Analysis and design of ion thruster for large space systems

Design analyses showed that an ion thruster of approximately 50 cm in diameter will be required to produce a thrust of 0.5 N using xenon or argon as propellants, and operating the thruster at a specific impulse of 3530 sec or 6076 sec respectively. A multipole magnetic confinement discharge chamber was specified

Retrofit and verification test of a 30-cm ion thruster

Twenty modifications were found to be necessary and were approved by design review. These design modifications were incorporated in the thruster documents (drawings and procedures) to define the J series thruster. Sixteen of the design revisions were implemented in a 900 series thruster by retrofit modification. A standardized set of test procedures was formulated, and the retrofit J series thruster design was verified by test. Some difficulty was observed with the modification to the ion optics assembly, but the overall effect of the design modification satisfies the design objectives. The thruster was tested over a wide range of operating parameters to demonstrate its capabilities

Primary Electric Propulsion Technology Study

An investigation of the 30-cm engineering-model-thruster technology with emphasis placed on the development of models for understanding and predicting the operational characteristics and wear-out mechanisms of the thruster as a function of operating or design parameters is presented. The task studies include: (1) the wear mechanisms and wear rates that determine the useful lifetime of the thruster discharge chamber; (2) cathode lifetime as determined by the depletion of barium from the barium-aluminate-impregnated-porous-tungsten insert that serves as a barium reservoir; (3) accelerator-grid-system technology; (4) a verification of the high-voltage propellant-flow-electrical-isolator design developed under NASA contract NAS3-20395 for operation at 10-kV applied voltage and 10-A equivalent propellant flow with mercury and argon propellants. A model was formulated for predicting performance

Evaluation of the use of on-board spacecraft energy storage for electric propulsion missions

On-board spacecraft energy storage represents an under utilized resource for some types of missions that also benefit from using relatively high specific impulse capability of electric propulsion. This resource can provide an appreciable fraction of the power required for operating the electric propulsion subsystem in some missions. The most probable mission requirement for utilization of this energy is that of geostationary satellites which have secondary batteries for operating at high power levels during eclipse. The study summarized in this report selected four examples of missions that could benefit from use of electric propulsion and on-board energy storage. Engineering analyses were performed to evaluate the mass saved and economic benefit expected when electric propulsion and on-board batteries perform some propulsion maneuvers that would conventionally be provided by chemical propulsion. For a given payload mass in geosynchronous orbit, use of electric propulsion in this manner typically provides a 10% reduction in spacecraft mass

Extended performance solar electric propulsion thrust system study. Volume 1: Executive summary

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed. Confirmation testing and analysis of thruster and power-processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. A program development plan was formulated that outlines the work structure considered necessary for developing, qualifying, and fabricating the flight hardware for the baseline thrust system within the time frame of a project to rendezvous with Halley's comet. An assessment was made of the costs and risks associated with a baseline thrust system as provided to the mission project under this plan. Critical procurements and interfaces were identified and defined. Results are presented

Retrofit and acceptance test of 30-cm ion thrusters

Six 30 cm mercury thrusters were modified to the J-series design and evaluated using standardized test procedures. The thruster performance meets the design objectives (lifetime objective requires verification), and documentation (drawings, etc.) for the design is completed and upgraded. The retrofit modifications are described and the test data for the modifications are presented and discussed

Characteristics of 30-centimeter mercury ion thrusters

The technology development of the 30 centimeter J series mercury ion thruster for prime propulsion application in solar electric propulsion systems is described. Thruster design is reviewed. A standardized set of test and data recording procedures formulated to allow for the characterization of the J series thruster is described. Characteristics measured are the magnetic baffle characterization, the neutralizer characterization, perveance, the minimum eV/ion measurement, and the electrical and propellant utilization efficiency measurements. Test results are presented

The 2.5 kW advanced technology ion thruster

A representative thruster was extensively documented with respect to performance parameters and characteristics at selected ion beam currents in the 0.5 to 2.75 A range, including measurements of thrust losses resulting from doubly-charged ions and ion beam divergence. Corrected total efficiency was shown to be relatively insensitive to operating parameter selection at any given power level. Factors affecting doubly-charged ionization were studied and it was found that the fraction of doubly-charged ions is directly proportional to the discharge chamber propellant utilization. The parameter that most affects this proportionality is the accel aperture diameter (which controls neutral atom loss). Thruster-power conditioner interactions were studied with the result that previous power supply specifications remain satisfactory. Options for reducing the number of power supplies required were demonstrated to be feasible. Gimbal actuator designs were studied with the goal of selecting a particular approach for design and development. The conclusion drawn was that optimum gimbal actuator design depends heavily on the thruster application and consequently the effort was concluded by developing a computer program to aid in specifying the gimbal requirements for the thrust vectoring required in a specific application

HOLLOTRON switch for megawatt lightweight space inverters

The feasibility of satisfying the switching requirements for a megawatt ultralight inverter system using HOLLOTRON switch technology was determined. The existing experimental switch hardware was modified to investigate a coaxial HOLLOTRON switch configuration and the results were compared with those obtained for a modified linear HOLLOTRON configuration. It was concluded that scaling the HOLLOTRON switch to the current and voltage specifications required for a megawatt converter system is indeed feasible using a modified linear configuration. The experimental HOLLOTRON switch operated at parameters comparable to the scaled coaxial HOLLOTRON. However, the linear HOLLOTRON data verified the capability for meeting all the design objectives simultaneously including current density (greater than 2 A/sq cm), voltage (5 kV), switching frequency (20 kHz), switching time (300 ns), and forward voltage drop (less than or equal to 20 V). Scaling relations were determined and a preliminary design was completed for an engineering model linear HOLLOTRON switch to meet the megawatt converter system specifications

Electric propulsion - characteristics, applications, and status

A comparative review of the principles of ion thruster and chemical rocket operations is presented. The 30cm mercury ion thruster development and the specifications imposed on it by the Solar Electric propulsion System program are discussed. The 30cm thruster operating range, efficiency, wear out lifetime, and interface requirements are described