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At least 1,711 records · Page 95

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NASA NTRS1972

Propellant sidefeed-short pulse discharge thruster studies

The technique of feeding a solid propellant into the discharge from the sides of the discharge was evaluated. The thrust/power ratio could be significantly effected by the included angle of V-shaped propellants and by the electrode length. This result implies that when results are compared at the same specific impulse it is possible to obtain higher thrust efficiencies. In particular, it was found that for a given discharge energy the thrust/power ratio correlated with propellant mass. Increasing the integral simultaneously increases both the gas dynamic and electromagnetic thrust. An analytic expression was formulated for ablated mass which comprehensively describes experimental data in terms of geometry and electrical parameters. The correlation of the product impulse x specific impulse with discharge energy was also described. It is suggested that the reliability of dry energy storage capacitors does not equal the reliability of liquid impregnated units when the comparison is made at the same joules/Kg rating.

Palumbo, D. J.↗

NASA NTRS1972

Grid-translation beam deflection systems for 5-cm and 30-cm diameter Kaufman thrusters

A 5-cm grid translation mechanism has been developed capable of 10 deg beam deflection. A 2026-hour endurance test was run at a preset 10 deg deflection angle and an extrapolated lifetime of better than 10,000 hours obtained. Response time data for grid translation are presented. Preliminary results for a 30-cm diameter system are given and results of a theoretical analysis of a dished grid system are discussed.

Lathem, W. C.↗

NASA NTRS1972

Exhaust flow and propulsion characteristics of a pulsed MPD arc thruster

An experimental investigation of the near-field megawatt, single-shot exhaust is presented for the self-field and auxiliary field cases (0-2 T). Plasma impact pressure and number density are correlated to provide velocity profiles (30,000 to 70,000 m/sec), thrust (10 to 120 N), impulse (3 to 16 N-sec) and mass accounting. The data agree with Hugel's self-field theory for the case where thrust is produced entirely by electromagnetic force. The data show that the thrust monotonically increases with auxiliary field.

Michels, C. J.↗

NASA NTRS1972

Quasi-linear theory of electron density and temperature fluctuations with application to MHD generators and MPD arc thrusters.

Fluctuations in electron density and temperature coupled through Ohm's law are studied for an ionizable medium. The nonlinear effects are considered in the limit of a third order quasi-linear treatment. Equations are derived for the amplitude of the fluctuation. Conditions under which a steady state can exist in the presence of the fluctuation are examined and effective transport properties are determined. A comparison is made to previously considered second order theory. The effect of third order terms indicates the possibility of fluctuations existing in regions predicted stable by previous analysis.

Smith, J. M.↗

NASA NTRS1972

Temporal survey of electron number density and electron temperature in the exhaust of a megawatt MPD-arc thruster.

Temporal and radial profiles are obtained 30 cm downstream from the anode for two peak arc currents (11.2 kA and 20 kA) and for various auxiliary magnetic fields (0, 1.0 T, and 2.0 T) using the Thomson scattering technique. Average density and temperature are relatively constant for over 100 microseconds with significant fluctuations. Radial profiles obtained are relatively flat for 4 cm from the axis. Compared to earlier 20 cm data, the exhaust density has decreased significantly, the average temperature has not changed, and the density ?hole' with an auxiliary magnetic field has enlarged.

Michels, C. J.↗

NASA NTRS1972

Optical radiation from regions downstream of mercury bombardment thrusters.

Use of a 0.5-m focal length, plane grating monochromator to measure the radiance of spectral radiation emanating from regions downstream of a mercury bombardment thrustor in the wavelength range investigated was 2800 to 6000 A. This radiation was due primarily to the radiative decay of excited mercury atoms exhausted from the thrustor. Radiance values ranged from 10 to the minus 11th to 10 to the min scm/ster varying with wavelength. For resonant radiation, the spectral radiance may exceed 10 to the minus 8th W/sq cm/ster. From such radiance measurements, it was concluded that the thrustor background radiation should not interfere with the control functions of a star tracker viewing through the thrustor exhaust, provided that the tracker is designed to operate with a sufficiently small field of view. Problems may be encountered, however, during the spacecraft acquisition phase where a larger field of view may be required. Here the thrustor exhaust radiation may be comparable to the star light flux. This problem may be circumvented by locating the tracker view axis so as not to view downstream of the thrustor.

Milder, N. L.↗

NASA NTRS1972

Electrothermal hydrazine thruster development.

Results of testing of several electrically-heated, thermal-decomposition hydrazine thrustors in the 5- to 70-mlb thrust range. Propellant supply pressures are compatible with those typical of larger catalytic thrustors. With a 5-W power input, pulsed specific impulse varies from about 165 to 215 sec, depending on duty cycle. The thrustors yield highly reproducible 1-mlb-sec impulse bits with 50-msec command pulse widths. The steady-state specific impulse exceeds 230 sec for thrust levels above 10 mlb.

Murch, C. K.↗

NASA NTRS1972

Ion thruster performance calibration.

The calibration of a typical 20-cm diameter ion thrustor was examined to determine performance penalties that must be assessed in projecting measured performance into a space environment. Four specific areas were investigated. These include (1) double ion content of the beam, (2) back ingestion from the vacuum facility, (3) beam spreading, and (4) propellant flow rate measurements. The double ion content was measured and found to be as high as 5.5% at an arc voltage of 35 volts. Back ingestion was observed to become significant above tank pressures of 6 microtorr. Beam spreading reduced effective thrust on the order of 2.5%.

Pawlik, E. V.↗

NASA NTRS1972

Grid-translation beam deflection systems for 5-cm and 30-cm diameter Kaufman thrusters.

