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In this paper we describe a new concept and prototype for dramtically reducing the cost of contact with planetary spacecraft. Known as spacecraft Demand Access, a suite of spacecraft and ground automation technologies, it enables future intelligent spacecraft to act as initiators of cost effective contact with the ground - doing it only when necessary.
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Scientific studies using spacecraft radio links have been conducted on almost every Solar System exploration mission in the past five decades and have led to numerous discoveries. Radio Science experiments have elucidated the thermal history of the Moon from high resolution gravitational field measurements, unveiled the interiors of Titan, Enceladus, Mercury, Phobos, Vesta and Ceres providing key evidence for identifying subsurface oceans on icy moons; sounded Titan, Saturn, and Pluto’s atmospheres, and refined models for the atmospheres, surfaces, and interior structure of Mars and Venus. A Juno experiment is in progress measuring the gravitational field of Jupiter to reveal its interior structures, as did a similar recent Cassini experiment with Saturn. Experiments at Mercury, the Jovian system, and other targets, are in development or planning phases. Over the next 30 years, significant advances in radio and laser link-science technologies, including nearly one order of magnitude improvement achievable in range-rate and range accuracy, could enable many new scientific breakthroughs. Future exploration concepts in many cases focus on applications of small spacecraft and can include spacecraft constellations for studies of atmospheric dynamics, interior structures, and surface properties. A set of science-enabling radio link technologies specific to small spacecraft instrumentation on future solar system missions are under study and development. Examples include field tests of radio scattering to determine soil properties, smallsat constellations for dense geographic and temporal atmospheric probing, small science-quality software-defined transponders, miniature ultra-stable oscillators, and advanced radio-metric calibrations at the Deep Space Network. This paper describes many of these technologies and their scientific applications.
Results are presented for a study on mission definition and design to determine broad technology directions and needs for advanced planetary spacecraft and future planetary missions. The discussion covers mission selection, system design, and technology assessment and review for a multicomponent spacecraft exploration facility provided with nuclear power propulsion. As an example, the Automated Planetary Space Station at Jupiter is examined as a generic concept which has the capability of conducting in-depth investigations of different aspects of the entire Jovian system. Mission planning is discussed relative to low-thrust trajectory control, automatic target identification and landing, roving vehicle operation, and automated sample analysis.
This paper considers one aspect of the complex radiation design considerations for planetary spacecraft, namely the approach used for solving the effects of radiation on materials problems. An overview of the approach developed at JPL for Voyager, and currently being used on Galileo, is treated briefly. Examples of the postulated Jovian charged particle levels are given. The types of computer analyses codes used, the mass shielding techniques that evolved and the recommended shielding techniques for future planetary spacecraft are treated.
The Center for Radiation Engineering and Science for Space Exploration (CRESSE) at Prairie View A&M University, Prairie View, Texas, USA, is establishing an integrated, multi-disciplinary research program on the scientific and engineering challenges faced by NASA and the international space community caused by space radiation. CRESSE focuses on space radiation research directly applicable to astronaut health and safety during future long term, deep space missions, including Martian, lunar, and other planetary body missions beyond low earth orbit. The research approach will consist of experimental and theoretical radiation modeling studies utilizing particle accelerator facilities including: 1. NASA Space Radiation Laboratory (NSRL) at Brookhaven National Laboratory; 2. Proton Synchrotron at Loma Linda University Medical Center; and 3. Los Alamos Neutron Science Center (LANSCE) at Los Alamos National Laboratory. Specifically, CRESSE investigators are designing, developing, and building experimental test beds that simulate the lunar and Martian radiation environments for experiments focused on risk assessment for astronauts and instrumentation. The testbeds have been designated the Bioastronautics Experimental Research Testbeds for Environmental Radiation Nostrum Investigations and Education (BERT and ERNIE). The designs of BERT and ERNIE will allow for a high degree of flexibility and adaptability to modify experimental configurations to simulate planetary surface environments, planetary habitats, and spacecraft interiors. In the nominal configuration, BERT and ERIE will consist of a set of experimental zones that will simulate the planetary atmosphere (Solid CO2 in the case of the Martian surface.), the planetary surface, and sub-surface regions. These experimental zones can be used for dosimetry, shielding, biological, and electronic effects radiation studies in support of space exploration missions. BERT and ERNIE are designed to be compatible with the experimental areas associated with the above facilities. CRESSE has broad