• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.4
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• pp.321-328
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• 2021

This paper derives numerically new buckling Knockdown factors for the lightweight design of the composite propellant tanks for space launch vehicles. A nonlinear finite element analysis code, ABAQUS, is used for the present postbuckling analysis of composite cylinders under compressive loads. Various thickness ratios (R/t) and slenderness ratios (L/R) are considered and Single Perturbation Load Approach is applied to represent the geometric initial imperfection of the composite cylinder. For the composite cylinder with thickness ratio of 500 and slenderness ratio of 2.04, the buckling Knockdown factor derived in this work is higher by 84.38% than NASA's previous buckling design criteria. Therefore, it is investigated that a lightweight design is possible when the present Knockdown factors are used for the design of composite propellant tanks. In addition, it is shown that global buckling loads and buckling Knockdown factors decrease as the thickness ratio or slenderness ratio of composite cylinders increases.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.6
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• pp.1-8
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• 2007

The objective of the present work is to characterize design parameters of solid propellant ignitor for composite, double base, and nitramine propellants using arc image furnace. Arc image furnace and fiber optics surface reflectometer were used to measure ignition delay time and reflected optical energy of several compositions of composite, double base and nitramine base rocket propellant at different pressure levels each other. The order of ignitability was double base > composite> nitramine propellants at initial pressure of over 75 psia. The highest ignition energy was needed to ignite nitramine propellant, however, the ignition delay time decreased abruptly as the pressure increased up to the range of $75{\sim}400$ psia. The absorbtion of radiation energy could be increased by the addition of small amount of opacifiers as carbon black, ZrC, WC and burning catalyst.

• Composites Research
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• v.36 no.2
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• pp.132-139
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• 2023

A linerless composite propellant tank was designed and manufactured by using the carbon fiber-reinforced composite materials which have superior strength-to-weight ratio in order to reduce weight of the tank. In this research, we designed a sub-scale composite propellant tank with a diameter of 800 mm to withstand an MEOP of 1.7 MPa. We manufactured the boss of the tank by using the same composite materials to reduce the thermal expansion difference between the boss and the secondary-bonded composite layers of the barrel in the cryogenic environment. We used the collapsible mandrel to manufacture the tank without any liner. The mandrel was made from epoxy-based composite tooling prepregs to reduce weight of the mandrel. We manufactured the test tanks by laying up the carbon fiber fabric prepregs manually on the mandrel and then applying the autoclave cure process. We performed a proof test, a helium tightness test, a repeated pressurization test, and a burst test in room temperature. The test results demonstrate that the proposed design and manufacture process satisfies all strength requirements as well as an anti-leakage requirement.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.5
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• pp.18-26
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• 2013

This study suggested techniques to reduce burning rate and their effects for the AP thermoplastic composite propellant. Burning rate obtained through ground tests using a small size motor were analyzed to investigate the effects of AP particle size and LiF of 0.5~2.0% on the inhibition reaction for the PEBAX/AP thermoplastic propellant. The results showed that utilization of large size particle of AP and addition of LiF under 2.0% can reduce the burning rate sufficiently and their quantitative effects were suggested in this paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.495-499
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• 2005

The INHA Rocket Research Institute changed the Ceramic nozzle material of their developed Solid Composite Propellant Motor with Teflon nozzle material. Static firings of the new Solid Rocket Motors was conducted on Thurst Tester to validate the increase in performance. The new enhanced Solid Roket Motor increased the total impulse by 18.3 percent while improving its reliability. The new process of manufacture reduced the time to produce a nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1994.11a
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• pp.39-45
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• 1994

Combustion characteristics of the solid composite propellants were studied from burning rate, ignition and steady combustion processes, and structure of the extinguished surfaces. Optical Vacuum Strand Burner (OVSB) system was desisted and configured to study those. Burning rates of the propellants were measured by OVSB at low pressure range by developed ten method. video camera(30 frames/s) was used to take potographs of the phenomena of ignition and combustion of propellant within the test cell of the OVSB. Burning surfaces of the propellants that were extinguished by rapid depressurization method were analyzed with Scanning Electron Microscope. (SEM).

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.26-32
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• 2015

In this study, it is evaluated for fracture toughness to analyze crack resistance properties of particulate reinforced composite propellant. Fracture toughness test using WST specimen is conducted by temperature conditions from $50^{\circ}C$ to $-60^{\circ}C$. Evaluation method for fracture toughness calculated using an equation suggested by ASTM E399 with linear elastic fracture mechanics. From these result, splitting loads and stress intensity factors of propellant increase according to decrease of test temperature. Also, the strain fields of specimen surface using digital image correlation increase as temperature decreased from $50^{\circ}C$ to $-40^{\circ}C$, but it sharply decreases at $-60^{\circ}C$ because of brittle behavior.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.383-386
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• 2006

Several HTPB/AP and HTPB/AP/HMX propellants were investigated experimentally for ignition characteristics in subatmospheric pressure. The threshold ignition pressure was 4psia for HTPB/AP composite propellant. The partial replacement of AP in HTPB/AP composite propellant by $5\sim15%$ of HMX, HNIW showed improvements in the threshold pressure was below 0.4psia. This appears to be due to the exothermic dissociation characteristics of HMX and HNIW at lower temperature $(\sim220^{\circ}C)$ than that of AP. The ignition substance B/KNO3 was coated thinly on the propellant surface for better ignition effect. As a result, ignition delay time of 15% was improved. NC is applied to $B/KNO_3$ ignition substance as a secondary binder and $NC-B/KNO-3$ suspension solution is coated to the propellant surface.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.2
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• pp.74-80
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• 2008

The feasibility study into the development of turbine spinners, which start up the turbo-pump, has been carried out and the design requirements and parameters ranges have been presented. Turbine spinners use the solid propellant as such composite propellant based AN compound with high energy plasticizers, coolants, and phase stabilizer which relieves a sensible volume change due to the phase transformation of AN near room temperature. Propellants which have a homing rate of $0.2{\sim}0.3\;mm/s$ and pressure exponent ranged from 0.3 to 0.6, showed stable burn-out in the standard motor tests. Both the magnitude of ignition energy and its thermal transfer mechanism have been proved to have a tangible effect on the ignition of the pyre starter, and the results of this study showed that a flame temperature of 1400K would be quite adequate to get a stable ignition for the AN composite propellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.7-10
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• 2009

The development and evaluation of solid propellant were performed for the turbopump pyro starter, which start up the liquid propellant rocket engine for the Space Launch Vehicle (SLV). Requirements for the turbopump pyro starter propellant include the production of low flame temperature, low burning rate and nontoxic gas to protect the mechanical corrosion or air pollution. This study describes the development of the solid propellant composition which is based on PCP binder. DHG (Dihydroxy glyoxime), which has advantages of oxygen balance and ignition, was used as coolant. The mechanical properties and burning rate of the propellants were measured. Finally, static fired test was performed to prove the possibility of development.