Downloads & Free Reading Options - Results

The potential use of an Inertial Electrostatic Confinement (IEC) power source for space propulsion has previously been suggested by the authors and others. In the past, these discussions have generally followed the charged-particle electric-discharge engine (QED) concept proposed by Bussard, in which the IEC is used to generate an electron beam which vaporizes liquid hydrogen for use as a propellant. However, in the present study, we consider an alternate approach, using the IEC to drive a conventional electric thruster unit. This has the advantage of building on the rapidly developing technology for such thrusters, which operate at higher specific impulse. Key issues related to this approach include the continued successful development of the physics and engineering of the IEC unit, as well as the development of efficient step-down dc voltage transformers. The IEC operates by radial injection of energetic ions into a spherical vessel. A very high ion density is created in a small core region at the center of the vessel, resulting in extremely high fusion power density in the core. Present experiments at the U. of Illinois in small IEC devices (less than 60-cm. dia.) have demonstrated much of the basic physics underlying this concept, e.g. producing approximately 10(exp 6) D-D neutrons/sec steady-state with deuterium gas flow injection. The ultimate goal is to increase the power densities by several orders of magnitude and to convert to D-He-3 injection. If successful, such an experiment would represent a milestone proof-of-principle device for eventual space power use. Further discussion of IEC physics and status will be presented with a description of the overall propulsion system and estimated performance.

tecnologie segrete

Electrostatic propulsion using C60 as the propellant has the potential to efficiently produce specific impulses in the range from 1000 to 3000 sec, and is therefore attractive for orbit transfer and station keeping missions. However, the viability of this concept depends on properties of the C60 molecule such as the ionization cross section, thermal stability, and resistance to fragmentation in a plasma environment. When this research began, insufficient information about these new materials was available. The objective of this program has been to determine properties of fullerenes relevant for electrostatic propulsion, to demonstrate ion extraction from a discharge, and to assess the implications for fullerene ion thrusters. The experiments we have undertaken fall into several distinct categories: time of flight mass spectrometry, electron energy loss spectroscopy, Fourier transform infrared spectroscopy, thermal stability, polarizability, and discharge chamber experiments.

An Inertial-Electrostatic Confinement (IEC) device was assembled at the Marshall Space Flight Center (MSFC) Propulsion Research Center (PRC) to study the possibility of using IEC technology for deep space propulsion and power. Inertial-Electrostatic Confinement is capable of containing a nuclear fusion plasma in a series of virtual potential wells. These wells would substantially increase plasma confinement, possibly leading towards a high-gain, breakthrough fusion device. A one-foot in diameter IEC vessel was borrowed from the Fusion Studies Laboratory at the University of Illinois @ Urbana-Champaign for the summer. This device was used in initial parameterization studies in order to design a larger, actively cooled device for permanent use at the PRC.

The operating characteristics of three types of atomic ion engines and three types of colloidal ion engines are evaluated. The major theoretical factors which will limit payload capacity are brought out for each type of engine. Total propulsion system masses for the production of ten pounds of thrust for 2,000 hours and for 10,000 hours of engine operation are compared. Examination of the payload capacities of the various engines for several lunar and planetary missions shows that the colloidal ion engines have a superior payload capacity for lunar missions and may be competitive with atomic ion engines for planetary missions. The specific mass of the power supply system is shown to be a much more dominant factor in determining the payload capacities of the atomic ion engines than for the colloidal ion engines.

The potential use of an INERTIAL ELECTROSTATIC CONFINEMENT (IEC) power source for space propulsion has previously been suggested by the authors and others. In the past, these discussions have generally followed the charged-particle electric-discharge engine (QED) concept proposed by Bussard, in which the IEC is used to generate an electron beam which vaporizes liquid hydrogen for use as a propellant. However, an alternate approach is considered, using the IEC to drive a 'conventional' electric thruster unit. This has the advantage of building on the rapidly developing technology for such thrusters, which operate at higher specific impulse. Key issues related to this approach include the continued successful development of the physics and engineering of the IEC unit, as well as the development of efficient step-down dc voltage transformers. The IEC operates by radial injection of energetic ions into a spherical vessel. A very high ion density is created in a small core region at the center of the vessel, resulting in extremely high fusion power density in the core. Experiments at the U. of Illinois in small IEC devices (is less than 60 cm. dia.) demonstrated much of the basic physics underlying this concept, e.g. producing 10(exp 6) D-D neutrons/sec steady-state with deuterium gas flow injection. The ultimate goal is to increase the power densities by several orders of magnitude and to convert to D-He-3 injection. If successful, such an experiment would represent a milestone proof-of-principle device for eventual space power use. Further discussion of IEC physics and status are presented with a description of the overall propulsion system and estimated performance.

