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Hybrid Design:
Hybrid engines are also a relatively new technology. Hybrid rocket engines
combine both a solid fuel with a liquid oxidizer. The combination provides a
performance greater than solid engines, but a cost and simplicity that can’t be achieved
by liquid engines. Another benefit of the hybrid engines over solid engines is that the
hybrid engine can be shut down if a problem occurs by simply shutting off the flow of
oxidizer to the solid fuel. This will allow a shutdown of the upper stage if a problem
occurs near the manned first stage.
The hybrid engine for the second stage was designed using conceptual design
methods 7 . Many different fuels were analyzed with HTPB/LOX combination providing
the best Isp. Unfortunately for the size of the second stage this results in a vehicle that is
almost 10 feet longer than the 42 foot payload bay designated in the baseline. Therefore
hydrogen peroxide (H2O2) was used the next best performance and a higher mixture ratio
than the HTPB/LOX configuration which results in a smaller vehicle (Since the oxidizer
is more dense than the fuel).
Once the propellant was determined the fuel chambers were designed to be as
short as possible while still having the proper length to diameter ratio to support
combustion. For hybrids this seemed to result in a seven port fuel chamber as shown in
Figure 15.
Figure 15: Seven Port Hybrid Design.
The Isp for this design can be estimated using conceptual design equations [7] and nozzle
parameters that were set to be the maximum nozzles to fit within the diameter as well as
the length constraints of the upper stage (approximated as 80% of a 15 degree half cone).
This resulted in an Isp of approximately 310 seconds. With this design and the propellant
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