Engineering topics

Hyman, J., Jr.

Publications and source records attributed to Hyman, J., Jr..

At least 19 records

NASA NTRS1984

Ion thruster system (8-cm) cyclic endurance test

This report describes the qualification test of an Engineering-Model 5-mN-thrust 8-cm-diameter mercury ion thruster which is representative of the Ion Auxiliary Propulsion System (IAPS) thrusters. Two of these thrusters are scheduled for future flight test. The cyclic endurance test described herein was a ground-based test performed in a vacuum facility with a liquid-nitrogen-cooled cryo-surface and a frozen mercury target. The Power Electronics Unit, Beam Shield, Gimal, and Propellant Tank that were used with the thruster in the endurance test are also similar to those of the IAPS. The IAPS thruster that will undergo the longest beam-on-time during the actual space test will be subjected to 7,055 hours of beam-on-time and 2,557 cycles during the flight test. The endurance test was successfully concluded when the mercury in the IAPS Propellant Tank was consumed. At that time, 8,471 hours of beam-on-time and 599 cycles had been accumulated. Subsequent post-test-evaluation operations were performed (without breaking vacuum) which extended the test values to 652 cycles and 9,489 hours of beam-on-time. The Power Electronic Unit (PEU) and thruster were in the same vacuum chamber throughout the test. The PEU accumulated 10,268 hr of test time with high voltage applied to the operating thruster or dummy load.

Dulgeroff, C. R.↗

NASA NTRS1984

IAPS (8-cm) ion thruster cyclic endurance test

Attention is given to the cyclic endurance qualification test, performed in a vacuum facility with a liquid nitrogen-cooled cryosurface and a frozen mercury target, to which a mercury-ion thruster representative of Ion Auxiliary Propulsion System (IAPS) thrusters was subjected. At the time that the endurance test was successfully concluded, upon consumption of the IAPS propellant tank's mercury, 8,471 hours of 'beam-on time' and 599 cycles had been accumulated. The system's Power Electronics Unit accumulated 10,268 hours of test time with high voltage applied to the operating thruster or dummy load.

Dulgeroff, C. R.↗

NASA NTRS1978

Engineering Model 8-cm Thruster System

Development of an Engineering Model 8-cm Mercury Ion Thruster System for Satellite Control has been successfully completed. This system operates at a specific impulse in excess of 2600 sec, produces a thrust of 5 mN with a total input power of 165 W; it has a dry mass of 16.6 kg and a mercury-propellant-reservoir capacity of 8.75 kg. This paper summarizes the development work, the system characteristics and performance, and the testing undertaken to verify the design.

Herron, B. G.↗

NASA NTRS1976

Modularized ion thruster

A modularized ion thruster system was developed for space-propulsion applications. Separate discharge chamber modules (DCMs) were optimized for operation at the thrust levels of T = 0.5 mlb and T = 2 mlb to accommodate the extended thrust range. These optimizations included modifications in the discharge-chamber components and the incorporation of ion machined accelerators in the beam-extraction systems. Performance of the optimized modules are summarized.

Hyman, J., Jr.↗

NASA NTRS1976

Modularized ion thruster development

A family of mercury ion thruster modules has been developed which extends the basic design of the Hughes-developed 1-mlb Engineering Model Thruster to accommodate the extended thrust range from 0.5 mlb to 4 mlb. In the subject program, separate Discharge Chamber Modules (DCM's) have been optimized for operation at nominal thrust levels T = 0.5 mlb and 2 mlb; DCM optimization is continuing at thrust level T = 4 mlb. Performance optimization required modification of the beam-extraction system and of discharge-chamber components; however, the cathodes and mechanical structure are unchanged from the Engineering Model design. Performance data on the optimized modules are given.

Hyman, J., Jr.↗

NASA NTRS1976

Development of an 8-cm engineering model thruster system

Electric propulsion has been shown to offer major advantages over the techniques currently employed for the control of earth satellites. For a user to realize these advantages, however, requires the availability of a proven, operationally flight-ready propulsion system. Currently an Engineering Model of an 8-cm ion thruster propulsion system is under development. The system includes the thruster unit with its associated reservoir, thruster gimbaling subsystem, and power processing unit. This paper describes the EM System with special emphasis on hardware design and system performance.

Herron, B. G.↗

NASA NTRS1975

One-millipound mercury ion thruster

A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.

Hyman, J., Jr.↗

NASA NTRS1974

Eight cm technology thruster development

A structural integrated ion thruster with 8-cm beam diameter (SIT-8) was developed for attitude control and stationkeeping of synchronous satellites. As optimized, the system demonstrates a thrust T=1.14 mlb (not corrected for beam V sub B = 1200 V (I sub sp = 2200 sec) total propellant utilization efficiency nu sub u = 59.8% (is approximately 72% without auxiliary pulse-igniter electrode), and electrical efficiency n sub E 61.9%. The thruster incorporates a wire-mesh anode and tantalum cover surfaces to control discharge chamber flake formation and employs an auxiliary pulse-igniter electrode for hollow-cathode ignition. When the SIT-8 is integrated with the compatible SIT-5 propellant tankage, the system envelope is 35 cm long by 13 cm flange bolt circle with a mass of 9.8 kg including 6.8 kg of mercury propellant. Two thrust vectoring systems which generate beam deflections in two orthogonal directions were also developed under the program and tested with the 8-cm thruster. One system vectors the beam over + or - 10 degrees by gimbaling of the entire thruster (not including tankage), while the other system vectors the beam over + or - 7 degrees by translating the accel electrode relative to the screen electrode.

Hyman, J., Jr.↗

NASA NTRS1973

Satellite control mercury ion thruster

Program status is reviewed for development of an 8-cm structurally integrated ion thruster for application to attitude control and stationkeeping of synchronous satellites. The system is designed to achieve a thrust level of 1 and 2 mlb at a beam voltage of 1200 V, a propellant-utilization efficiency of 85%, and electrical efficiency of 75%. Thrust vectoring will be implemented by a dished-gird roll misalignment system. Design features are incorporated which ensure thruster life of 20,000 hours, with 10,000 thruster-start cycles and 100,000 beam-deflection cycles.

Hyman, J., Jr.↗

NASA NTRS1973

Performance optimized, small structurally integrated ion thruster system

A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS1972

SIT-5 system development.

A 5-cm structurally integrated ion thruster (SIT-5) has been developed for attitude control and stationkeeping of synchronous satellites. With two-dimension thrust-vectoring grids, a first generation unit has demonstrated a thrust of 0.56 mlb at a beam voltage of 1200 V, total mass efficiency of 64%, and electrical efficiency of 46.8%. Structural integrity is demonstrated with a dielectric-coated grid for shock (30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.8 cm long by 13.9 cm flange bolt circle, with a mass of 8.5 kg, including 6.2 kg mercury propellant. Characteristics of a second-generation unit indicate significant performance gains.

Hyman, J., Jr.↗

NASA NTRS1971

Design and development of a small structurally integrated ion thruster system

A 5 cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. With 2-D vectorable grids, thrust is 0.56 mlb at a beam voltage of 1200 V, total mass efficiency is 64%, and electrical efficiency is 46.8%. Nonvectoring thrust of 0.41 is demonstrated at very low specific impulse with dielectric coated grids. Structural integrity is demonstrated with dielectric coated grids for shock (30 G), sinusoidal (9 G), and random (19 G squared/Hz) accelerations. System envelope is 31.6 cm long by 14 cm flange B.C. with a mass of 8.5 kg including 6.2 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS1970