Engineering topics

Ewen, R. L.

Publications and source records attributed to Ewen, R. L..

9 records

NASA NTRS1989

Heat Flux In A Dual-Throat Rocket Engine

Report describes tests of dual-throat rocket engine burning hydrogen in oxygen. In tests, heat-flux profiles measured in inner nozzle and outer chamber. Dual-throat engine being considered for advanced space transportation. Consisting of two combustion chambers in series operating separately or together, engine makes it possible to change mode of operation in flight to obtain performance in most advantageous regime. Produces high thrust at sea level or low thrust with higher performance at high altitudes or in space.

Ewen, R. L.↗

NASA NTRS1986

Dual-throat thruster thermal model

The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

Ewen, R. L.↗

NASA NTRS1986

Dual throat thruster test results

Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of a gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A bleed flow of less than two percent of the primary chamber flow was required.

Ewen, R. L.↗

NASA NTRS1981

Advanced oxygen-hydrocarbon rocket engine study

This study identifies and evaluates promising LO2/HC rocket engine cycles, produces a consistent and reliable data base for vehicle optimization and design studies, demonstrates the significance of propulsion system improvements, and selects the critical technology areas necessary to realize an improved surface to orbit transportation system. Parametric LO2/HC engine data were generated over a range of thrust levels from 890 to 6672 kN (200K to 1.5M 1bF) and chamber pressures from 6890 to 34500 kN (1000 to 5000 psia). Engine coolants included RP-1, refined RP-1, LCH4, LC3H8, LO2, and LH2. LO2/RP-1 G.G. cycles were found to be not acceptable for advanced engines. The highest performing LO2/RP-1 staged combustion engine cycle utilizes LO2 as the coolant and incorporates an oxidizer rich preburner. The highest performing cycle for LO2/LCH4 and LO2/LC3H8 utilizes fuel cooling and incorporates both fuel and oxidizer rich preburners. LO2/HC engine cycles permitting the use of a third fluid LH2 coolant and an LH2 rich gas generator provide higher performance at significantly lower pump discharge pressures. The LO2/HC dual throat engine, because of its high altitude performance, delivers the highest payload for the vehicle configuration that was investigated.

Obrien, C. J.↗

NASA NTRS1980

Advanced oxygen-hydrocarbon rocket engine study

Preliminary identification and evaluation of promising LO2/Hydrocarbon rocket engine cycles were used to produce a consistent and reliable data base for vehicle optimization and design studies. cycles G and C were chosen for design analysis. Preliminary design analysis of the heat transfer subsystem was performed to establish major technology requirements.

Obrien, C. J.↗

NASA NTRS1980

Advanced oxygen-hydrocarbon rocket engine study

The program consists of parametric analysis and design to provide a consistent engine system data base for defining advantages and disadvantages, system performance and operating limits, engine parametric data, and technology requirements for candidate high pressure LO2/Hydrocarbon engine systems. The parametric chamber and nozzle cooling analysis was completed for the four potential coolants: RP-1, LCH4, LO2, and LH2. A summary of the cooling capability of each propellant is presented.

Obrien, J. O.↗

NASA NTRS1977

Combustion effects on film cooling

The effects of: (1) a reactive environment on film cooling effectiveness, and (2) film cooling on rocket engine performance were determined experimentally in a rocket thrust chamber assembly operating with hydrogen and oxygen propellants at 300 psi chamber pressure. Tests were conducted using hydrogen, helium, and nitrogen film coolants in an instrumented, thin walled, steel thrust chamber. The film cooling, performance loss, and heat transfer coefficient data were correlated with the ALRC entrainment film cooling model which relates film coolant effectiveness and mixture ratio at the wall to the amount of mainstream gases entrained with the film coolant in a mixing layer. In addition, a comprehensive thermal analysis computer program, HOCOOL, was prepared from previously existing ALRC computer programs and analytical techniques.

Rousar, D. C.↗

NASA NTRS1973

Hydrogen film cooling investigation

Effects of flow turning, flow acceleration, and supersonic flow on film cooling were determined experimentally and correlated in terms of an entrainment film cooling model. Experiments were conducted using thin walled metal test sections, hot nitrogen mainstream gas, and ambient hydrogen or nitrogen as film coolants. The entrainment film cooling model relates film cooling effectiveness to the amount of mainstream gases entrained with the film coolant in a mixing layer. The experimental apparatus and the analytical model used are described in detail and correlations for the entrainment fraction and film coolant-to-wall heat transfer coefficient are presented.

Rousar, D. C.↗

NASA NTRS1972

Hydrogen film/conductive cooling

Small scale nozzle tests using heated nitrogen were run to obtain effectiveness and wall heat transfer data with hydrogen film cooling. Effectiveness data are compared with an entrainment model developed from planar, unaccelerated flow data. Results indicate significant effects due to flow turning and acceleration. With injection velocity effects accounted for explicitly, heat transfer correlation coefficients were found to be the same with and without film cooling when properties are evaluated at an appropriate reference temperature for the local gas composition defined by the coolant effectiveness. A design study for an O2/H2 application with 300 psia (207 N/sq cm) chamber pressure and 1500 lbs (6670 N) thrust indicates an adiabatic wall design requires 4 to 5 percent of the total flow as hydrogen film cooling. Internal regenerative cooling designs were found to offer no reduction in coolant requirements.

Ewen, R. L.↗