HRE100K OUR FIRST HYBRID ROCKET ENGINE

Hybrid Rocket Engine 100K (HRE100K for short) is the codename of the project  Skyward Experimental Rocketry is currently working on in order to develop the best fitting engine for Rocksanne II-X. The engine owes its nickname to the fact that it has been designed to reach a Total Impulse 100,000Ns, which will boost Rocksanne over 10km in height: it will have to propel a 7m tall rocket, with an approximate external diameter of 22cm.

The engine will give a critical contribution to one of the main goals of Skyward: designing, building and launching a rocket that will reach the old record-breaking apogee of 12.5km. To top it off, as things stand today, HRE100K is to be officially the first Hybrid Rocket Engine to be flown by an Italian-based student rocketry association.

What’s that?

A hybrid rocket engine is a type of chemical rocket engine in which the reagents are stored in different phases (in our case, a liquid oxidizer and solid fuel). Its constitutive principle goes along the lines of the Latin motto “In medio stat virtus”, as it combines both the benefits of solid rocket motors and the ones of liquid rocket engines.

The typical components of a hybrid rocket motor include:

• An oxidizer tank
• Valves, piping, and other feed system elements
• Turbomachinery, when not using a pressure feed system
• An igniter, or hypergolic ignition
• An injector (which sprays the oxidizer in the combustion chamber)
• A combustion chamber, where the sprayed oxidizer is ignited and reacts with the solid fuel grain
• A Thrust Vector Control (T.V.C.) to control and correct flightpath (this can be supported by fins and other movable surfaces)
• A supersonic nozzle

Compared to a solid rocket motor, a hybrid rocket engine yields a higher Isp and is more reliable: since the chemical reagents are physically separated, the system is less prone to unexpected spontaneous combustions and explosions; also, it is controllable by adjusting the position of a valve placed at the end of the oxidizer tank (typically by the means of a remote controlled servo), so that the combustion can be stopped (e.g. in the case of an aborted mission) and even restarted.

The core principle we have exploited for designing the engine is strictly related to the famous thrust equation:

$$T= \dot{m}u_e + A_e(p_e – p_a)$$

In fact, a good initial choice of a parameter to work on is mass flow rate, which is tightly linked to the regression rate, i. e., the speed at which solid fuel is consumed  in the combustion chamber, governed by the equation

$$\dot{r}=aG^{n}x^{m}$$

Where $$a$$, $$n$$ and $$m$$ are constants which depend on the fuel, $$G$$ is the total propellant mass flux and $$x$$ is the distance down the port. The relation between mass flow rate and regression rate is given by

$$\dot{m}=\dot{m}_{ox} + \int_0^x \rho_{f}\pi d_{p}\dot{r}(\tilde{x})d\tilde{x}$$

Where $$d_p$$ is the port diameter.

Simplifying things a bit, to maximize the thrust, it is needed to burn a propellant as dense as possible in the shortest amount of time: in this way, because of Newton’s third law, a fairly big mass is given a good acceleration by the engine in a certain direction (i.e. a force is beign applied) and, as a reaction, the engine is receiving a force with the same intensity, but opposite in direction.

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A 4m, lubricant free aluminum tank stores around 120l of N2O in a vessel resistant up to 150bar pressures.

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A servo-adjustable Stainless Steel 316 valve feeds the oxidizer to our custom-made aluminum injector.

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3 Igniters are symmetrically mounted on the bulkhead.

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The combustion of N2O and a mixture of paraffin and SEBS, burning for 20 seconds, generates enough thrust to reach a total impulse of 100,000Ns.

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The engine will measure around 210mm in external diameter and it will be 80cm long.

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A carbon-phenolic nozzle conveys 5000N of mean thrust.