jffisher

This is the first launch of the Mark I Viper. The Mk I Viper cleared the rail guide at 2 m/sec and it was under thrust for ~ 7 seconds. I used 50 ml of unstabilized ~ 90% HTP plus 2 ml of ethanol for the oxidizer and PLA infused with KMnO4 for the fuel. Ignition occurred in ~ 0.3 seconds. The mass flow rate was about 13.29 gm/sec resulting in a total propellant mass of ~ 93.0 gm for the 7.0 seconds of thrust. I expected the Mk I viper to clear the rail guide at 4 m/sec. As such, I didn't have the aerodynamic control I needed to correct the trajectory. To prevent it from pitching forward, I'll shift the center of mass more in line with the thrust vector. Also, I'll increase the thrust to give me more aerodynamic control.

This is a static engine test of the MKI Viper flight system. Ignition occurs in ~0.3 sec, burn time is ~7 sec, thrust ranges from 15 N at ignition to 20 N at shut down. The short ignition time is due to a preheating event during initialization of the RC receiver. Read more in the April End of Month report at FisherSpaceSytems.com.

In an attempt to increase the thrust of the class I engine, I increased the O/F ratio by shortening the fuel grain from 15 cm to 12.5 cm. This reduced the contact surface area of the fuel grain and also, reduced the mass of the rocket engine by about 20 gm. Ignition occurred around 0.9 sec. I observed a net positive thrust of greater than 19.1 N at ignition and a c* efficiency of over 100%.

OOPS! I forgot to add a one kilogram mass to the rocket engine and it took off. The rocket ran into the load cell at the top, came off the rail guide, and landed on the ground. An accidental launch but still pretty cool. There was no damage to the rocket engine and diagnostics. After some recalibration, the system is up and running again.

The first two flight test with the new throat diameter of 6.7 mm did not go well! In the first test, the Mk I Viper made it to the top of the rail guide, blew out the nozzle, lost thrust, broke through the stop, came off the rail guide, and crashed onto the concrete pad. In the second test, the Mk I Viper left the rail guide at ~ 1.5 m/sec, began to pitch up, blew out the nozzle, flipped over once, and crash landed next to the test stand.

I used some similarity scaling parameters to increase the thrust of the HTPE/PLA hybrid rocket engine. The objective is to increase launch velocity and thus aerodynamic control of the Mk I Viper rocket glider. The three scaling parameters I looked at were initial gas flux, initial surface area flux, and length to diameter ratio. I designed and printed a higher flux fuel grain and inserted a variable PLA flow restrictor at the end of the fuel core. My average thrust went from ~19 N to ~23 N and overall performance was on par with previous test.

I printed a 15 cm PLA/Al fuel core using off the shelf aluminized PLA with an aluminum content of ~ 13%. I infused the PLA/Al fuel core with KMnO4 and assembled a rocket engine based on the PLA/Al/KMnO4 fuel core. All other parameters were the same. Ignition occurred in ~ 1.3 sec and burn time was ~ 5.8 sec. The c* was 1,477 m/s with a c* efficiency of 96%. This too looks promising!

In this test there was a peak combustion pressure of ~140 psia with a linear decline to ~110 psia and a peak thrust of ~35 N declining to ~24 N. I believe the peak pressure and thrust were due to a brief plug of PLA in the throat. This is clearly visible in the video of the test. Ignition occurred in about 1.3 sec with a burn time of 3.0 sec. Although I had a really good thrust, my c* efficiency and thrust coefficient efficiency was ~74% and ~66% respectively. However, my O/F ratio was ~7.5, way to high (theoretical at 2.75). Theoretically, a high oxidizer to fuel ratio leads to poorer performance.

