extension in solid propellant at home and abroad are introduced. Experimental study and theoretical analysis of the crack extension in solid propellant are summarized. The advantages and disadvantages of such study are analyzed. The influences of various factors on crack extension in solid propellant are generalized. The results show that the tension ma- chine and high-speed photography are main means for studying the mechanics characteristic of pro- pellant with crack, and the combustion property of propellant and the extension of crack; the working pressure in combustion chamber, pressure boost rate, style and size of crack, and burning rate of pro- pellant arc the important factors for the crack extension in solid propellant. It is pointed out that it is necessary to research the quantitative relationship between various factors and crack extension.%介绍国内外关于固体推进剂裂纹扩展的研究现状,总结固体推进剂裂纹扩展的试验研究和理论分析方法,归纳影响固体推进剂裂纹扩展的各种因素。分析认为:采用拉伸装置研究有裂纹推进剂力学性能和利用高速摄影系统研究裂纹燃烧和扩展情况是当前主要的研究手段;发动机燃烧室内压力、升压速率、裂纹形状尺寸和推进剂燃速是装药裂纹扩展的重要因素;必须进一步开展各因素和裂纹扩展的定量关系研究。