P18102: RIT Launch Initiative Hybrid Rocket
/public/Engine Specifications/Propulsion/


Table of Contents

Propulsion Summary

The propulsion system is responsible for the combination and reaction of the oxidizer and fuel, as well as the energy conversion that ultimately generates thrust to propel the rocket. The propellants are combined by injecting a liquid oxidizer through the injector into the combustion chamber where the solid fuel grain resides. This process commences after the chamber has been brought up to temperature and pressure by the igniter and pre-combustion chamber. After the oxidizer is released, the propellants react until the oxidizer or the fuel runs out. For our engine design, the burn will last approximately 7 seconds. The combustion products from the reaction reach pressures up to 500 PSI and temperatures above 5000 degrees Fahrenheit. Downstream of the fuel grain, any un-reacted combustion products mix in the post-combustion chamber, greatly increasing the efficiency of the system. Lastly, the thermal energy is converted into kinetic energy through the use of a converging-diverging nozzle. The gases go from negligible speeds at the entrance of the nozzle to supersonic speeds at the exit of the nozzle where they are blown into the atmosphere. The result is an engine that produces 1300 lbs of thrust for 7 seconds, propelling a 140lb rocket to 30,000 feet.

A summary of the key propulsion values is given in the table below:

Parameter Value Units
Thrust 1300 lbf
Thrust Coefficient 1.5 [-]
Specific Impulse 240 s
OF Ratio 5 [-]
Gamma 1.23 [-]
C-star 1566.4 [m/s]
Oxidizer Nitrous Oxide (N2O) [-]
Fuel Parrafin [-]
Chamber Pressure 500.00 psi
Chamber Temperature 3025.4 K
Used Oxidizer Mass 34.868 lbm
Used Fuel Mass 6.9735 lbm
Oxidizer Flowrate 4.9245 lbm/s
Fuel Flowrate 0.9849 lbm/s
Total Flowrate 5.9094 lbm/s
Chamber Inner Diameter 5.0000 in
Chamber Thickness 0.7500 in
Chamber Length 36.125 in

Combustion Chamber Assembly

Overall Assembly Layout

Combustion Chamber Assembly

Combustion Chamber Assembly


To ensure proper sealing of the chamber, we employed the use of O-rings in the nozzle and injector sections. The specific O-rings were selected using Parker's O-ring Handbook, linked below:

Parker O-ring Handbook

For each application, the O-rings were selected by a set-by-step approach in order to selecting the correct cross-sectional area, diameter/dash number, material and durometer/hardness. In addition, we used the Parker guide to properly design the O-ring glands and installation considerations (such as chamfers). Listed below are some of the O-rings used in our design and associated info related to their selection.

The chosen material for these O-rings are Fluorocarbon Rubber or FKM. This material was chosen for its high heat capabilities, its cost, and its compatibility with nitrous oxide. After reading through the Swiss Propulsion Lab Power Point found on the Safety Protocols page, we became aware of the fact that FKM rubbers swell significantly when exposed to nitrous oxide. This is not a huge concern for us as we do not plan on nitrous oxide coming on contact with the O-rings. We will conduct post-fire inspections of these O-rings to ensure that they are not damaged, and will replace these O-rings if needed.

General Propulsion Requirements

  1. Produce 1300 lbf of thrust
  2. Operate with a maximum total impulse of 9,209 lb-sec
  3. Safely convert propellant chemical energy to usable kinetic energy
  4. Avoid the formation of shock waves inside the nozzle
  5. Survive all thrust, vibrational, and thermal loads

Oxidizer Selection

Our oxidizer was carefully selected by analyzing the advantages, disadvantages, and impact on the entire engine design process. The 2 final candidates were nitrous oxide and gaseous oxygen. Nitrous oxide is a highly used oxidizer for hybrid rocket engines because of its self-pressurization properties and high vapor pressure. This allows for a simpler feed system design. On the other hand, oxygen is the best oxidizer when considering the combustion process. Regardless of the oxidizer chosen, it is the oxygen molecules themselves that react with the fuel. Gaseous oxygen was the only form of pure oxygen considered because of the complexities associated with designing and storing liquid oxygen (LOX) systems, a cryogenic liquid.

An extensive Pugh Chart of the two oxidizer options is shown below.

N2O (liquid) O2 (gas)
Vapor pressure + -
Performance (Isp) - +
Performance density + -
Handling + -
Cost 0 0
Regression complexity + -
Thrust chamber size + -
Storage 0 0
Total engine weight + -
Flow rate feasibility + 0
Oxygen content - +
Adiabatic temperature - +
Storage pressure 0 -
Thermal risks 0 -
Instabilities - +
Oxidation/Erosion rate + -
Sum of results +5 -3

Ultimately, the above analysis shows that nitrous oxide will be the most advantageous oxidizer for our rocket engine.

Fuel Selection

Our fuel selection was based on careful analysis of the performance, handling, and manufacturing of the different options. For the 2019 IREC competition, we are limited to using non-toxic propellants. Non-toxic propellants require no special handling, storage, or protective equipment. Therefore, fuel options were limited to hydro carbons, which happen to be easily accessible and cost effective.The fuel options were eventually limited to 2 choices, both of which are commonly used in hybrid engines. The first is HTPB (an elastomer) and Paraffin (a wax). Both of these fuels are stable in the solid state at room temperature and therefore are very safe to handle. They will not spontaneously combust unless they are somehow brought up to ignition temperature and exposed to a flow of oxidizer.

Another Pugh Chart was developed to compare both of these options.

Paraffin (Solid) HTPB (Solid)
Performance (Isp) 0 0
Performance Density 0 0
Regression Rate + -
Turbulent Heat Transfer + -
Heat of Vaporization + -
Heat of Combustion + 0
Handling 0 0
Storage 0 0
Total Weight 0 0
Mass flow rate uniformity + -
Cost 0 0
Sum of results +5 -4

The fuel of choice for our application is Parrafin wax. Among all of the reasons in the above chart, the largest reasons are because of it's superior regression rate and regression uniformity. This will ensure a uniform burn, constant flow rate, and predictable thrust over our burn.

O/F Ratio

The OF ratio represents the ratio of oxidizer to fuel by mass. Our research shows that a ratio of 6.5 would be optimum in order to achieve the stoichiometric ratio for the propellants. If the engine operates at the stoichiometric ratio, the reaction will be it's most efficient and burn the hottest. This is desireable because it produces the highest thrust for the given propellants. However, the OF ratio for our system has been chosen to be 5 for several strategic reasons. This allows us to run our combustion "fuel rich" in order to reduce combustion chamber temperatures and lower our thermal management system requirements. This allows for the use of more readily available and cost effective materials. Another reason for running fuel rich is that it reduces the amount of oxidizer that will need to be carried in the rocket and therefore reduces the size and weight of the oxidizer tank. Lastly, with less oxygen in the combustion products there will be less oxidation and erosion of critical engine components such as the nozzle. This is extremely important because it will not only extend the life of the nozzle, but also maintain the critical dimensions of the nozzle that govern the performance of the engine. This ensures repeatability and reliability.

Thrust Validation

Thrust Validation

Thrust Validation