After studying the theoretical framework behind LPRE’s it became evident that the P-5 would require a systematization different from that of more complex high-power rocket engines. The following systems were chosen and characterized separately through MEP’s and testing.
Figure 2. Pie chart showing the six engineering systems
The propulsive system refers to the combustion chamber and nozzle design. This includes choosing the propellants, the chamber pressure, and the desired thrust level. To make these decisions, trade studies were made in July of 2018 with a sample objective in mind: The P-5 engine must be simple enough to be economically and logistically feasible, but complex enough to explore the underlying technicalities involved in low power liquid rocket engines. Therefore, very advanced features such as: cryogenics, hypergolics, turbopumps, thrust vectoring, throttling, high chamber pressures/ thrust levels, etc.; were avoided. Moreover, most of the aerospace rated components are significantly expensive, and therefore it was decided to use commercially available valves and regulators. This set a limit to the maximum operating pressure of the engine, and thus an idea of the mass flow rates that could be achieved (which is proportional to thrust). Perhaps the most difficult decision was the oxidizer choice, since the only non-cryogenic ones (i.e. nitric acid [and derivatives], nitrous oxide, and hydrogen peroxide) are difficult to handle and obtain. Finally, due to its liquid state at STP and commercial availability in Costa Rica, hydrogen peroxide (50% w/w) was chosen. It was clear that the low concentration of the peroxide would lower the engine’s power, but most concerningly, there was a chance the water content would be high enough to not even support combustion. In August 2018, a sample of this oxidizer was bought and mixed with different fuels and was proven to support combustion (further confirmed one year later). The geometry and thermochemistry of the engine was handled with an inhouse software called the Rocket Engine Designer (RED) which links with NASA’s CEARUN to solve for an ideal rocket.
Figure 3. General specs of the P-5 engine original design
Figure 4. Evolution of the propulsive system
The injection system consists of the design and prototyping of the injector which is often considered the heart of the engine. G. S. Gill, and W. H. Nurick [1] present an injector design algorithm for four different injector element types, of which three were selected, designed, and evaluated with a MEP. Afterwards, instances of both injector B and C were manufactured and cold flowed if both Purdue’s Zucrow Propulsion Labs, and in AREX’s Albireo Low Power Propulsion Lab in Costa Rica (see “Testing the Engine”). It was then determined that pintle injectors (B) are not very scalable for engines with a small chamber diameter. Therefore, injector C which was chosen and further improved based on cold flow data.
Figure 5. CADs of injectors considered for the P-5 engine
Table 1. Specs of the injectors
The feed system consists of the tanks, valves, regulators, propellant lines, and all other components required to deliver the propellants to the engine. The system is pressurized with carbon dioxide and in-line regulators allow achieving relatively constant mass flow rates. These regulators and some valves must be manually adjusted, but afterwards, the engine can be controlled electronically at distance. Two stainless steel 2.1-liter tanks were manufactured locally, and the rest of the components were bought from commercial suppliers according to the system’s requirement. The maximum operating (safety) pressure was set to 200 psig.
Figure 6. Final version of the Piping and Instrumentation Diagram of the feed system.
The electrical system includes the wiring and programming of all electronic components. A software called the Rocket Engine Testing Center (RETC) was programmed to control the feed system both manually and automatically, as well as to log data. This software ran on a Raspberry Pi connected to a computer 30 meters away through an ethernet cable. An Arduino UNO was used to record data from sensors such as pressure transducers, thermocouples, and flow meters. The whole system was powered by individual DC power supplies.
Figure 7. A screenshot of the RETC software in manual opperation mode
The structural system refers to the physical structure that holds the engine static, and the space used to arrange all the systems together. A test stand/ bench was designed such that engine could fire vertically, and the components could be stored safely.
Figure 8. CAD design of the test bench.
The cooling system refers to the method used to remove heat from the engine. It was decided to keep the engine uncooled because of the following reasons. First, the theorized temperature of the combustion gases was relatively low compared to other rocket engines (below 2000 K) due to the low concentration of the oxidizer and low chamber pressure. Second, a heat transfer simulation for a three second fire of a much powerful conceptual stainless-steel engine with a 4 mm chamber thickness was made before and proved the engine to withstand much higher temperatures. Third, thanks to the IMEP, the final version of the engine ended up having an immense safety factor due to the chamber thickness going up from 4 mm to 35 mm for more than 62% of the chamber length.