HTP/PLA/KMnO4 Hybrid Rocket Engine

[Jerry Fisher]

Solaire plans used dry ice to pressurize the fuel and HTP

Thanks for the tip. I didn't realize the vapor pressure of dry ice was that high (~830 psi @ 20 C). I will be experimenting with it for the class II engine. I was going to use my porous ceramic mixed metal oxide (that I've been working on for SEVEN YEARS) with 70% HTP as the pressurizing subsystem but, dry ice seems a lot simpler. However, there are questions of mass and duration. Does it weigh less than the HTP subsystem and can it sustain a pressure for greater than 60 seconds of run time?

The flight system is a work in progress. I'm working on a lightweight air-frame for the class I engine. I used fiberglass panels but the mass is to great (~450 gm). Ordering some carbon fiber cloth to try that next. For flight control, I'm using an RC transmitter, receiver, and spare servos.

Keep those tips and comments coming.

[Reuben]

How would the co2 be vaporised? with liquids they can be ran through some sort of cooling jacket for the engine but this obviously can't be done with ice. or would it be stored above boiling temperature and just at pressure, where it could sublimate when more pressurant is needed, and the tank pressure is reduced?

[Jerry Fisher]

How would the co2 be vaporised?

I did a quick search on sublimation rate of dry ice. Most research is done for the packing and transportation industries at room temperature and atmospheric pressure. Very little research at 830 psi and room temperature. I looked at the Clapeyron-Clausius Equation but it contains an experimental coefficient. So, no matter what equation you use, some research is required.

I plan on putting some dry ice in a stainless steel container, letting it come up to pressure, and measuring the change in pressure as I release the vapor through a stainless steel orifice. The diameter of the orifice will depend on how much gas I need to keep my tank pressurized. Once I have the coefficient, I can scale up to larger systems. There are additives such as ammonia to increase the sublimation rate if required. 

[aceshigh]

How would the co2 be vaporised?

I did a quick search on sublimation rate of dry ice. Most research is done for the packing and transportation industries at room temperature and atmospheric pressure. Very little research at 830 psi and room temperature. I looked at the Clapeyron-Clausius Equation but it contains an experimental coefficient. So, no matter what equation you use, some research is required.

I plan on putting some dry ice in a stainless steel container, letting it come up to pressure, and measuring the change in pressure as I release the vapor through a stainless steel orifice. The diameter of the orifice will depend on how much gas I need to keep my tank pressurized. Once I have the coefficient, I can scale up to larger systems. There are additives such as ammonia to increase the sublimation rate if required.

BASE: we have some good news and bad news for you

The bad news is that the rocket and the ship that would save you had a problem. They wonīt be able to take off before you die from hunger and thirst OR lack of oxygen

Stranded astronaut: damn... and what is the good news?

BASE: we discovered the problem with your ship. A simple cable got disconnected. We will be sending the solution to your computer. Then you can return home. But beware, your velocity will be higher than normal on reentry, but donīt worry, the cookie tiles are much better than old ceramic tiles. Haha, of course, unless one of them is missing. You didnīt eat any, have you?

Stranded Astronaut: well, at least from HUNGER I wonīt be dying.

[Jerry Fisher]

   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 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. https://rumble.com/voo0fx-plakmno4-high-flux-test.html

[Reuben]

Do you know what the effect of a longer residence time of the HTP on the ignition delay? perhaps prior to ignition, the HTP injection pressure could be lowered to allow it longer to catalyse whilst not wasting all that isn't combusted. otherwise watching that video was very satisfying the combustion looked very very consistent- awesome build!

[Jerry Fisher]

Do you know what the effect of a longer residence time of the HTP on the ignition delay?

I'm not sure there would be an advantage here. Lowering the propellant tank pressure to below 120 psig introduces oscillations in the combustion chamber. I need at least a 30 psig pressure drop across the 12V NC solenoid valve, the check valve, and the injector nozzle to dampen out the oscillation. Also, the lower chamber pressure results in a lower thrust and characteristic velocity. Finally, the mechanics of starting out with a lower tank pressure then ramping up to a higher pressure would add more mass to the class I flight system. This may be feasible in a class II system but not in a class I system. Thanks for the comment.

