![[Under Construction]](undercon.gif){kind=small}
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.
| History Design Considerations System Description Engine Parameters Injector Design System Components Normal Operation Test results Resources |
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History
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.
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 |
Oxidizer density: 1.
Resources
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