Double-click here to edit the text.
Blog
Fisher Space Systems LLC
Viper Rocket Trike
Updated Jan 2022
12-31-2021: I began this month by making a wooden template to hand layup the carbon fabric for the Mk I Viper. I found that by laying Saran Wrap on the wooden template it was easier to remove the carbon fabric panel from the template. I cut three wooden panels, made some PLA mold brackets to hold everything together, and inserted some supports and a 1" CPVC pipe. I wrapped some Saran Wrap around the mold and hand laid up the carbon fabric in two layers. The mold and carbon fabric fuselage is shown on the left.
Next, I worked on the nose cone and forward strut. Originally, I tried to make the nose cone out of PLA. But, it look terrible and the mass was greater than 50 gm. I decided to go with a styrofoam nose cone. I cut a styrofoam block into a rough shape of a nose cone and then sanded down the edges until I reached a tapered shape resembling a nose cone. I glued some hinges on it and attached it to the forward strut. The forward strut is made of PLA and is secured to the fuselage by three small screws. When inserted into the fuselage, the propellant tank rest against the forward strut, pushing on the forward strut during take off. The forward strut and nose cone are shown on the left.
Finally, after Christmas, I worked on the aft strut. The aft strut centers the rocket engine and supports the fin assembly. I made three fins using the thinnest wall setting on the 3D printer. I glued some K'Nex pieces to the fins to hold everything together. The aft strut and fin assemble are shown on the left.
I made some plumbing modifications to the propellant tank, solenoid valve, check valve, injector, and rocket engine assembly. Before the modifications, the subsystem mass was ~ 930 gm. After the mods, the subsystem mass, shown on the left, is now ~ 575 gm, a savings of 355 gm.
The forward and aft cockpit struts are next. The forward cockpit strut supports the fuselage as well as a micro servo motor to open the cover and deploy the paraglider . The aft cockpit strut also supports the fuselage and three micro servo motors for the fins. The cockpit itself houses the battery, receiver, and paraglider. The total flight system mass is an estimated 1,292 gm, well below the 1.5 kg class I requirement. The mass allotment is summarized in the table below. Read more in the December EOM report.
11-30-2021: There were three test this month. In this series of test, 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.
A special thank you for the VIP observers on 11/10/2021. I've run enough test to feel confident of the safety in the HTP/PLA hybrid. So, Fisher Space Systems, LLC played host to a group of future scientist and engineers. After a brief safety overview, I loaded the oxidizer, ran the test, and everything was perfect. They, of course, are invited back this summer to observe the launch and landing of the Mk I Viper. Read more in the November EOM report.
10-31-2021: 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 ignitor. Eliminating the ignitor 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 ratio 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. Read more in the October EOM report.
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.
Next, I worked on the nose cone and forward strut. Originally, I tried to make the nose cone out of PLA. But, it look terrible and the mass was greater than 50 gm. I decided to go with a styrofoam nose cone. I cut a styrofoam block into a rough shape of a nose cone and then sanded down the edges until I reached a tapered shape resembling a nose cone. I glued some hinges on it and attached it to the forward strut. The forward strut is made of PLA and is secured to the fuselage by three small screws. When inserted into the fuselage, the propellant tank rest against the forward strut, pushing on the forward strut during take off. The forward strut and nose cone are shown on the left.
Finally, after Christmas, I worked on the aft strut. The aft strut centers the rocket engine and supports the fin assembly. I made three fins using the thinnest wall setting on the 3D printer. I glued some K'Nex pieces to the fins to hold everything together. The aft strut and fin assemble are shown on the left.
I made some plumbing modifications to the propellant tank, solenoid valve, check valve, injector, and rocket engine assembly. Before the modifications, the subsystem mass was ~ 930 gm. After the mods, the subsystem mass, shown on the left, is now ~ 575 gm, a savings of 355 gm.
The forward and aft cockpit struts are next. The forward cockpit strut supports the fuselage as well as a micro servo motor to open the cover and deploy the paraglider . The aft cockpit strut also supports the fuselage and three micro servo motors for the fins. The cockpit itself houses the battery, receiver, and paraglider. The total flight system mass is an estimated 1,292 gm, well below the 1.5 kg class I requirement. The mass allotment is summarized in the table below. Read more in the December EOM report.
11-30-2021: There were three test this month. In this series of test, 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.
A special thank you for the VIP observers on 11/10/2021. I've run enough test to feel confident of the safety in the HTP/PLA hybrid. So, Fisher Space Systems, LLC played host to a group of future scientist and engineers. After a brief safety overview, I loaded the oxidizer, ran the test, and everything was perfect. They, of course, are invited back this summer to observe the launch and landing of the Mk I Viper. Read more in the November EOM report.
10-31-2021: 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 ignitor. Eliminating the ignitor 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 ratio 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. Read more in the October EOM report.
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.
"Simplicity is the ultimate sophistication."
- Leonardo da Vinci
This work is licensed under the Creative Commons Attribution-ShareAlike 4.0 International License. To view a copy of this license,
visit
http://creativecommons.org/licenses/by-sa/4.0/
Copyright (c) 2021 Jerry F Fisher
http://creativecommons.org/licenses/by-sa/4.0/
Copyright (c) 2021 Jerry F Fisher
| Subsystem | Mass (gm) |
|---|---|
|
| |
| Prop Tank, valves, rocket engine with fuel |
575
|
| Forward strut and nose cone | 42 |
| Cockpit forward strut (includes servo motor) | 40 (est.) |
| Cockpit (includes batt, receiver, & paraglider) | 130 (est.) |
| Cockpit aft strut (includes three servo motors) | 70 (est.) |
| Aft strut and fins | 70 |
| HTPE Oxidizer | 75 |
| Carbon Fabric Fuselage | 290 |
| Total System | 1,292 (est.) |
