Hybrid engine subsystems
To differentiate areas of responsibility, the hybrid engine is separated into three so-called subsystems: the OX system, the burn chamber and the filling station. You can read more about them below. Coming soon, we will upload the full system documentation for each subsystem for you to download!
The OX, or oxidizer, subsystem of the engine is responsible for delivering liquid N2O to the burn chamber. It consists mainly of a tank to hold the N2O and a system of valves and pipes to allow for filling/”fueling” (although of course, the fuel is already in the burn chamber), delivery to the burn chamber, and some auxiliary functions such as an emergency vent valve. The weight of fuel systems relative to the payload is pretty much the crux of rocket engineering, so making a lightweight and sufficiently small OX system to safely handle the N2O pressure of 60 Bar while also fitting inside the fuselage is no easy task.
The OX system is mostly designed around the properties of N2O: The tank diameter needs to be wide enough to allow for continuous evaporation to maintain pressure while firing, and the temperature needs to be controlled so that the tank is neither too hot or too cold. N2O goes supercritical when it reaches 36 C, at which point the density drops and it runs the risk of decomposing into N2 and O2. Decomposition is an exothermic process, and could lead to the tank detonating. Conversely, too low a temperature will yield a pressure drop which will affect the engine’s performance somewhat unpredictably - it could fizzle out or it could result in a “blowback” where the fire from the burn chamber travels upline into the tank again resulting in detonation. These parameters must be tracked from a safe distance, and so the OX tank has a pressure transducer and multiple temperature sensors installed on its surface, with the data transmitted to the ground station computer.
For practical reasons, most of the control of the on-board part of the engine systems has been incorporated into the OX system. If you are interested in sensoring and data extraction, systems control/state machines or mechanical optimization, you should apply to the OX system.
The burn chamber subsystem of a rocket engine is the most easily recognizable part of it: it’s where the fire comes out. Much like in the engine of a car, the burn chamber is a cylindrical space where fuel and oxygen mix and combust, ignited by a pyrotechnical charge or similar. That combustion product is then directed downwards toward the nozzle, which accelerates it and eventually results in the thrust propelling the rocket upwards. The burn chamber can largely be broken down into the injector, pre combustion chamber, fuel grain/combustion chamber, post combustion chamber and nozzle. These are stacked directly on top of eachother inside the burn chamber casing.
The injector is essentially the carburetor of the engine, serving to “atomize” the liquid N2O into small droplets which decompose into N2 and O2. It contains no moving parts, but looks like a plate with several holes drilled through it - very much like the showerhead in your bathroom. The injector largely controls how much N2O is let into the burn chamber, and the exact geometry of the holes pose a complicated problem of fluid mechanics. The droplets are “injected” into the pre combustion chamber where they are heated and given time to decompose before they reach the fuel grain. In our engine, the fuel grain is a hollow cylinder of solid paraffin wax.
The now hot, gaseous O2 reacts with the wax to produce even more heat and CO2 gas (colloquially known as “fire”) which flows through the post combustion chamber to finish reacting before it finally reaches the nozzle or engine bell in the fiery spectacle we all know and love. The geometry of the nozzle greatly affects the efficiency (Isp) of the engine, and poses yet another fluid mechanics problem. All these components need to survive very high temperatures, and so must be made from materials that are resistant to heat and oxidizing.
If you are interested in fluid mechanics, pyrotechnic chemistry, simulations or materials science, the burn chamber is probably where you should apply.
All rockets need infrastructure on the ground. The filling station contains the storage tank for the N2O, valve systems that allow for safe filling of the OX tank, a mechanism to remotely detach the filling umbilical from the rocket, camera systems and various sensors for surveillance.
For ease of handling, the filling station comes in two parts: an aluminum framework to keep the N2O storage tank safely in place, and a stand where the valves and control electronics are installed.
The main challenge of the ground station is to design a way to safely and reliably detach the filling umbilical from the rocket fuselage. You might have seen rocket launches where the umbilicals are seemingly ripped from the fuselage as the rocket ascends off the pad. That would likely be catastrophic for a small rocket like ours. Modifying a valve connection to allow remote detachment has proven difficult, and is a combined problem of actuation and mechanics. Furthermore, the filling station must communicate wirelessly with the rocket and/or ground control station to transmit sensor data and valve control signals as well as the camera feed. We currently have no complete solution for this problem, and rely on maintaining line of sight when we are doing tests. The control system for the filling station must cooperate with that of the OX system, and when working with our prototype, the test bench, the two have for all intents and purposes been one system.
If you are interested in wireless communications, state machines, data extraction or valve systems, you should apply for the filling station.
Note: the injector is part of the burn chamber subsystem, but because of its complexity we've given it its own section.
The purpose of the injector is to deliver the liquid oxidizer as a highly atomized spray into the combustion chamber, so that the droplets can quickly evaporate into gas and partake in combustion. Additionally, the small injector orifices function as the mass flow rate limiter of the system. The mass flow rate of oxidizer is one of the most important factors of the rocket’s overall performance. This is particularly because of the resulting O/F ratio, or oxidizer to fuel mass ratio. The oxidizer we are using is N2O, which has a very flat curve in regards to efficiency or ISP. This means that even if the O/F ratio is off by a considerable margin, the efficiency will not be harmed substantially. This is not the case for other oxidizers such as liquid oxygen or H2O2. Regardless, knowledge of the O/F ratio inside the burn chamber is essential to understand and develop the rocket engine. This is a difficult problem, because the main factor in deciding the fuel regression rate (analog to fuel mass flow ratio) is the oxidizer mass flow ratio. As such, it is clear that knowing and controlling the oxidizer mass flow ratio to be as close to desired values as possible is absolutely critical for the rocket to function as intended.
There are many injector styles, the most relevant types being showerhead injectors, impinging injectors, swirl injectors and vortex. Showerheads are the simplest variety, with simple, straight holes. Impinging injectors have the streams out of the injector impinge onto each other, so that they break up and atomize more which helps combustion efficiency. Swirl injectors also atomize well, and introduce a swirling flow which has been shown to potentially increase regression rates significantly. Vortex injectors also introduce swirling flow, but in a more simplistic manner.
The chosen liquid oxidizer is nitrous oxide, N2O. Its main advantage is that it has a high vapor pressure at room temperature, which means that the tank can reach high pressures without necessarily requiring external pressurization systems or turbopumps to drive the flow. However, this also greatly complicates the analysis as two-phase flow is introduced, especially across the injector. Additionally, the pressure drop in the tank as it empties is a complicated process that is difficult to accurately model.
N2O is very similar to CO2, both linear molecules of similar size and thermodynamic properties. Therefore, CO2 can potentially be used as an analog for N2O while testing. This is useful because while N2O is considered a relatively safe oxidizer, there are still some important safety concerns with using it (nitrous oxide decomposition events).
A CFD approach to the problem was originally intended, but we were recommended against it by the previous students advisor (James Dawson) and CFD professor (Reidar Kristoffersen) when they saw the problem. This was largely due to computational power needed by the very fine meshes required for atomization. A different approach, where mass flow rate models have been coded in Python was used - this code will be available to you. However, you may find a CFD approach that is useful after further consultation with professors and research.
As for the nozzle, the composition of the gasses flowing through it is not known to precision, and neither is the temperature or pressure. This means that any work simulating the nozzle will include a considerable amount of simplification and guessing. The nozzle is of a simple DeLaval design, so the only changes in design would be the angles in front and behind the mouth of the nozzle.
Got questions? Drop us an email! Note that applications must be submitted through the applications page.