*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.

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

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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 (

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 (

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. Engine sizing is constrained by three Three main variables constrain engine sizing,

   Eazy math inline
   body P_{c}

   , thrust

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   body F_{t}

   , and

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   body \dot{m}

   . We used a Simulink simulation to decide on an

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   body F_{t}

   of

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   body 2500N

   and metallurgy and COTS valve/fitting pressure ratings to arrive at a

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   body P_{c}

   of

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   body 2500 \mathit{kPa}

   or

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   body 362.6\mathit{Psi}

   .

3. Using NASA CEA. We input

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   body P_{c}

   ,

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   body P_{e}

   , and

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   body \mathit{MR}

   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

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   body {M}

   ,

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   body {\gamma}

   (equivalent to

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   body \frac {C_{p}} {C_{v}}

   ),

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   body \frac {A_{e}} {A_{t}}

   ,

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   body {V_{e}}

   , and

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   body \textrm{ISP}

   .

4. The following parameters are fed into a Matlab script:

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   body \dot{m}

   ,

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   body \mathit{MR}

   ,

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   body P_{c}

   ,

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   body P_{e}

   , and

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   body L^\star

   (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