A 5-cm grid translation mechanism has been developed capable of 10-deg beam deflection. A 2026-hour endurance test was run at a preset 10-deg deflection angle, and an extrapolated lifetime of better than 10,000 hours was obtained. Response time data for grid translation are presented. Preliminary results for a 30-cm diameter system are given, and results of a theoretical analysis of a dished grid system are discussed.

Lathem, W. C.↗

NASA NTRS1972

Exhaust flow and propulsion characteristics of a pulsed MPD arc thruster.

Experimental investigation of the near-field, megawatt, single-shot exhaust for the self-field and auxiliary field cases (0 to 2 tesla) for a pulsed MPD arc thrustor. Plasma impact pressure and number density are correlated to provide velocity profiles (30,000 to 70,000 m/sec), thrust (10 to 120 N), impulse (3 to 16 N-sec) and mass accounting. The data agree with Hugel's self-field theory for the case where thrust is produced entirely by electromagnetic force. The data show that the thrust monotonically increases with auxiliary field.

Michels, C. J.↗

NASA NTRS1972

Optical properties of thermal control coatings contaminated by MMH/N2O4 5-pound thruster in a vacuum environment with solar simulation.

Cat-a-lac Black and S13G thermal control coatings were exposed to the exhaust of a thrustor in a simulated space environment. Vacuum was maintained at less than 10 microtorr during thrustor firing in the liquid helium cooled facility. The thrustor was fired in a 50-millisecond pulse mode, and the accumulated firing time was 224 seconds. Solar absorptance and thermal emittance of the coatings were measured in-situ at intervals of 300 pulses, using a calorimetric technique. The Cat-a-lac Black coatings showed no change in solar absorptance or thermal emittance. The S13G showed up to 25% increase in solar absorptance but no change in thermal emittance.

Sommers, R. D.↗

NASA NTRS1972

Ion-thruster propellant utilization.

Maximum propellant utilization in a mercury electron-bombardment thrustor is evaluated. The primary-electron region in the ion chamber of a bombardment thrustor is analyzed at maximum utilization. Both the analysis and experimental data from a range of ion-chamber configurations show a nearly constant loss rate for un-ionized propellant at maximum utilization over a wide range of total propellant flow rate. The discharge loss level of 1000 eV/ion was used to define maximum utilization. The exact level of this definition has no effect on the qualitative results and little effect on the quantitative results. The results obtained are particularly significant whenever efficient throttled operation is required.

Kaufman, H. R.↗

NASA NTRS1972

Development of a multikilowatt ion thruster power processor

A feasibility study was made of the application of silicon-controlled, rectifier series, resonant inverter, power conditioning technology to electric propulsion power processing operating from a 200 to 400 Vdc solar array bus. A power system block diagram was generated to meet the electrical requirements of a 20 CM hollow cathode, mercury bombardment, ion engine. The SCR series resonant inverter was developed as a primary means of power switching and conversion, and the analog signal-to-discrete-time-interval converter control system was applied to achieve good regulation. A complete breadboard was designed, fabricated, and tested with a resistive load bank, and critical power processor areas relating to efficiency, weight, and part count were identified.

Schoenfeld, A. D.↗

NASA NTRS1973

Hydrogen-oxygen auxiliary propulsion for the space shuttle. Volume 1: High pressure thrusters

Technology for long life, high performing, gaseous hydrogen-gaseous oxygen rocket engines suitable for auxiliary propulsion was provided by a combined analytical and experimental program. Propellant injectors, fast response valves, igniters, and regeneratively and film-cooled thrust chambers were tested over a wide range of operating conditions. Data generated include performance, combustion efficiency, thermal characteristics film cooling effectiveness, dynamic response in pulsing, and cycle life limitations.

Source record↗

NASA NTRS1973

Investigation of hollow cathode performance for 30-cm thrusters

A parametric investigation of 6.35 mm diameter mercury hollow cathodes was carried out in a bell jar. The parameters that were varied were the amount of initial emissive mix, insert position, emission current, cathode temperature, orifice diameter, and mercury flow rate. Flow characteristic curves and performance as a function of time were obtained for the various cathodes. The results of a 3880 hr life test of a main cathode run at 15 amps emission current with no noticeable changes in keeper and collector voltages are also presented.

Mirtich, M. J.↗

NASA NTRS1973

Exhaust pressure and density of various pulsed mpd-arc thruster systems

Exhaust flow in a new, 155 cm i.d., vacuum facility was compared with earlier measurements in a small (15.2 cm i.d.) duct. Reductions in post transient impact pressure were about 5:1 in the larger facility. Corresponding reduced electron number densities (approximately 2 x 10 to the 13th power/cucm) were noted. A new, 125 miscrosecond, pulse forming network power source produced no major differences in impact pressure compared to the crowbarred condenser bank used earlier. Comparing a puff gas feed of the arc chamber with a new 10 millisecond steady gas feed also showed no major difference in impact pressure for 125 miscrosecond powering.

Michels, C. J.↗

NASA NTRS1972

Development of a radioisotope-fueled thruster for satellite propulsion.

Description of the design and testing of two distinct thrustor concepts which use ammonia propellant and have a specific impulse as high as 310 sec during pulsed operation. Considerations of thrustor design, material selection, and system testing are discussed in the light of performance, lifetime, and nuclear safety requirements. Results of a detailed study of the advantages and disadvantages of the use of such a thrustor are presented.

Rosen, S. G.↗

NASA NTRS1972

Space simulation experiments on reaction control system thruster plumes.

A space simulation procedure was developed for studying rocket plume contamination effects using a 5-lb bipropellant reaction control system thrustor. Vacuum chamber pressures of 0.00003 torr (70 miles altitude) were achieved with the thrustor firing in pulse trains consisting of eight pulses - 50 msec on, 100 msec off, and seven minutes between pulse trains. The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid-helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.

Cassidy, J. F.↗