We can see from table A.4.2.2.4.1 that for our four types of propellants (cryo, storable, hybrid, and solid) specific impulse does not vary greater than ten seconds, with the exception of liquid oxygen / liquid hydrogen. Due to great variation between liquid hydrogen / liquid oxygen and the other cryogenic fuels, we specifically chose liquid oxygen / liquid hydrogen as our cryogenic fuel for Model Analysis simulations. Selection for the other three systems required more analysis. Generally speaking, the cost of the propellant will be negligible in comparison to the cost of the engine and sub systems of the launch vehicle. In modern rocket design, most propellants are manufactured in bulk quantities which tend to limit costs. However, certain propellants, regardless of propellant price, can be extremely expensive due to the handling costs associated with their use. The main qualities we consider in the selection of the four propellants are the costs of equipment, personnel, and design alterations associated with the propellant characteristics. Some of the main considerations are: Highly Pressurized Propellants Low Temperature Propellants (cryogenics) Toxicity Volatile Nature Highly pressurized propellants require stronger, and thus, thicker tanks. These thicker tanks are heavier and more costly which makes them detrimental to our design. Propellants under high pressure also pose a safety hazard as they are less stable. Low temperature propellants pose a problem due to the tank, piping, and engine requiring insulation against the frigid temperatures. This also demands additional weight and cost associated with the insulation. Additional person hours are needed to fill the launch vehicle due to precooling the lines. A potential risk associated with cryogenic systems is boil-off leading to over pressurization of the fuel tanks and pressurant lines. This requires the addition of a bleed system to prevent explosion which increase complexity and cost.
A.4.2.8 Testing When it comes to testing our rocket engines we decided to try and design a launch vehicle which could be tested here at Purdue’s Zucrow High Pressure Laboratories. After some research into the test labs at Purdue we as a team got a much better understanding of what we could accomplish here. For testing rocket engines there are certain things that a facility must have in order to have a successful test. For example the test stand at Purdue’s high pressure laboratories contains not only a 10,000 pound thrust stand, but also the capability to test liquid propellants such as hydrogen peroxide, liquid oxygen and liquid hydrocarbon. The facility currently has propellant tanks and tubing for these fuel and oxidizer types. This was also a factor in determining the propellant type for some of our designs. For example our hybrid rocket engines incorporate hydrogen peroxide (H2O2) as the oxidizer. Because of this Purdue would be a suitable candidate for testing. As for solid rocket engines there is no propellant that needs to be fed into the system because the test engine already has the solid propellant (fuel and oxidizer) built in. Because of this fact the only parameter that needs to be looked at for Purdue testing would be the test stands maximum thrust capability. After researching Zucrow laboratories on Purdue’s website and through various web articles and a visit to Zucrow we determined exactly how big of a rocket the facility could handle.1 Some of the restrictions at Zucrow Laboratories are given below in Table A.4.2.8.1. Table A.4.2.8.1: Purdue’s Zucrow Laboratories facility restrictions. Maximum Capability Value Units Thrust 44,480 N Chamber Pressure 4.137 MPa Mass Flow Rate 6.803 kg/s In the beginning of our design we to stay within these parameters. The first iteration of designs tried to keep these parameters as a limit to the performance of our rocket engine. On top