FisherSpaceSystems

This is a test of the MkI Viper canopy release mechanism. The canopy has four post which compress four springs mounted on the frame. The canopy is held into place by Kanthal A-1 wire. By remote control, a high current passes through the wire melting a small section of the wire causing a break. The canopy is ejected from the frame. The canopy and mounting frame have a mass of ~125 gm.

This month, in an effort to get maximum thrust with minimum hardware, I ran two tests without the check valve and with a 7.5 mm nozzle throat diameter. The HTPE tank pressure was 140 psig. The first test streamed PLA out the nozzle and the second test blew up. Calculations show that the ignition surface flux (ISF) was ~ 0.3 gm/cm2/sec as compared to ~ 0.2 gm/cm2/sec on previous test using the check valve. Last month, the two successful test without the check valve was with the same ISF of ~ 0.3 gm/cm2/sec but with a nozzle throat diameter of 6.0 mm. I surmise that the smaller throat diameter slowed the flow rate enough for ignition to occur while the larger throat diameter did not. After careful consideration, I have decided to add the check valve back into the plumbing.

This month, I ran two tests without the check valve. As expected, there was an increase in both the chamber pressure and the thrust. Ignition time was about 0.3 sec and burn time was about 6.0 sec for each run. The average chamber pressure was ~ 107 psia. The initial thrust was ~ 22 N and increases to ~ 28 N at the end of the run with an average thrust at ~ 25 N.

This test examined the ignition time of the PLA/KMnO4 fuel core after weeks of storage in a dry bag. Ignition was in ~ 0.6 sec with a burn time of ~ 6.0 sec. The shorter the ignition time the longer the burn time for a given amount of propellant. The total propellant mass is ~ 115 gm, just under the FAA regulation for a class I rocket. The fuel core was in a zip lock bag with desiccant for 24 days. There were three test, the first was stored in dry bag for ~ 6 months, the second for 24 days, and the third for 31 days. All other parameters were the same. Ignition times varied from 0.4 sec to 0.6 sec. As such, The procedure is to store PLA/KMnO4 hybrid fuel cores in a dry bag for a minimum of two weeks and as long as four weeks. The endoscope video shows a nice even burn of the fuel core. The video begins at the injector end and backs out to the exit. The six point design is clearly visible. The injector evenly distributes the spray around the fuel core. On closer examination, one can see the cone angle of the spray. Backing out, there is an even burn distribution. The purplish color is due to the reflection from the LED light on the endoscope, not the color due to the fuel. Reference the blog on www.fisherspacesystems.com for end of month report.

This month I showed that the class I engine performance was consistent, reliable, and ignition occurs in ~ 1.1 sec. The parameters were the same for each of the three test: propellant tank pressure, 130 psig; HTPE blend O/F ratio, 27.5, initial HTPE flow rate, 14.8 ml/sec; mass flow rate for HTP and ethanol, 19.7 gm/sec and 0.4 gm/sec, respectively; cross sectional area for the fuel cores, ~1.1 cm2 ; and the ignition "oxidizer" flux, ~ 17.6 gm/cm2 /sec. All three test used the same batch of distilled HTP with 2.0 ml of ethanol. The thrust was ~ 16.5N. The results were about the same across all three test. The average: ignition, 1.1 sec; mass flow rate, 12.4 gm/sec; chamber pressure, 105 psia; characteristic velocity, 1280 m/sec; efficiency, 86%, thrust, 16.5 N; regression rate, 0.23 mm/sec; O/F ratio, 3.0.

There were two test in October. I eliminated the glow wire for ignition and decreased the cross sectional area of the 15 cm PLA/KMnO4 fuel cell to increase the oxidizer flux. All other parameters were the same. I used a blend of 55 ml of ~ 85% HTP and 1.7 ml of denatured ethanol (O/F = 37.4) as the oxidizer. 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 increased flux had on the operation of the engine and if auto ignition would occur without the glow wire. The ignition oxidizer flux was ~14 gm/cm2-sec and was the same for both test. Ignition occurred in ~1.9 sec for the low flux fuel core and ~1.5 sec for the high flux fuel core. The ignition times are about the same as with a glow wire igniter. Eliminating the igniter simplifies the system. The run-time oxidizer flux was ~5.4 gm/cm2-sec for the low flux fuel core and 9.1 gm/cm2-sec for the high flux fuel core. The fuel core regression rate and O/F ration was approximately the same in both test. The deciding factor was the chamber pressure and the characteristic velocity. The propellant tank was pressurized to 130 psig using CO2 gas as the pressurant in both test. The low flux test had a higher chamber pressure with corresponding higher characteristic velocity with a c* efficiency of ~91%. Based on these results and despite the longer ignition time, I've selected the low flux 15 cm fuel core for the class I flight system.

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.

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%. I still plan on launching a class I HTPE hybrid before the end of the year. Next month I'll lock down the thrust and begin building the flight system.

This is the first test of ~85% hydrogen peroxide blended with ethanol. In the test, I used porous ceramic cylinders infused with mixed metal oxides and sintered at 600C. The test shows ignition in about 9 sec followed by combustion for another 9 sec. The initial flow rate was 1.7 ml/sec at 120 psia. At the end of the test, there was burn through at the injector. I recovered three out of the four cylinders and reused them in a follow on test with an initial flow rate of 5.4 ml/sec but got no ignition.

This is a repeat of the 01-27-2021 test but with a 1/4" stainless steel mist injector with a 1.0 mm diameter orifice, a 3.0 mm fuel core wall thickness, L/D=12.5 (vs 2.0 mm, L/D=10.5 in 01-27 test), and 85% HTP (vs 90%). Combustion instabilities seem to be under control due to a higher pressure drop across the injector. The net positive thrust occurs at the end of the video corresponding to the erosion of the graphite nozzle throat from 3.6 mm to 5.2 mm. The increase in throat diameter reduces chamber pressure and increases mass flow rate resulting in a net positive thrust. Net positive thrust occurred for ~2 sec at a thrust of greater than 12 N. PLA fuel core segments were infused with potassium permanganate at high temperature and pressure.

This is a test of 3, 3D printed poly lactic acid (PLA) fuel core segments with 20% hexagon infill infused with potassium permanganate (KMnO4) at high pressure and temperature. Each segment is 5 cm long and 1.7 cm in diameter making the total length 15 cm. The mixing chamber is preheated for about 10 sec using a high resistance glow wire. Ninety percent hydrogen peroxide (HTP) is introduced into the fuel core under a pressure of 120 psi by opening a normally closed solenoid valve. The KMnO4 catalyzes the HTP producing the initial temperature and pressure for the thermal environment. Once the high temperature environment is established, thermal decomposition of HTP continues.

This video is an introduction to Fisher Space System, LLC and the Viper Rocket Trike including a short clip on the HTP/PLA rocket motor. I believe the time is right for a privately owned and operated spacecraft. It will have to be safe, reliable, and inexpensive. The Viper Rocket Trike is a single place ground launched VTHL rocket glider. It is powered by a pressure fed hydrogen peroxide/polylactic acid (HTP/PLA) hybrid rocket motor. At altitude, it deploys an inflatable paraglider for return to launch site and will have a backup parachute for emergencies. The pilot will experience three minutes of thrust, several minutes of zero gravity, and 14 minutes of glide. The pilot will have full control of the launch, free fall, glide, and landing.