• Journal of the Korean Society of Propulsion Engineers
• /
• v.4 no.4
• /
• pp.50-58
• /
• 2000

Burning rate of the matrix propellant($r_{sb}$) and burning rate along the metal wire($r_w$) were measured and analyzed for the HTPB/AP/Al propellant embedded with Ag wire($\phi$0.15mm) according to weight % of RDX(0~20%). Variation of burning rate increment ratio($r_w$/$r_{sb}$) and pressure exponent(n) was studied for the nitramine propellant having 10% RDX embedded with three kinds of metal(Ag, Cu, and Ni-Cr) of which diameter range is between 0.1~0.6mm. Maximum burning rate increment ratio of the nitramine propellant embedded with Ag wire($\phi$0.1mm) was 5.94 at $20^{\cire}C$, 1000 psia, 16.4% faster than that of HTPB/AP propellant, it is because that autoignition temperature of the nitramine propellant was higher than that of HTPB/AP propellant. Standard deviation of absolute ($r_{wc}$/$r_{we}$)/$r_{we}$ calculated by using new empiracal equation composed of four dimensionless groups, is 6.11% less than that calculated by using original empirical equation composed of three dimensionless group. The new empiracal equation is derived from Buckingham pi theorem using the parameters such as thermal diffusivity, melting temperature. wire diameter, propellant sample diameter, frame temperature, autoignition temperature and matrix burning rate which influence on $r_w$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.508-512
• /
• 2017

In this study, we carried out the study on the combustion characteristics of HTPB/AP propellants with Al and Zr as fuel metal in order to develop the solid propellant with high burning rate. The major combustion characteristics of propellant were investigated as measuring of the burning rate and pressure exponent, and the HTPB/AP solid propellants were prepared with introducing Butacene as burning rate catalyst for the enhancement of burning rate. The propellant with Al and Zr was demonstrated the improvement of propellant performance and combustion characteristic.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.10 no.4
• /
• pp.77-84
• /
• 2006

With the development of micro/nano spacecraft, concepts of micro propulsion are introduced for orbit transfer and drag compensation as well as attitude control. Micro solid propellant thruster has been attention as one of possible solution for micro thruster. In this paper, micro solid propellant thruster is introduced and research on basic components of a micro solid propellant thruster is reported. Micro Pt igniter was fabricated through negative patterning and quantitative effect of geometry was estimated. The characteristic of HTPB/AP solid propellant was investigated to measure the homing velocity. A combustion chamber was fabricated by means of anisotropic etching of photosensitive glass. Finally, micro solid propellant thrusters having various geometries were fabricated and tested.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.04a
• /
• pp.351-355
• /
• 2011

In this study, 2 kinds of HTPE insensitive propellants composed of HTPE/BuNENA binder, AP, AN and Al were investigated for combustion characteristics, ignition delay time, sensitivity and insensitive properties compared with HTPB propellant. HTPE propellant showed almost same sensitivity results as HTPB propellant, showed 2~3 times higher value than the value of HTPB propellant, ignition delay time respectively, and met the standard criteria, while HTPB propellant failed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 1996.05a
• /
• pp.121-127
• /
• 1996

To compare the performance of aluminized solid propellants, the theoretical calculation was performed for the propellants using HTPB and PEG binder and four kinds of oxidizers such as AP, HMX, ADN, and HNIW. PEG/HMX/Al and PEG/HNIW/Al showed the maximum performance at 17% of aluminum level and there was no difference in maximum performance when HMX was partially replaced with AP in PEG/HMX/AP/Al propellant. The order of performance magnitude of various propellants which the specific impulse loss calculation was considered by semi-empirical equation was like the following; PEG/HNIW/AI>[$\frac{PEG/HMX/AI}{PEG/HMX/AP/AI}$>HTPB/AP/AI>PEG/ADN/AI

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.2
• /
• pp.39-45
• /
• 2013

Basically, in order to apply solid propellant as ignition source to high energy metal particle combustion system, we analyzed combustion characteristics of the HTPB/AP/Al, HTPE/AP/Al propellants by using a strand burner. The propellants were tested in a high-pressure windowed strand burner, which was pressurized up to 300 psia with pure argon gas. Strand burner was visualized with two quartz windows and ignition was accomplished by a 10 W $CO_2$ laser. The burning rate of propellant was measured with high-speed camera method for frame analysis and photodiode method for combustion time analysis. Emission spectrum was measured with spectrometer at 300 nm ~ 800 nm and 1500 nm ~ 5000 nm and then we analyzed species during propellant combustion.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.3 no.1
• /
• pp.41-47
• /
• 1999

The burning rate of solid propellant is one of the key parameter associated with the dynamic characteristics of combustion and the combustion performances. In the AP propellants, the evaporation on the reacting surface as well as the decomposition of the propellant is of great importance in determining the overall burning rate. In this study, a theoretical analysis was conducted to obtain the expression for burning rate in the steady state combustion with the energy and species equations in the condensed phase when the radiative heat flux partially contributes to the total heat transfer to the propellant surface.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.175-178
• /
• 2008

A Numerical Simulation which uses preconditioning algorithm to examine unsteady combustion processes for the AP/HTPB propellant with a converging-diverging nozzle has been compared with experimental data for solid rocket motor. To analysis reacting flow of solid rocket motor, unsteady pressure, temperature contour was simulated by grid moving of propellant.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.37 no.8
• /
• pp.774-779
• /
• 2009

A numerical modelling was performed to predict unsteady combustion processes for the AP/HTPB/Al propellant in a solid rocket motor. Its results were compared with the experimental data. Temporal pressure development was found to match quite well with measured data. A change in propellant surface was traced using the moving grid. The propellant thickness change was also observed to confirm the erosive burning effect.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.191-195
• /
• 2009

In the solid rocket propellant combustion, dynamic phase change from solid to liquid to vapor occurs across the melt layer. During the burning surface, micro scale bubbles form as liquid and gas phases are mixed in the intermediate zone between the propellant and the flame. The experimentally measured thickness of this layer called the foam layer is approximately 1 micron at 1 atmosphere. In this paper, we present a new melting layer model derived from the classical phase change theory. The model results show that the surface of burning grows and propagate uniformly at a velocity of $r=ap^n$.