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Goldstein, R.

Publications and source records attributed to Goldstein, R..

At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Fiber optic rotation sensor /FORS/ signal detection and processing

The recent development of low-loss single-mode optical fiber waveguides for light has made possible a new class of inertial reference devices built on the principle of a closed loop interferometer. Light circulating through the loop in both directions experiences a relative phase delay proportional to rotation rate about the loop axis. This paper derives the phase delay and discusses signal detection, signal processing techniques and error sources. It is concluded that synchronous modulation and demodulation and an active gain control at the signal calculation level are required to eliminate drift errors. Potential performance is extraordinarily good; rotation rate sensitivity of a few milli-arc seconds per second and angular position random walk errors of an arc second per square root hour appear feasible.

Goss, W. C.↗
NASA NTRS1979

Single mode glass fiber welding

The electric-arc welding of commercially available single-mode optical fiber has been demonstrated. A mean transmission of 92% and a maximum transmission of 98% are reported for welds of fiber waveguide of 4.5 microns core diameter.

Nelson, M. D.↗
NASA NTRS1978

IAU nomenclature for albedo features on the planet Mercury

The International Astronomical Union has endorsed a nomenclature for the albedo features on Mercury. Designations are based upon the mythological names related to the god Hermes; they are expressed in Latin form. The dark-hued albedo features are associated with the generic term Solitudo. The light-hued areas are designated by a single name without generic term. The 32 names adopted are allocated on the Mercury map.

Dollfus, A.↗
NASA NTRS1978

Solar sailing - The concept made realistic

The practicability of using solar sails as a means of spacecraft propulsion is discussed with reference to a Haley's comet rendezvous mission in 1981. Two types of sails are examined: square and heliogyro. Because of ease of deployment (release into space from a space shuttle, and unfolding of the blades through solar radiation pressure) the latter is preferred. Structural and mission parameters are given for the heliogyro configuration, with attention to operational temperature limits and material coefficients. A cranking orbit is described, where a sail-powered spacecraft flies first toward the sun, and then away from it in order to achieve rendezvous with a comet. Other missions using solar sails, including a back-up mission to Encke's comet in 1983, are also discussed.

Friedman, L.↗
NASA NTRS1978

Fiber optic rotation sensor /FORS/ laboratory performance evaluation

A single axis rotation sensor, utilizing a 200 m single mode optical fiber in a 30 cm diameter coil has been constructed and tested in the laboratory. A He-Ne laser is the light source for the Sagnac interferometer-mode sensor. The outputs of Si detectors viewing two complementary interference patterns are electronically differenced to obtain a rotation-dependent signal. Linear response over a rotation rate range of about 0.0001 to 0.1 rad/sec was obtained.

Goldstein, R.↗
NASA NTRS1977

Active control of potential of the geosynchronous satellites ATS-5 and ATS-6

The ATS-6 geosynchronous satellite carries two cesium ion thrusters as part of the technology experiment package. Their effectiveness in controlling the spacecraft potential was shown during normal operation of the thrusters. Auroral particle detectors were used to estimate spacecraft potential, amount of cesium propellant backflow to the spacecraft, and electron signature of the thruster. Operation of the thruster clamps the spacecraft to about -10 v in the presence of a wide range of ambient particle flux. Some positive ions from the thruster are returned to the spacecraft during this period. The plasma bridge neutralizer of the thruster is also effective in controlling the spacecraft potential. Laboratory simulation tests of the ATS-5 ion thruster thermionic emitter neutralizer showed that negative differential charging on the spacecraft surface probably prevents the emitter from completely discharging the spacecraft in flight.

Goldstein, R.↗
NASA NTRS1977

Plasma distribution and spacecraft charging modeling near Jupiter

To assess the role of spacecraft charging near Jupiter, the plasma distribution in Jupiter's magnetosphere was modeled using data from the plasma analyzer experiments on Pioneer 10 (published results) and on Pioneer 11 (preliminary results). In the model, electron temperatures are kT = 4 eV throughout, whereas proton temperatures range over 100 or equal to kT or equal to 400 eV. The model fluxes and concentrations vary over three orders of magnitude among several corotating regions, including, in order to increasing distance from Jupiter, a plasma void, plasma sphere, sporadic zone, ring current, current sheet, high latitude plasma and magnetosheath. Intermediate and high energy electrons and protons (to 100 MeV) are modeled as well. The models supply the information for calculating particle fluxes to a spacecraft in the Jovian environment. The particle balance equations (including effects of secondary and photoemission) then determine the spacecraft potential.

