Engineering topics
Explore propulsion reports from NASA NTRS and DOE OSTI. Search rocket engines, electric propulsion and thrusters; verify designs and performance in the source documents.
Evaluation of extraterrestrial resources to supply propellants for refueling rockets at various stations during interplanetary flight
The Resonant Combustion Program uses a bombing technique to precipitate the transition from the steady state to the resonant, or oscillatory, rocket combustion mode. The initial pressure disturbance produced by the high-explosive device, when exploded in a nonreactive environment of cold N2 gas, is a blast-wave having an initial amplitude of several thousand psi (within a radius of approximately one inch from the bomb case) which is still several hundred psi strong even at a radius of 4 to 5 in. The use of such a severe triggering disturbance has brought up the question of whether or not the violent sustained oscillatory mode observed in this program is forced by the character of the initial disturbance itself.
Size, type, and cost of facility simulating 200,000 ft altitude for testing small chemical rockets and components
Mars and Venus atmospheres as working fluids in gas turbines and liquid oxidizers for rocket engines
Experimental data on performance and cooling rates of rocket thrust chambers using FLOX/light- hydrocarbon propellants
Experimental equipment for impingement test of electrostatic rocket exhaust jet
Test equipment, immersion tests, and spectral distribution of solar simulator for studying effects of electrostatic rocket exhaust impingement on spacecraft
Solid propellant rocket motor predictive design by correlating performance and weight statistics via computer program
Chemical effects of rocket exhaust impingement on spacecraft surfaces
Computer program characterizing behavior of pressure-fed rocket engine system
Tolerance level for deleterius effects of electrostatic rocket exhaust impinging on spacecraft surfaces
Method for producing porous tungsten plates for ionizing cesium compounds for propulsion of ion engines
Developing space-storable liquid propulsion components for future space missions
Digital computer program for studying transient performance of pressure fed rocket engines
Optimum methods for orbital resupply, maintenance, and repair of hybrid propellant rocket engine subsystem for orbiting space station attitude control
During recent cryodeposit tests with an 0.18-N thruster, the mass flux in the plume back field was measured for the first time for nitrogen, carbon dioxide, and a mixture of nitrogen, hydrogen, and ammonia at various inlet pressures. This mixture simulated gases that would be generated by a hydrazine plenum attitude propulsion system. The measurements furnish a base upon which to build a mathematical model of plume back flow that will be used in predicting the mass distribution in the boundary region of other plumes. The results are analyzed and compared with existing analytical predictions.
Choice of type and placement of thrusters on spacecraft (s/c) should include consideration of their effects on other subsystems. Models are presented of the exhaust plumes of mercury, cesium, colloid, hydrazine, ammonia, and Teflon rockets. Effects arising from plume impingement on s/c surfaces, radio frequency interference, optical interference, and earth environmental contamination are discussed. Some constraints arise in the placement of mercury, cesium, and Teflon thrusters. Few problems exist with other thruster types, nor is earth contamination a problem.
A new reference design was prepared for the 5 kW(e) thermionic power supply. The safety equipment in this design is a passive containment system which does not rely on the operation of any mechanisms such as a launch escape rocket or deployment of parachutes. It includes: (1) a blast shield to protect against the explosion of the launch vehicle; (2) a combination of refractory thermal insulation and heat storage material to protect against a sustained launch pad fire; (3) a reentry body with a spherical nose and a large conical flare at the aft end to stabilize the reentry attitude and lower the terminal velocity in air; (4) composite graphite thermal protection to sustain the reentry heat pulse; (5) crushable honeycomb behind the nose to limit the deceleration of the radioisotope source due to impact on land at terminal velocity; (6) a double-walled secondary containment vessel surrounding the isotopic capsules; (7) neutron shielding to reduce external dose rates; (8) an auxiliary cooling system employing redundant heat pipes to remove the radioactive decay heat from the heat source and reject it to the surroundings or to a forced convection loop.