1000 Pound Thrust Regeneratively Cooled LOX-Kerosene Rocket

Update, 20-June-2012: since this project has been put on hold for the past 5 year, I've decided it is time to sell the components.  Soon I'll be selling the lr101 engine, stainless steel lox and kerosene tanks, composite helium pressure tank, a couple pressure regulators, a composite pressure tank from a Pershing missile, and a few other things.

It's been over 10 years since I've updated this page.  It's about time I continue to document the design.

[Under Construction]Under construction. Revised: 06/23/12.
This page will be updated from time to time as I transcribe the design notes from my trusty yellow pad

Design Considerations
System Description
Engine Parameters
Injector Design
System Components
Normal Operation
Test results

[BACK]      [HOME]


After seeing a few liquid fueled engine tests and flights at the Mojave Test Area of the Reaction Research Area, I figured I would take up the challlenge to design, build, and fly a 1000 pound thrust liquid fueled rocket.

Design Considerations
The engine I'm using is a Rocketdyne LR101 vernier engine I purchased from Norton Sales in North Hollywood, CA.  The LR101 is a regenerative cooled engine using liquid oxygen (LOX) and kerosene as propellants and producing 1000 pounds of thrust.  The optimum oxidizer to fuel ratio for LOX/kerosene is 2.56.  I plan to use a fuel rich mixture of 2.20 to help with cooling.

System Description

Engine Parameters
Measurements of the LR101 are as follows:

Chamber diameter = 2.725 in area = 5.832 in2
Throat diameter = 1.625 in area = 2.074 in2
Nozzle Exit diameter = 3.675 in area = 10.607 in2

From these measurements the contraction and expansion ratios can be calculated

Contraction ratio = chamber area/throat area = 5.832/2.074 = 2.8
Expansion ratio = nozzle exit area/throat area = 10.607/2.074 = 5.1

Target Isp

Operating charateristics and principal dimensions

Propellants LOX / Jet-A
O/F mixture ratio 2.30
Characteristic velocity c*, ft/s 5400
Thrust coefficient Cf (sea level) 1.489
Specific impulse (Is)tc (sea level), s 249
Total propellant flow rate, lb/s  
Thrust (sea level), lb 1000
Chamber pressure (injector end), psia 0
Chamber pressure (nozzle stagnation), psia 0
Average gas specific-heat ratio 1.233
Combustion-chamber cross-section area, in.2 5.83
Throat area, in.2 2.07
Nozzle exit area,in.2 10.61
Combustion-chamber volume, in.3  
Combustion-chamber length (to throat), in.  
Characteristic chamber length L*, in.  
Overall thrust-chamber length, in.  
Design contraction area ratio 2.8
Design expansion area ratio 5.1

Back to Top

Injector Design

Oxidizer density:  1.

Back to Top

System Components & Layout

Back to Top

Normal Operation

Back to Top

Test Results

Back to Top

Rocket Propulsion Elements, 7th Edition
By George P. Sutton, Oscar Biblarz

Modern Engineering for Design of Liquid-Propellant Rocket Engines (Progress in Astronautics and Aeronautics, Vol 147)
By Dieter K. Huzel, David H. Huang

How to Design, Build and Test Small Liquid-Fuel Rocket Engines
Rocketlab / Chine Lake, Calif.
by Leroy J. Krzycki

Amateur Liquid Propellant Rocket Experimentation 1943 to 1998, Volume II
RRS Liquid Rocket History, 1990 to 1993
RRS News, Volume 54, Number 4, December 1997
By Reaction Research Society

Pocket Ref, 2nd Edition
Compiled by Thomas J. Glover

CRC Handbook of Mathematical Sciences, 6th Edition
William H. Beyer, Ph.D. (editor)

Flometrics, Inc
Solana Beach, CA

Norton Sales
North Hollywood, CA

Burden's Surplus Center
Lincoln, NE


Back to Top


Revised: 06/23/12.

[BACK]     [HOME]