expertise in space radiation in the areas of space radiation environment modeling, Monte-Carlo radiation transport modeling, space radiation instrumentation and dosimetry, radiation effects on electronics, and multi-functional composite shielding materials. The BERT and ERNIE testbeds will be utilized in individual and collaborative research incorporating this expertise. The research goal is to maximize the technical readiness level (TRL) of radiation instrumentation for human and robotic missions, optimizing the return value of CRESSE for NASA exploration and international co-operative missions. Outcomes and knowledge from research utilizing BERT and ERNIE will be applied to a variety of scientific and engineering disciplines vital for safe and reliable execution of future space exploration missions, which can be negatively impacted by the space radiation environment. The testbeds will be central to a variety of university educational activities and educational goals of NASA. Specifically, BERT and ERNIE will enhance educational opportunities in science, technology, engineering and mathematics (STEM) disciplines for engineering and science students at PVAMU, a historically black college/university. Preliminary data on prototype testbed configurations, including simulated lunar regolith (JSC-1A stimulant based on Apollo 11 samples), regolith/polyethylene composites, and dry ice, will be presented to demonstrate the usefulness of BERT and ERNIE in radiation beam line experiments.
Mass is the most important limiting parameter for present-day planetary spacecraft design, In fact, the entire design can be characterized by mass. The more efficient the design of the spacecraft, the less mass will be required. The communications system is an essential and integral part of planetary spacecraft. A study is presented of the mass attributable to the communications system for spacecraft designs used in recent missions in an attempt to help guide future design considerations and research and development efforts. The basic approach is to examine the spacecraft by subsystem and allocate a portion of each subsystem to telecommunications. Conceptually, this is to divide the spacecraft into two parts, telecommunications and nontelecommunications. In this way, it is clear what the mass attributable to the communications system is. The percentage of mass is calculated using the actual masses of the spacecraft parts, except in the case of CRAF. In that case, estimated masses are used since the spacecraft was not yet built. The results show that the portion of the spacecraft attributable to telecommunications is substantial. The mass fraction for Voyager, Galileo, and CRAF (Mariner Mark 2) is 34, 19, and 18 percent, respectively. The large reduction of telecommunications mass from Voyager to Galileo is mainly due to the use of a deployable antenna instead of the solid antenna on Voyager.
An experimental investigation of the vibration characteristics of a 60 deg conical shell model of a planetary entry vehicle is described and the results presented. Model configurations include the shell with or without one or two Z-ring stiffeners and with or without a simulated payload. Tests were conducted with the model clamped at the small diameter and with the model suspended at the simulated payload. Additionally, calculated results obtained from application of several analytical procedures reported in the literature are presented together with comparisons between experimental and calculated frequencies and meridional mode shapes. Generally, very good frequency agreement between experimental and calculated results was obtained for all model configurations. For small values of circumferential mode number, however, the frequency agreement decreased as the number of ring stiffeners increased. Overall agreement between experimental and calculated mode shapes was generally good. The calculated modes usually showed much larger curvatures in the vicinity of the rings than were observed in the experimentally measured mode shapes. Dual resonances associated with modal preference were noted for the shell without Z-ring stiffeners, whereas the addition of stiffeners produced resonances for which the model responded in two or more modes over different sections of the shell length.
Nuclear-electric propulsion evaluated for unmanned planetary and interplanetary spacecraft, comparing thermionic and turbogenerator power plants
Electric propulsion requirements for planetary and interplanetary spacecraft
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Electric propulsion requirements for planetary and interplanetary spacecraft
The Cassini project at the Jet Propulsion Laboratory (JPL) is implementing a spacecraft telemetry system based on the Consultative Committee for Space Data Systems (CCSDS) packet telemetry standards. Resolving the CCSDS concepts with a Ground Data System designed to handle time-division-multiplexed telemetry and also handling constraints unique to a deep-space planetary spacecraft (such as fixed downlink opportunities, small downlink rates and requirements for on-board data storage) have resulted in spacecraft and ground system design challenges. Solving these design challenges involved adapting and extending the CCSDS telemetry standards as well as changes to the spacecraft and ground system designs. The resulting spacecraft/ground system design is an example of how new ideas and philosophies can be incorporated into existing systems and design approaches without requiring significant rework. In addition, it shows that the CCSDS telemetry standards can be successfully applied to deep-space planetary spacecraft.
