Weekly Meetings:
Wednesdays at 12:00 PM in WH 113
Workdays 1:00 PM to 5:00 PM in WH 113
Division Milestones:
2016: Division Founded
2017: Placed 1st in Undergrad Research at Southwest Aerospace Symposium

Rocketry Research Division


The Rocketry Research Division is the organization's youngest and most ambitious division. Created in 2016, the objective of this division is to explore and research various rocketry-related academic-focused topics. The current long term goal of this division is to Design, Build, Test, and Fly a liquid-fueled rocket engine. We anticipate this engine to be the first liquid-fueled rocket engine ever flown by UTA students.

Division News and Updates

2017 October 6

With the start of a new semester, the team has almost doubled the amount of active members and productivity has improved. We look forward to a productive semester. We are happy to announce that the team participated in the College of Engineering REU, and was awarded $2,000 to help research costs.

Due to the high atmospheric temperatures of the possible testing sites, the current pressure of the nitrous oxide could allow for evaporation in the feed lines. In order to counteract the problem the team has increased the pressure of the combustion chamber, which has caused the other systems to change. The combustion chamber has been changed to a smaller diameter in order to withstand the extra pressure load, this makes it so cooling system's efficiency had to be improved to counteract the increased heat flux. Finally the Inconel has been changed to an Inconel 625 because it has a higher melting point.

The team has finished the schematic for the feed system and now is in talks with a local company to streamline the design and contract valves. The injector has been updated and will soon be ready for simulations. Once simulations and redesigning is complete the manufacturing can be contracted. A code is being developed for the cooling system in order to calculate the size needed for the channels.


2017 January 14

The team has been working hard this past semester and significant progress has been made. However, the team did not complete all of the goals we had set for this semester and are therefore slightly behind schedule. Some of the progress that was completed includes determining several major design parameters for the injector, the fuel and oxidizer tanks, and the feed system.

It was decided that the feed system would be a pressure-regulated system, rather than a blow down system, utilizing helium gas as the pressurant. The type of discharge system for the tanks has been narrowed down to either a direct gas interface, or a piston displacement mechanism. Both systems will be studied further in order to determine which will be used on the final vehicle. For the time being, a direct gas displacement system will be designed for the fuel tank, while a piston displacement system will be used for the oxidizer tank.
Image Reference: Sutton, G.P. and Biblarz, O.B. “Rocket propulsion elements,” John Wiley & Sons, 7th Ed., New York, 2001, p.6

Pressure Regulated Feed System

Pintle Style Injector

For the pintle injector, both a fuel centered and oxidizer centered flow were examined, and it was decided that the central fluid would be the fuel, while the annulus fluid will be the oxidizer. The basic geometry (of the 1st iteration) was determined, and chosen to be similar to the figure shown. CAD models have been started and several variations of the manifold design are in progress. It is expected that after the 1st iteration of the injector is manufactured, it will be tested in order to experimentally determine several major parameters such as the discharge coefficients and pressure drops. After testing has been completed, the injector will be redesigned in order to meet all of the design requirements.
Image Reference: Yang, V., Habiballah, M., Hulka, J., Popp, M., “Liquid Rocket Thrust Chambers: Aspects of Modeling Analysis and Design”, Progress in Astronautics and Aeronautics, Volume 200, AIAA, Reston, VA., 2004. pp. 160

This past semester, RPA (Rocket Propulsion Analysis) software was used in order to get first order approximations of the heat exchanger design characteristics. After several iterations of simulations however, it was decided that a different approach would be taken for the spring semester. Next semester the team intends to perform analytical calculations by hand (in MATLAB) in order to approximate the characteristics of the system and to determine the major geometrical parameters. After an initial geometry is determined, simulations will be performed in RPA to verify the design, and CAD models will be created. After the CAD models are complete, further analysis will be performed using ANSYS.


Shown in the figure below is a bare-bones skeleton of the thrust chamber the team is working on (please note: this model is incomplete). As seen in the figure, we are using a conical nozzle for the first iteration. This is to simplify the process so that we can focus on other features of the system. In future design iterations, the team intends to perform the axisymmetric method of characteristics in order to design a contoured bell nozzle.

Next semester the team is planning to continue design on the feed system, injector, and heat exchanger, while also starting to begin analysis on the control system and structural design.


2016 September 09

The research team has made significant progress over the summer; the preliminary design for the thrust chamber has been completed using quasi-1D analysis. The team utilized MATLAB, RPA, and CEA software packages for the preliminary design. Several other major design parameters have been determined for the pressure fed rocket engine, and the approximate values are summarized in the table below. Work has also been started on the design of the regenerative cooling system and the pintle style injector. The goals for the fall semester are to complete the injector and heat exchanger design, as well as complete the plumbing and feed system design, and also start the preliminary design of the control system. The team also plans to create CAD models of all the components and run high-fidelity simulations using ANSYS.



Fuel CH3OH(L) (Methanol)
Oxidizer N2O(L) (Nitrous Oxide)
O/F ratio 3.5
Total Propellant Flow Rate 2.6 [kg/s]
Expansion Ratio 4
Chamber Pressure 20 [bar]
Optimum Thrust 5 [kN]
Is(Optimum Specific Impulse) 211 [s]
c*(Optimum Characteristic Velocity) 1494 [m/s]
c(Optimum Effective Exit Velocity) 2070 [m/s]
CF(Optimum Thrust Coefficient) 1.4