Tests were conducted to determine the effects of electrostatic propulsion beam divergence effects on spacecraft surfaces. The subjects discussed are: (1) sensitive surfaces on the ATS 6 spacecraft, (2) the cesium ion source and testing facility, (3) cesium ion effects on thermophysical properties, and (4) simulated charge-exchange ion exposure. The compatibility of the ATS 6 ion engine experiment with the engineering subsystems and other experiments aboard the ATS 6 spacecraft was analyzed.

The third phase of a program to develop understanding of and tolerance-level criteria for the deleterious effects of electrostatic rocket exhaust (Cs, Cs(+), Hg, Hg(+)) and materials of rocket construction impinging on typical classes of spacecraft (S/C) surfaces was completed. Models of ion engine effluents and models describing the degradation of S/C surfaces by these effluents are presented. Experimental data from previous phases are summarized and Phase 2 data and analysis are presented in detail. The spacecraft design implications of ion engine contaminants are discussed.

The analysis of ion exhaust beam current flow for multiply charged ion species and the application to propellant utilization for the thruster are discussed. The ion engine in use in the experiments is a twenty centimeter diameter electromagnet electron bombardment engine. The experimental technique to determine the multiply charged ion abundance ratios using ion time of flight is described. An analytical treatment of the discharge action in producing various ion species has been carried out.

Field evaporation of solid metal electrodes has been proposed as an ion source for an electrostatic propulsion device. The chief advantage over existing ion sources is the prospect of 100 percent fuel utilization efficiency. This advantage arises as a result of the elimination of the need for a gaseous precursor state for propellant ionization. The attainment of required high surface field strengths is achieved through field-induced extrusion of the electrode geometry at elevated temperatures. Contributions of both surface and bulk transport mechanisms are taken into account.

tecnologie segrete

tecnologie segrete

tecnologie segrete

This effort examines the spectroscopy of the second spectrum of the iodine atom (I II) in order to determine one, or more, useful transitions for laser-induced fluorescence of an accelerated atomic iodine singly charged ion (I+). While the second spectrum of iodine has been analyzed, it is not particularly well characterized. Nor has it been studied substantially within a plasma such as those of interest to the spacecraft propulsion community. Our goal is to examine the spectral data available in the literature and determine transitions suitable for development into diagnostics tools, such as laser- induced fluorescence (LIF), to examine the plasma acceleration within an electrostatic plasma propulsion thruster. While xenon remains the preferred propellant for electrostatic spacecraft propulsion, a number of alternative propellants are being analyzed in various laboratories. Some of the propellants that have been investigated in the recent literature include krypton, bismuth, and iodine. Of these alternative propellant candidates, iodine is the least well investigated. However, due to its close mass (127 versus 131 amu) compared to xenon, it has strong potential for use as an electrostatic propulsion propellant. Iodine's benefis include a solid density of 4.9 g/cc, a low boiling point of 183 degrees C. Compared to xenon storage density of 1.2 g/cc at 2,000 psi, or the bismuth boiling point of 1,564 degrees C, there appear to be system level advantages to iodine fueled electrostatic spacecraft propulsion. This effort focuses on the development of a laser-induced fluorescence diagnostic tool for the iodine ion.

Nuclear electric generator for accelerating charged propellant particles in electrostatic propulsion system

The suitability of the baseline 30 cm thruster for future space missions was examined. Preliminary design concepts for several advanced thrusters were developed to assess the potential practical difficulties of a new design. Useful methodologies were produced for assessing both planetary and earth orbit missions. Payload performance as a function of propulsion system technology level and cost sensitivity to propulsion system technology level are among the topics assessed. A 50 cm diameter thruster designed to operate with a beam voltage of about 2400 V is suggested to satisfy most of the requirements of future space missions.

tecnologie segrete