The Cassini project at the Jet Propulsion Laboratory (JPL) is implementing a spacecraft telemetry system based on the Consultative Committee for Space Data Systems (CCSDS) packet telemetry standards. Resolving the CCSDS concepts with a Ground Data System designed to handle time-division-multiplexed telemetry and also handling constraints unique to a deep-space planetary spacecraft (such as fixed downlink opportunities, small downlink rates and requirements for on-board data storage) have resulted in spacecraft and ground system design challenges. Solving these design challenges involved adapting and extending the CCSDS telemetry standards as well as changes to the spacecraft and ground system designs. The resulting spacecraft/ground system design is an example of how new ideas and philosophies can be incorporated into existing systems and design approaches without requiring significant rework. In addition, it shows that the CCSDS telemetry standards can be successfully applied to deep-space planetary spacecraft.
ORBSIM program developed for accurate extraction of parameters of geophysical models from Doppler-radio-tracking data acquired from orbiting planetary spacecraft. Model of proposed planetary structure used in numerical integration along simulated trajectories of spacecraft around primary body. Written in FORTRAN 77.
The work performed on the Ion Propulsion System Technology Task in FY90 is described. The objectives of this work fall under two broad categories. The first of these deals with issues associated with the application of xenon ion thrusters for primary propulsion of planetary spacecraft, and the second with the investigation of technologies which will facilitate the development of larger, higher power ion thrusters to support more advanced mission applications. Most of the effort was devoted to investigation of the critical issues associated with the use of ion thrusters for planetary spacecraft. These issues may be succinctly referred to as life time, system integration, and throttling. Chief among these is the engine life time. If the engines do not have sufficient life to perform the missions of interest, then the other issues become unimportant. Ion engine life time was investigated through two experimental programs: an investigation into the reduction of ion engine internal sputter erosion through the addition of small quantities of nitrogen, and a long duration cathode life test. In addition, a literature review and analysis of accelerator grid erosion were performed. The nitrogen addition tests indicated that the addition of between 0.5 and 1.0 percent of nitrogen by mass to the xenon propellant results in a reduction in the sputter erosion of discharge chamber components by a factor of between 20 and 50, with negligible reduction in thruster performance. The long duration test of a 6.35-mm dia. xenon hollow cathode is still in progress, and has accumulated more than 4,000 hours of operation at an emission current of 25 A at the time of this writing. One of the major system integration issues concerns possible interactions of the ion thruster produced charge exchange plasma with the spacecraft. A computer model originally developed to describe the behavior of mercury ion thruster charge exchange plasmas was resurrected and modified for xenon propellant. This model enables one to calculate the flow direction and local density of the charge exchange plasma, and indicates the degree to which this plasma can flow upstream of the thruster exhaust plane. A continuing effort to investigate the most desirable throttling technique for noble gas ion thrusters concentrated this year on experimentally determining the fixed flow rate throttling range of a 30-cm dia. thruster with a two-grid accelerator system. These experiments demonstrated a throttling capability which covers a 2.8 to 1 variation in input power. This throttling range is 55 percent greater than expected, and is due to better accelerator system performance at low net-to-total voltage ratios than indicated in the literature. To facilitate the development of large, higher power ion thrusters several brief studies were performed. These include the development of a technique which simulates ion thruster operation without beam extraction, the development of an optical technique to measure ion thruster grid distortion due to thermal expansion, tests of a capacitance measurement technique to quantify the accelerator system grid separation, and the development of a segmented thruster geometry which enables near term development of ion thrusters at power levels greater than 100 kW. Finally, a paper detailing the benefits of electric propulsion for the Space Exploration Initiative was written.
Design, alignment, testing, and specifications of S-band transponder Mark I with 20-cycle bandwidth for planetary spacecraft
Technology of systems analysis and contribution to design and flight of lunar and planetary spacecraft