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Results for “COMBUSTION EFFICIENCY”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Effect of Ozone Addition on Combustion Efficiency of Hydrogen: Liquid-Oxygen Propellant in Small Rockets

An experimental study shows that 2 percent by weight ozone in oxygen has little effect on overall reactivity for a range of oxidant-fuel weight ratios from 1 to 6. This conclusion is based on characteristic-velocity measurements in 200-pound-thrust chambers at a pressure of 300 pounds per square inch absolute with low-efficiency injectors. The presence of 9 percent ozone in oxygen also did not affect performance in an efficient chamber. Explosions were encountered when equipment or procedure permitted ozone to concentrate locally. These experiments indicate that even small amounts of ozone in oxygen can cause operational problems.

Miller, Riley O.↗
NASA NTRS1979

Design, fabrication and delivery of a high pressure LOX-methane injector

An injector for 3000 psia chamber pressure using liquid oxygen and gaseous methane propellants is presented. The injector is intended to be evaluated during a series of pressure-fed test firings using a water-cooled calorimeter chamber and a milled-slot regenerative chamber. Combustion efficiency, combustion stability, ignition and injector face heat transfer assessments were made for candidate injector body and pattern design approaches. This evaluation resulted in baselining an oxidizer post type manifold with a 60 element platelet coaxial swirler injector pattern. An axial acoustic resonator cavity was created at the injector/chamber interface.

Valler, H. W.↗
NASA NTRS1950

Altitude Performance and Operational Characteristics of 29-inch-diameter Tail-pipe Burner with Several Fuel Systems and Fuel-cooled Stage-type Flame Holders on J35-A-5 Turbojet Engine / Richard L. Golladay and Harry E. Bloomer

An investigation of tail-pipe burning was conducted in the NACA Lewis altitude wind tunnel with a full-scale turbojet engine and an 29-inch-diameter tail-pipe burner. Effects of fuel distribution and number and arrangement of stages on performance and operational characteristics of several fuel-cooled flame holders are presented and discussed. Operation with a three-stage flame holder having the large stage upstream was the most efficient. Combustion efficiency was slightly increased at high altitudes by injecting fuel upstream of the flame holder.

TURBINES, GAS - AFTERBURNING↗
NASA NTRS2005

"Powdered Magnesium: Carbon Dioxide Combustion for Mars Propulsion"

Powdered magnesium - carbon dioxide combustion is examined as a potential in-situ propellant combination for Mars propulsion. Although this particular combination has relatively low performance in comparison to traditional bi-propellants, it remains attractive as a potential basis for future Martian mobility systems since it could be partially or wholly manufactured from indigenous planetary resources. As a means of achieving high mobility during long-duration Mars exploration missions, the poorer performing in-situ combination can, in fact, become a superior alternative to conventional storable propellants, which would need to be entirely transported from earth. Thus, the engineering aspects of powdered metal combustion devices are discussed including transport/injection of compacted powder, ignition, combustion efficiency, combustion stability, dilution effects, lean burn limits, and slag formation issues. It is suggested that these technological issues could be effectively addressed through a multi-phase research and development effort beginning with basic feasibility tests using an existing dump configured atmospheric pressure burner. Follow-on phases would involve the development and testing of a pressurized research combustor and technology demonstration tests of a prototypical rocket configuration.

Foote, John P.↗
NASA NTRS2007

Powdered Magnesium-Carbon Dioxide Rocket Combustion Technology for In Situ Mars Propulsion

Powdered magnesium (Mg) carbon dioxide (CO2) combustion is examined as a potential in situ propellant combination for Mars propulsion. Although this particular combination has relatively low performance in comparison to traditional bipropellants, it remains attractive as a potential basis for future martian mobility systems, since it could be partially or wholly manufactured from indigenous planetary resources. As a means of achieving high mobility during long-duration Mars exploration missions, the poorer performing in situ combination can, in fact, become a superior alternative to conventional storable propellants, which would need to be entirely transported from Earth. Thus, the engineering aspects of powdered metal combustion devices are discussed including transport/injection of compacted powder, ignition, combustion efficiency, combustion stability, dilution effects, lean burn limits, and slag formation issues. It is suggested that these technological issues could be effectively addressed through a multiphase research and development effort beginning with basic feasibility tests using an existing dump configured atmospheric pressure burner. Follow-on phases would involve the development and testing of a pressurized research combustor and technology demonstration tests of a prototypical rocket configuration.

Foote, J. P.↗
NASA NTRS1958

A Preliminary Investigation of the Performance of a Short-length Turbojet Combustor Using Vaporized Hydrocarbon Fuels

Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Jones, R. E.↗