[CameronD]

Do you know what the effect of a longer residence time of the HTP on the ignition delay?

I'm not sure there would be an advantage here. Lowering the propellant tank pressure to below 120 psig introduces oscillations in the combustion chamber. I need at least a 30 psig pressure drop across the 12V NC solenoid valve, the check valve, and the injector nozzle to dampen out the oscillation. Also, the lower chamber pressure results in a lower thrust and characteristic velocity. Finally, the mechanics of starting out with a lower tank pressure then ramping up to a higher pressure would add more mass to the class I flight system. This may be feasible in a class II system but not in a class I system. Thanks for the comment.

How about keeping tank pressure the same and using a slowly-opening ball valve instead of the solenoid?  Wouldn't that have the same/similar effect??

Impressive work, BTW!

[edzieba]

Or printing a flow-restrictor as part of the combustion chamber that rapidly burns away just after ignition.

[Jerry Fisher]

...a slowly-opening ball valve instead of the solenoid?

Yes. It would have a similar effect but, the ball valve weighs ~3 times as much as a solenoid valve (345 gm vs 102 gm). I will be using a ball valve on the class II engine. The ball valve I have is rated at 150 psig. I will be testing at higher pressure and throttling up after ignition on the class II engine.

Or printing a flow restrictor....

Good idea! I could print a PLA flow restrictor and place it at the end of the mixing chamber. After ignition, the orifice diameter will increase and the graphite nozzle throat diameter would take over. The PLA flow restrictor would not add much mass to the system. I would start with an orifice diameter of 4 mm (vs 5mm nozzle throat diameter) and work my way to smaller orifice diameters. I'll have to be careful and not plug the nozzle with PLA melt. That has happened before with resulting explosion.

[Reuben]

Do you know what the effect of a longer residence time of the HTP on the ignition delay?

I'm not sure there would be an advantage here. Lowering the propellant tank pressure to below 120 psig introduces oscillations in the combustion chamber. I need at least a 30 psig pressure drop across the 12V NC solenoid valve, the check valve, and the injector nozzle to dampen out the oscillation. Also, the lower chamber pressure results in a lower thrust and characteristic velocity. Finally, the mechanics of starting out with a lower tank pressure then ramping up to a higher pressure would add more mass to the class I flight system. This may be feasible in a class II system but not in a class I system. Thanks for the comment.

I am aware that a 30% pressure decrease is healthy for the negation of oscillation between the fuel tank and the combustion chamber, however that oscillation only arises ad a product of combustion, and would not begin until ignition had occurred, and chamber pressure could be raised again. The point you brought up about weight however does sound prohibitive.

[Jerry Fisher]

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 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. Link to Nov 10, 2021 test https://rumble.com/vqb7of-htppla-hybrid-reliability-test.html. Read more in the November end of month report.

[CameronD]

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 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. Link to Nov 10, 2021 test https://rumble.com/vqb7of-htppla-hybrid-reliability-test.html. Read more in the November end of month report.

Looks good, Jerry!  Seems fairly stable in your video so you might be onto something with this.  Keep up the great work!

[Jerry Fisher]

This month I made a wooden mold for the fuselage. I wrapped some Saran Wrap around the mold and hand laid up the carbon fabric in two layers. The carbon fabric fuselage has a mass of ~ 290 gm.

For the 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. I attached the nose cone 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 aft strut centers the rocket engine and supports the fin assembly. The aft strut is also made of PLA and is secured to the fuselage by three small screws. 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.

I made some plumbing modifications to the propellant tank, solenoid valve, check valve, injector, and rocket engine assembly reducing its mass to ~ 575 gm.

Next will be the cockpit struts and cockpit. The total flight system mass is an estimated 1.3 kg (includes the fuel and oxidizer), well below the 1.5 kg class I requirement. Read more in the December, 2021 EOM report.

