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Viper Rocket Trike
Updated Sept 2021
09-30-2021: There were five test in September (four succeeded & one failed). I varied the length of the PLA/KMnO4 fuel core as follows: 16.5 cm, 15.0 cm, 13.5 cm, and 12.0 cm. All other parameters were the same. I used a blend of 55 ml of ~ 87% HTP and 2.1 ml of denatured ethanol (O/F = 26.2) as the oxidizer. The propellant tank was pressurized to 130 psig using CO2 gas as the pressurant. I used a 1/4" stainless steel mist nozzle with a 1.0 mm orifice as the injector and a graphite phenolic nozzle with an initial throat diameter of 5 mm. The objective was to determine what effect the length of the fuel core had on the operation of the engine and to select the best length to continue. The ignition oxidizer flux of ~14 gm/cm2-sec, the run-time oxidizer flux of ~6 gm/cm2-sec, the fuel core regression rate of ~0.4 mm/sec, and the characteristic velocity of ~1390 m/sec were consistent on three out of the four successful test. The deciding factor was the oxidizer to fuel ratio, the thrust, and the burn time. For the 15 cm fuel core the O/F ratio was 2.3, close to theoretical. Ignition occurred in one second and lasted for ~7 sec. There was a net positive thrust of greater than 16.2 N at ignition and lasted throughout the burn. Based on these results, I've selected the 15 cm fuel core for the class I flight system. Read more in the September EOM report.
08-24-2021: Of the last five test in August 2021, the test on August 24 was the best. I increased the throat diameter to 5 mm, decreased the characteristic length to 33 in, increased the oxidizer tank pressure to 130 psig, increased the length of the fuel core to 16.5 cm, and added a pressure probe to the mixing chamber. Ignition occurred in 1.5 to 2.0 sec. The chamber pressure rose to ~93 psig in 2.0 sec and was steady throughout the ignition. Burn time was ~5 sec. The video shows a net positive thrust greater than 14 N (3.2 lb) at ignition and held throughout the burn time. Shut down was instantaneous. The oxidizer to fuel ratio was ~2.3 and total mass flow rate was ~13.4 gm/sec resulting in a characteristic velocity of 1,163 m/sec with a c* efficiency of ~77%. Read more in the August EOM report.
07-29-2021: In the lastest series of test, to shorten the ignition time, I blended ~85% hydrogen peroxide with ethanol using oxidizer to fuel ratios of 30 and 20. The first test baselined the series using just HTP/PLA/KMnO4 hybrid motor. The next test used an HTPE blend with a 30:1 ratio (i.e 1 ml of ethanol added to 30 ml of HTP). The third test used an HTPE blend with a 20:1 ratio (i.e. 1.5 ml of ethanol added to 30 ml of HTP). The HTP oxidizer was at 85% and the fuel grain, PLA infused with KMnO4, core was 2.3 cm in diameter, 15 cm long, and in the star configuration for all three test. Why would I add ethanol to 85% hydrogen peroxide?
Welllll! Several months ago I read a couple of articles on blended HTP as a monopropellant for micro satellites (Ref). In the articles, the authors stated that ethanol was chosen because it was somewhat compatable with HTP and significantly increased the specific impulse. In theory (using the NASA CEA program), adding a small amount (O/F = 30) of ethanol not only doubled the specific impulse but, also tripled the combustion chamber temperature (from 380 C to 1178 C). This necessitates the use of a more robust catalyst. In a catalyst pack based on silver, the silver would melt away quickly (silver melts at ~961 C) and in a catalyst pack based on platinum, the platinum would wash out (platinum melts at ~1770 C). And they cost a lot, platinum more so than silver.
"Simplicity is the ultimate sophistication."
- Leonardo da Vinci