Experimental Evaluation of Six Ablative-Material Thrust Chambers As Components of Storable-Propellant Rocket Engines.
Abstract
Six ablative-material thrust chambers were tested as components of storable-propellant (nitrogen tetroxide N2O4 and a 50-50 blend of unsymmetrical dimethyl hydrazine with hydrazine N2H4) rocket engines. The nominal initial nozzle-throat diameters were 7.82 inches (19.8 cm). Engine operating conditions were held constant at a chamber pressure of 100 psia (689 kN/sq meters) and an oxidant-to-fuel ratio of 2.0. The six ablative-material thrust chambers provided both material and geometry variables. p4
Document Details
- Document Type
- Technical Report
- Publication Date
- Jun 01, 1967
- Accession Number
- ADA306740
Entities
People
- Arthur M. Shinn Jr.
Organizations
- Glenn Research Center