[Jerry Fisher]

This month I worked on the forward and aft struts for the cockpit, the paraglider box, the battery pack, and the placement of the RC receiver and the servos. I inserted the oxidizer tank and fuel grain (pictured below) into the carbon fabric fuselage. Everything fits and is ready for static testing. The mass is ~1.3 kg including the oxidizer and fuel. Read more in the January 2022 end of month report.

[Jerry Fisher]

This month, 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 caused the rocket engine to run oxidizer rich. Also, it 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%. https://rumble.com/vwtclh-htppla-hybrid-oxidizer-rich.html

Try adding metal powder to your fuel and you might get better thrust.

Also this month, as suggested by colbourne (reply #14), 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.  If ignition time is reduced by running the engine oxidizer rich as noted above, it may be worth pursuing. Also, using a finer grain of Al with the PLA may improve performance. https://rumble.com/vwtcy9-htpplaal-hybrid.html

[Jerry Fisher]

This month, I mounted a static engine test stand, a load cell to measure thrust, and a test article with a blast shield around the oxidizer tank to my rocket engine test stand. To run a static engine test, I need to pressurize the oxidizer tank to 140 psig and turn off the fill valve. Adding a blast shield around the oxidizer tank makes it safer and is part of my KISSES principle of "Keeping it Safe". Over the next month or two I'll be testing fuel grain lengths of 11.5, 12.5, and 13.5 cm to narrow done the ideal fuel grain lengths for best overall performance.

[Jerry Fisher]

This month, I did a static engine test of the flight system. A video of the test is at https://rumble.com/v13dqgl-mki-viper-static-engine-test.html. The thrust ramped up from ~15 N to around 20 N over a 7 sec burn time. The longer than usual burn time is attributed to ignition occurring in ~0.3 sec and thus leaving more oxidizer to fuel the engine. This was due to a small amount of HTPE leaking into the fuel grain during initialization of the receiver, a preheating event. 

The center of mass (CM) is below the cockpit area and ~ 2 cm off the thrust vector. The off axis thrust will cause the Viper to rotate after leaving the guide rail. Due to the low thrust at ignition, the liftoff velocity will not be enough to correct the rotation of the MkI Viper using the elevators and the vertical stabilizer. I will rearrange the masses inside the fuselage and bring the CM closer to the line of thrust. Also, I believe I can increase the thrust by increasing the initial nozzle throat diameter from 5 to 6 mm.

Also, during my first test (with a 12.5 cm fuel grain) on the static engine test stand, I forgot to put a one kilogram mass on the rocket and it took off. I definitely need a pretest checklist. The liftoff was pretty cool. A video of the test is available at https://rumble.com/v13dp0n-liftoff.html.

[CameronD]

Also, during my first test (with a 12.5 cm fuel grain) on the static engine test stand, I forgot to put a one kilogram mass on the rocket and it took off. I definitely need a pretest checklist. The liftoff was pretty cool. A video of the test is available at https://rumble.com/v13dp0n-liftoff.html.

Well, there can be no arguing with actual results.. Well done, Jerry!

[Jerry Fisher]

Well, there can be no arguing with actual results.. Well done, Jerry!

Thanks! Although it was an accidental launch, I got some useful data. That contraption had a mass of 1.9kg. As such, there was a net positive thrust greater than 19N and it sure was pretty. The throat diameter eroded to 6 mm.

As a result of the previous test, I increased the initial throat diameter from 5 to 6 mm which resulted in a slight increase in mass flow rate. I measured the thrust at 19.1 N and it looked pretty steady rumble.com/v17imbi-20-n-thrust-in-may-2022.html.  The characteristic velocity was calculated to be ~1662 m/sec with an efficiency of 104% (theory @ 1593 m/sec).

Also, I've been working on the flight system. I have rearranged the mass some with the purpose of moving the CM more in line with the thrust vector. Next month, I plan to test for consistency and reliability and launch the Mark I Viper.