Goldstein, R.↗
NASA NTRS1977

Active spacecraft potential control system selection for the Jupiter orbiter with probe mission

It is shown that the high flux of energetic plasma electrons and the reduced photoemission rate in the Jovian environment can result in the spacecraft developing a large negative potential. The effects of the electric fields produced by this charging phenomenon are discussed in terms of spacecraft integrity as well as charged particle and fields measurements. The primary area of concern is shown to be the interaction of the electric fields with the measuring devices on the spacecraft. The need for controlling the potential of the spacecraft is identified, and a system capable of active control of the spacecraft potential in the Jupiter environment is proposed. The desirability of using this system to vary the spacecraft potential relative to the ambient plasma potential is also discussed. Various charged particle release devices are identified as potential candidates for use with the spacecraft potential control system. These devices are evaluated and compared on the basis of system mass, power consumption, and system complexity and reliability.

Beattie, J. R.↗
NASA NTRS1977

Active control of electric potential of spacecraft

Techniques are discussed for controlling the potential of a spacecraft by means of devices which release appropriate charged particles from the spacecraft to the environment. Attention is given to electron emitters, ion emitters, a basic electron emitter arrangement, techniques for sensing electric field or potential, and flight experiments on active potential control. It is recommended to avoid differential charging on spacecraft surfaces because it can severely affect the efficacy of emitters. Discharging the frame of a spacecraft with dielectric surfaces involves the risk of stressing the dielectric material excessively. The spacecraft should, therefore, be provided with grounded conductive surfaces. It is pointed out that particles released by control systems can return to the spacecraft.

Goldstein, R.↗
NASA NTRS1976

Active control of spacecraft potentials at geosynchronous orbit

Tests have been conducted concerning the active control of the potentials of the geosynchronous satellites ATS-5 and ATS-6. The ATS-5 tests show that a simple electron emitter can be used to reduce the magnitude of the potential of a spacecraft which has been charged negatively by the environment. The ATS-6 ion thruster had also a pronounced effect on the potential barrier. In this case, the flux of high-energy primary ions and of low-charge exchange ions produces a space-charge neutralization effect which the electron gun alone cannot achieve.

Goldstein, R.↗
NASA NTRS1975

Production of CO /a3 Pi/ in a mercury discharge

The vibrational intensity distribution of CO (a3 Pi - X1 Sigma+) Cameron bands was measured. Data combined from the present and from previous experiments suggest that a3 Pi vibrational temperatures are more characteristic of the parent molecule than are the details of the excitation mechanism. The temperatures were obtained using Franck-Condon factors to deconvolute the bands at 194, 199, and 206 nm. The values were 4800 K for CO2 and 1900 K for CO.

Monahan, K. M.↗
NASA NTRS1975

Optical properties of mercury ion thruster exhaust plumes Significance for candidate SEP science instruments

Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume emission rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter was considered incompatible with simultaneous thruster operation.

Goldstein, R.↗
NASA NTRS1975

Spacecraft charging control demonstration at geosynchronous altitude

ATS-6, which was launched on May 30, 1974, has now completed its first eclipse season. Potentials as great as 11,000 volts were observed between the spacecraft and ambient plasma. Both ATS-5 and 6 have ion engine experiments onboard. These experiments have been used to clamp the spacecraft potential slightly negative when emitting an ion beam and momentarily discharge the spacecraft when only operating the ion engine's neutralizing electron source. Descriptions of the neutralizer's operation and their effect on potential are presented. In addition, spacecraft charging data for ATS-5 and 6 are presented and correlated over the same eclipse season.

Bartlett, R. O.↗
NASA NTRS1974

Optical properties of mercury ion thruster exhausts and implications for science instruments

Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. Most of the emitted radiation came from UV lines of excited mercury atoms and ions, with no observable continuum emission. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter, was considered incompatible with simultaneous thruster operation.

Monahan, K. M.↗
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

Cathode effects on thrust subsystem performance predictability.

Development of a model of mercury vapor electron bombardment thrustor hollow cathode operation. The proposed model predicts that the primary electron emission mechanism is thermionic when sufficient low work function material is present, that the cathode is heated by power extracted from the plasma to a temperature sufficient to emit the current demanded by the external circuit; that for a given total arc current, attempts to reduce cathode temperature by changes in the external thermal coupling will affect only the amount of power extracted from the plasma; and that depletion of low work function material in the cathode results in higher cathode temperature.

Goldstein, R.↗