For the past year, I've been using a 1/4" stainless steel mist nozzle with a 1.0 mm orifice for all test. With the 1.0 mm orifice, I get an initial oxidizer flow rate of ~ 14.8 ml/sec when pressurized to 140 psig. The same mist nozzle design with a 1.5 mm orifice gives me an ~ 23.0 ml/sec flow rate at 140 psig. Keeping all other parameters the same I got ~ 140 psig pressure spike in the mixing chamber. The pressure spike blew out the nozzle. Next test will have a variable flow control. This should reduce the pressure spike and produce a stable combustion.

I launched the Mk I Viper at a 30 degree angle. Gross liftoff mass was ~ 1.3 kg. Liftoff velocity was between 3 and 4 m/sec. I expected the Mk I Viper nose to pitch up after leaving the rail guide but it pitched down instead. In slow motion, I observed a small amount of rudder control. The Mk I Viper appeared to yaw a little to port

I upgraded my rocket engine test stand and ran two tests on the HF fuel core, one of which was on the new test stand. On second test, I increased the HTPE load to 62.4 ml. Ignition occurred in ~ 0.8 seconds and lasted ~ 5.0 seconds. The thrust was ~ 24 N and the c* efficiency was ~ 91%. Total propellant load was ~ 105 gm.

This month I tested a 13.5 cm high flux PLA/KMnO4 fuel core. The objective of this test was to increase the O/F ratio. Ignition occurred in ~ 1.3 sec with ~ 4.2 sec of thrust. From the video, I was able to deduce the mass flow rate. As such, the O/F ratio was ~ 4.0. The L/D was 8.8 and the initial surface flux was ~ 0.24 gm/cm2/sec. From the test, it appears as though the L/D is the dominate parameter as opposed to the initial surface flux.

I needed more thrust to exit the rail guide at a higher velocity. So, I increased the throat diameter to 6 mm. Not only did I get more thrust, but my characteristic velocity increased to 1662 m/sec. A few more test like this and I'll be ready to launch the Mark I Viper. Read the May 2022 end of month report at www.fisherspacesystems.com.

On Aug 11, 2023, I used the ignition surface flux scaling parameter (0.2 gm/cm2/sec) to designed a new 6-point star fuel core. The 6-point star configuration increased the surface area over the 5-point design. As such, the new fuel core can be made ~ 1.5 cm shorter than the 5-point design. The decrease in fuel core length will result in a mass savings for the overall flight system. Ignition occurred in 0.6 sec and burn time was 5.8 sec. The average thrust was ~ 24.0 N. The initial test results show a slight increase in performance over the 5-point design. The second part of the video shows the inner core of PLA/KMnO4 fuel core after the test. The recession of the PLA/KMnO4 fuel core seems to be evenly distributed along the length of the fuel core. This implies that, if designed properly, most of the fuel core can be consumed in the burn.

On Aug 11, 2023, I repeated the test from February 23rd but used a 500 ml Soda Stream® bottle as the pressure/propellant tank versus the 1000 mL tank. The objective was to test the performance of the rocket engine to see if there was a noticeable difference. I plan to use the 500 ml bottle in the rocket glider. All other parameters where the same as the 02/23 test. Time to ignition was 0.4 sec and burn time was 6.0 sec. The average thrust was ~ 22.1 N. There was no degradation of performance between the two tanks. The second part of the video shows the inner core of PLA/KMnO4 fuel core after the test. The recession of the PLA/KMnO4 fuel core seems to be evenly distributed along the length of the fuel core. This implies that, if designed properly, most of the fuel core can be consumed in the burn.

The objective of this test was to examine the effect of an optimum expansion ratio on the coefficient of thrust. The optimum expansion ratio at sea level is ~ 1.7. By cutting off ~ 4 mm from the end of the nozzle, I get an expansion ratio of ~ 1.8. I had a steady average thrust of ~ 19.3 N with a burn time of ~ 4.4 seconds. Ignition occurred in ~ 1.4 seconds. Although I had a really good thrust, my thrust coefficient efficiency was ~ 75%. However, my O/F ratio was ~ 5.2, way to high (theoretical at 2.75). So, there is room for improvement!