SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
FORTRAN 4 subroutines for calculating critical flow factor to determine mass flow rate of gaseous helium and nitrogen through critical flow nozzles
Axial running clearance effect on efficiency and mass flow rate of Brayton cycle radial inflow turbines
Mass flow rate calculation of methane and natural gas mixtures through critical flow nozzles
Colloid annular thrustor performance, using low mass flow rate data and space charge formulas
Tube heat transfer augmentation by helical vane inserts, noting Reynolds number and mass flow rate effects
Rocket engine injector orifice to accommodate changes in density, velocity, and pressure, thereby maintaining constant mass flow rate of propellant into rocket combustion chamber
Methods for determining mass flow rates of gases for use in calibrating gas flowmeters
Small curvature radius/throat radius ratio supersonic nozzles mass flow rate coefficients at high Reynolds numbers, appraising isentropic flow prediction methods
Real gas mass flow rate computation through sonic nozzle
Methane and natural gas flow through critical flow nozzles, calculating real gas effects on mass flow rate
Subroutines, FORTRAN 4 type, were developed for calculating isentropic natural gas mass flow rate through nozzle. Thermodynamic functions covering compressibility, entropy, enthalpy, and specific heat are included.
A computer program was developed which performs an error analysis of a minimum-fuel, finite-thrust, transfer maneuver between two Keplerian orbits in the vicinity of Mars. The method of analysis is the Monte Carlo approach where each off-nominal initial orbit is targeted to the desired final orbit. The errors in the initial orbit are described by two covariance matrices of state deviations and tracking errors. The function of the program is to relate these errors to the resulting errors in the final orbit. The equations of motion for the transfer trajectory are those of a spacecraft maneuvering with constant thrust and mass-flow rate in the neighborhood of a single body. The thrust vector is allowed to rotate in a plane with a constant pitch rate. The transfer trajectory is characterized by six control parameters and the final orbit is defined, or partially defined, by the desired target parameters. The program is applicable to the deboost maneuver (hyperbola to ellipse), orbital trim maneuver (ellipse to ellipse), fly-by maneuver (hyperbola to hyperbola), escape maneuvers (ellipse to hyperbola), and deorbit maneuver.
Data are presented, in plotted form, of tests for determining the interference effects of space shuttle booster cruise engine location for power-on and power-off conditions. The tests were conducted in a 7 x 10 foot transonic wind tunnel; the model was a 0.015-scale space shuttle booster specially equipped for propulsion effects testing. Data were obtained over a Mach number range of 0.4 to 1.13 at angles of attack from -4 deg to 20 deg at zero degrees sideslip and at angles of sideslip from -6 deg to +6 deg at constant angles of attack of 0 deg, 6 deg, 15 deg, and in some cases 10 deg. Additional parameters investigated were: elevon deflection, canard deflection, aileron deflection, rudder deflection, canard position, and mass flow rate.
Measurements of arc voltage and energy loss to the cooled electrodes of a magnetoplasmadynamic (MPD) arc, operating without an applied magnetic field, were made at chamber pressures of 26 to 950 torr, argon mass flow rates of 0.08 to 44 g/s and current of 200 to 2000 A. The resulting arc thermal efficiency varied from 22% at a chamber pressure of 26 torr to 88% at 950 torr. Thermal efficiency was only weakly dependent on arc current. It is concluded that the MPD arc operating without an applied magnetic field and at higher pressure than normally used in thruster applications is a reliable and efficient steady-state plasma source.
Spectroscopic and photographic investigations are reported that reveal a complex azimuthal species structure in the exhaust plume of a quasi-steady argon MPD accelerator. Over a wide range of operating conditions the injected argon remains collimated in discrete jets which are azimuthally in line with the six propellant injector orifices. The regions between these argon jets, including the central core of the exhaust flow, are occupied by impurities such as carbon, hydrogen and oxygen ablated from the Plexiglas back plate of the arc chamber. The features of this plume structure are found to be dependent on the arc current and mass flow rate. It is found that nearly half the observed velocity is attained in an acceleration region well downstream of the region of significant electromagnetic interaction. Recombination calculations show that the ionization energy is essentially frozen.
Experimental and theoretical study of the performance of a gas-core nuclear rocket, showing that in a uranium plasma nuclear rocket with an 8-ft cavity diameter the cavity wall can be cooled up to a power level of 7400 MW when the rocket operates at a pressure of 1000 atm and the propellant mass flow rate is 10 lbm/sec. A maximum cavity specific impulse of 5800 sec was obtained under such operating conditions. The fact that the wall heat flux was much lower for reactor powers below this level is linked to the presence of a relatively cool, opaque insulating layer of seeded propellant between the hot plasma and the solid wall. The additional coolant required for nozzle protection reduced the maximum cavity specific impulse to 5200 sec.
A unique method for solving the design problem of determining the areas required to meet an imposed pressure differential limit is presented. The associated analytical technique for determining the transient pressure differentials in a multicompartment series-parallel connected system is also presented. In an effort to verify the results obtained by using the analytical methods presented, a test was conducted using a seven compartment system. It was found in the investigations that a computer program based on the analytical model described can determine the pressure, temperature, pressure differentials, and mass flow rates into and out of each compartment at subsonic or sonic velocities.
The design of cryogenic flow control components, specifically the prevalve, for the Saturn 5 vehicle is discussed. Design systems are assessed and component requirements revealed significant improvements in mass flow rate, temperature, and pressure.