• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.21-26
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• 2019

Exposure to a rapidly depressurized environment causes extinction of a burning solid propellant. Experiments have been conducted to determine the rate of depressurization required to extinguish a burning solid propellant. For this purpose, a depressurization combustor was designed and fabricated. The results of this experiment were used to determine the boundary between extinction and non-extinction of AP/HTPB solid propellants under different propellant compositions. Experimental results show that the initial and final pressures have a considerable effect on the critical depressurization rate.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.3
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• pp.12-17
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• 2020

The mechanical properties of the propellant with yellow iron oxide were slightly increased compared to the propellant with red iron oxide. The propellant with yellow iron oxide used two types of AP. As the ratio of small particles of AP increased, the burning rate increased. The propellant may be applied to the propellant under operating conditions of 17.5 mm/sec or less having a pressure index of 0.5. The burning rate downs in the mixer scale-up. The stress at maximum load of propellant decreased and the strain at maximum load increased in the mixer scale-up. The yellow iron oxide did not affect the adhesive force between the insulation/liner/propellant.

• Journal of Mechanical Science and Technology
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• v.15 no.4
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• pp.473-481
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• 2001

By depicting the transfer of heat and combustion reaction to take place within thin gas layers close to the propellant surface burning in a steady-state fashion, a mathematical equation has been deduced to describe the burning rate of solid propellant as a function of initial grain temperature and chamber pressure. It has been also assumed that chemical reaction could take place in premixing-diffusing zone but were carried out mainly in the reaction-flame zone. All these phenomena taken place in each zone of combustion have been assumed to be steady-state. In the present investigation, the equation, γ=$\kappa$$.$(1/R(T(sub)i+C))(sup)n$.$exp(-E(sub)a/R(T(sub)i+C))(P/z) is being presented and it is compared with experimental data. The proposed model has been tested and evaluated vis-a-vis strand burner data for three different propellants based on CTPB, and it has been found that the deviation of the computed burning rates from the measured rates ranged up to 2%.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.78-85
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• 1999

Burning characteristics of solid propellants embedded with four kinds of metal wires(Ag, Cu, Al, Ni-Cr wire) were studied with varying wire diameters(O.10.8 mm) lot N-5 propellant. It was found that the order of the burning rate increment ratio($r_w$/$r_sb$) was Ag wire > Cu wire > Al wire> Ni-Cr wire which was the same as the order of the magnitude of thermal diffusivity. The burning rate increment ratio($r_w$/$r_sb$) of N-5 propellant was less than that of composite Propellant because of the difference of adiabatic flame temperature and flame structure. When Ag, Cu and Al wire having high thermal diffusivity were embedded in N-5 propellant, the plateau and mesa characteristics of the double base propellant were disappeared, but not disappeared in the case of propellant embedded with Ni-Cr wire due to its poor thermal conductivity.

• Composites Research
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• v.33 no.5
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• pp.309-314
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• 2020

Polymer composites were used to reduce the weight of the spacecraft's cryogenic propellant tank. Since these materials were directional, the permeability performance of the gas permeated or delivered in the stacking direction was an indicator directly related to performance such as tank stability and onboard fuel quantity estimation. In addition, the results of permeation measurements and optical analysis of the surface to verify the effect of the number of cycles exposed to the cryogenic-room temperature environment are included. As a result, the permeability was inversely proportional to the thickness and was proportional to the number of thermal shocks, and it was verified that the permeability performance was suitable for the cryogenic propellant tank material for the space launch vehicle.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.79-88
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• 2017

Ignition test of the fuel-rich propellant coated with ignition support material in the ramjet combustor condition was conducted. Ignition delay and flame holding was measured. Fuel grain consist of HTPB mixed with AP particle 15 wt.%, Al particle 5 wt.%. To cause the short ignition delay, ignition support consist of $NC/BKNO_3$ and composite propellant was coated to the fuel grain. Ethanol blended $H_2O_2$ gas generator control the temperature, pressure, $O_2$ concentration in the oxidizer gas in the air. Gas is supplied with mass flux of $200kg/m^2s$. Through the test ignition support operated well and ignition delay of 0.6 second and the Flame was sustained.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.1
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• pp.8-16
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• 2017

Structural integrity of solid rocket depends on the residual reactions between constituents of its composition(post cure, migration etc.), the oxygen(or anti-oxydent) in the free volume and humidity (desiccant) under the perfect sealed condition. Mechanical Properties of composite solid propellant arising from those factors are very complex. Moreover the propulsion are faced with thermal loads from diurnal & seasonal cycle till firing. In this study, the fast evaluation method of long term mechanical properties is suggested based on Thermo-Rheological Simplicity from curing oven to cool-down stage in view point of thermal stabilization. For this subject, endurance tester having temperature control capability are devised. From the results from incremental load and strain, non-linear characteristics are discussed.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.3
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• pp.9-17
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• 2002

This Study is to investigate 8 composite propellants including Butacene and ${Bi_2}{O_3}$ by Shotgun/RQ Bomb test. Burning rate and mechanical properity are known to be major factors in determining the deflagrability of propellant. Propellant including over 5.5% Butacene(Ferrocene grafted HTPB) burned out over 135 m/s of impact velocity during Shotgun/RQ Bomb test. It was blown that Butacene was very sensitive material under high velocity impact. In the test results, propellants under 25mm/s in burning rate at 1500 psia could meet the requirements for IM of UN Test Series 7c(ii). Propellant deflagrabillity depends on burning rate at performance in the results of the present.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.215-218
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• 2007

Internal Ballistic modeling and performance prediction for solid propellant micro thruster was performed with heat loss to the chamber wall as an important factor of miniaturization. Simple l-D end-burner type thruster and general HTPB-AP type composite propellant were selected for computation model. The results showed that the performance loss with the heat loss to the surroundings becomes larger as the surface-to-volume ratio is increased. In this case, the total impulse was reduced about 3% of the case in adiabatic condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.78-86
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• 2019

Aluminum, which is a metal fuel contained in the composite solid propellant, is not ignited and burned on the combustion surface by the oxide film, and it partially melts and coalesces with surrounding aluminum particles. For the evaluation and design of the propellant performance, modeling was performed to predict the size and distribution of agglomerated particles, and the size and distribution of agglomerates were compared and verified through experiment. The predicted values showed the tendency to decrease with pressure as in the experiment, but the error increased as the pressure increased. The agglomerated particle distribution graph showed a difference in the volume fraction although the diameter at the peak was the same.