*Note: This page is not meant as a primer on the design or sizing of liquid engines. There are many fantastic resources created by other groups, here are some:

Liquid Rocket Engine Sizing - USC Viterbi

Charlie Garcia's Playlist

Sizing a Rocket Engine - Liquid Propulsion Lab

USC Liquid Propulsion Laboratory

Overview

Liquid engine sizing is the process that relates the chemical processes of combustion to physical parameters. It is separate, but related, to engine design which will consider materials, heat transfer, and many other physical constraints.

Engine sizing can seem somewhat complex, much of this is due to the berth of prerequisite knowledge that is needed to fully grasp the formulas, concepts, and processes. The reality is that the engine sizing approach used by many collegiate teams benefits from, but does not require, advanced knowledge of fluid mechanics, gas dynamics, or propulsion. Again, engine sizing may seem complex and while it can be the basics can be grasped without an extensive technical background.

In reality, the process is straightforward, not simple, but straightforward. Figure 2 shows the process RPL used to size its CPLC engine.

Selecting Initial Parameters

Engine sizing involves a large number of variables both quantitative and qualitative. However, nearly all of these can be constrained by the selection of a propellant combination and two of the following three variables: chamber pressure (), thrust (), and mass flow rate (). Selecting two of these and a propellant combination will allow you to proceed to NASA CEA and define engine performance and size. Oftentimes, these parameters are driven by systems-level requirements such as certain vehicle performance/flight profile requirements, propellant selection dictated by a challenge/competition/school rules, or any number of program requirements.

See the following pages for more details on: Selecting Propellant Combination , Selecting a Design Thrust, and Selecting a Preliminary Chamber Pressure.

Again, we chose to size with respect to chamber pressure () and thrust (). It would be theoretically possible to select a , , and and to then drive propellant requirements but this technique is not feasible or productive. Propellant combination selection is driven by overarching program or organization requirements (ex. our school does not allow LOX effectively requiring us to use N2O).

NASA Combustion Equilibrium Analysis

1. The propellant combination is one of the most significant factors in engine sizing. Nitrous Oxide, N2O is our oxidizer, and Isopropyl Alcohol, IPA is our fuel. IPA was selected as fuel due to its low cost and accessibility.

2. Three main variables constrain engine sizing, , thrust, and . We used a Simulink simulation to decide on an of and metallurgy and COTS valve/fitting pressure ratings to arrive at a of or .

3. Using NASA CEA. We input,, and for our N2O oxidizer and IPA fuel. CEA allows you to simulate various mass fractions. We chose 4.0 as it gives the best ISP performance, this burns hot but not as hot as a more OX-rich mix. CEA outputs the , (equivalent to ),, , and .

4. The following parameters are fed into a Matlab script: , , , , and (L* is somewhat arbitrary and selected) as well as chemical properties. m_dot is adjusted! We chose an L* of 1 m, and m_dot 1.1 kg/s. That yields V_c. We pick 0.15 m L which gives a D_c of 7.73cm and exit diameter of 6.3 cm and a throat diameter of 2.99 cm.

Variable Table