Combustion characterization of ammonium perchlorate based solid rocket propellant
There were many studies on ammonium perchlorate (AP) based solid rocket propellant (SRP) and some data already exist. However, there is no complete data available in Malaysia especially the information on the combustion characteristics of baseline data of AP based SRP. This thesis discusses on the c...
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| Format: | Thesis |
| Language: | English |
| Published: |
2011
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| Subjects: | |
| Online Access: | http://eprints.utm.my/id/eprint/16929/1/AmirAzizMFKM2011.pdf |
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| Summary: | There were many studies on ammonium perchlorate (AP) based solid rocket propellant (SRP) and some data already exist. However, there is no complete data available in Malaysia especially the information on the combustion characteristics of baseline data of AP based SRP. This thesis discusses on the combustion characteristics of AP based SRP including the methods of the propellant selection and fabrication, burning rate test and static thrust test. Together with literature study and theoretical performance, thirteen sets of different propellant mixture were finalized with consideration of the mechanical and processability factors. The propellant was a mixture of AP as an oxidizer, Aluminum (Al) as fuel and Hydroxy-Terminated Polybutadiene (HTPB) as binder/fuel. For each mixture, HTPB binder was fixed at 15% and cured with Isophorone isocyanate (IPDI) (9.33% per mass of HTPB). The percentage of the solid materials was set at a constant value of 85%. By varying the AP and Al, the effect of oxidizer-fuel (O/F) ratio on the whole propellant can be determined. The propellant strands were manufactured using press-molding method and burnt in the strand burner at ambient pressure to obtain the initial burning characteristics. Then four propellant compositions were selected, namely p60, p66, p74 and p80 for further evaluation over a range of pressure from 6 atm to 31 atm. The results show that the increasing of O/F ratio and combustion pressure lead to the increase in burning rate. The fastest burning rate achieved was 12 -1 mmsec at combustion pressure of 31 atm for propellant p80 which has O/F ratio of 4.0. It was found that, the formulated propellant in this study have the normal burning characteristics with pressure exponent lies within the range of 0.501 to 0.561. Based on theoretical evaluation, formulation for p66 gives highest specific impulse, sp I . Thus, p66 has been selected to be evaluated in static thrust testing to obtain its performance characteristics. The results showed that the maximum thrust obtained is 162 N with generating sp I of 143.92 sec |
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