摘要:

Radioisotope Electric Propulsion (REP) has the potential to enable small spacecraft to orbit outer planetary targets with trip times comparable to flyby missions. The ability to transition from a flyby to an orbiter mission lies in the availability of continuous low power electric propulsion along the entire trajectory. The electric propulsion system’s role is to add and remove energy from the spacecraft’s trajectory to bring it in and out of a heliocentric hyperbolic escape trajectory for the outermost target bodies. Energy is added and the trajectory is reshaped to rendezvous with the closer‐in target bodies. Sample REP trajectories will be presented for missions ranging for distances from Jupiter orbit to the Pluto‐Kuiper Belt. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649581

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

This paper describes research on combining a microelectromechanical system (MEMS) nuclear battery with a field‐emission electric propulsion (FEEP) thruster, thereby providing potentially attractive solutions to precise satellite stationkeeping and propulsion requirements. The MEMS nuclear battery, under development at the University of Wisconsin, consists of multiple layers of a radioisotope source alternating with pn junction semiconductor energy converters. Many radioisotopes were assessed for this purpose, typically with average beta‐particle energies of 50–250 eV, and the beta‐emitter Cs‐137 tentatively has been identified as most suitable. A slit‐style, cesium‐propellant FEEP thruster was chosen for the present study because it is a relatively mature technology. For use with a FEEP thruster, many modular MEMS nuclear batteries must be arrayed in series in order to achieve a sufficiently high voltage (∼10 kV). Critical issues include achieving an attractively high MEMS nuclear battery efficiency, maximizing the battery’s lifetime against radiation damage, producing the relatively high voltage (∼10 kV) required for a FEEP thruster, and providing an effective interface between the MEMS nuclear battery modules and the FEEP thruster. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649582

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

We have developed a suspension system using pre‐tensioned titanium alloy wires to support a 1 Watt Radioisotope Heating Unit (RHU) for a mission to Mars. This suspension is very strong in all directions and has quite low thermal conduction between the RHU and its colder surroundings. This will allow the RHU to operate at 250 °C for generating electrical power from an attached thermoelectric converter (TEC) after surviving multiple 300 G impacts on the Martian surface. We have tested the suspension under impact loads that have the same duration as the impacts expected from the air‐bag cushions planned for the Mars mission. The suspension survives impacts of 500 G with the RHU canister at room temperature and at 250 °C. We calculate that with the RHU at 250 °C there will be only 86 mW of heat conducted through the support structure. This design in intended for use with multilayer insulation in a good vacuum. We estimate that the conduction through such insulation would be 75 mW. This leaves more than 800 mW of the heat from the RHU to flow through the TEC, yielding more than 40 mW of electrical power. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649583

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

A shock‐tolerant thermal enclosure has been designed for use in distributed landed missions. Missions such as Pascal and the Mars Long‐Lived Landed Network require low power sources capable of surviving an omnidirectional load at impact and delivering reliable power for several Martian years. With the use of a radioisotope heat source and a thermoelectric converter, power can be generated reliably, but the challenge of developing an insulating canister that delivers sufficient power at end of life and is shock tolerant has been elusive. We describe a manufacturable design using conventional materials that meets mission requirements and show preliminary analysis of impact load response. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649584

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

The Titan Corporation conducted a systems engineering study to develop an overall architecture that meets both the articulated and unarticulated requirements on the Prometheus Program with the least development effort. Key elements of the Titan‐designed ENABLER system include a thermal fission reactor, thermionic power converters, sodium heat pipes, ion thruster engines, and a radiation shield and deployable truss to protect the payload. The overall design is scaleable over a wide range of power requirements from 10s of kilowatts to 10s of megawatts. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649585

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

The National Aeronautics and Space Administration’s Mars Exploration Rover (MER) 2003 project is designed to place two mobile laboratories (Rovers) on Mars to remotely characterize a diversity of rocks and soils. Milestones accomplished so far include two successful launches of identical spacecraft (the MER‐A and MER‐B missions) from Cape Canaveral Air Force Station, Florida on June 10 and July 7, 2003. Each Rover uses eight Light Weight Radioisotope Heater Units (LWRHUs) fueled with plutonium‐238 dioxide to provide local heating of Rover components. The LWRHUs are provided by the U.S. Department of Energy. In addition, small quantities of radioactive materials in sealed sources are used in scientific instrumentation on the Rover. Due to the radioactive nature of these materials and the potential for accidents, a formal Launch Approval Process requires the preparation of a Final Safety Analysis Report (FSAR) for submittal to and independent review by an Interagency Nuclear Safety Review Panel. This paper presents a summary of the FSAR in terms of potential accident scenarios, probabilities, source terms, radiological consequences, mission risks, and uncertainties in the reported results. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649586

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

On 10 June 2003 at 1:58 p.m. Eastern Daylight Time (EDT) and 7 July 2003 at 11:18 p.m. EDT, two separate spacecraft/rovers were successfully launched to Mars atop a Delta II 7925 and Delta II 7925H, respectively. Each spacecraft/rover carried eight Light Weight Radioisotope Heater Units (LWRHUs) for thermal conditioning of electronics during the cold Martian nights. As a part of the joint National Aeronautics and Space Administration/U. S. Department of Energy safety effort, a contingency plan was prepared to address the unlikely events of an accidental suborbital reentry or out‐of‐orbit reentry. The objective of the contingency plan was to develop and implement procedures to predict, within the first hour, the probable Earth Impact Footprints (EIFs) for the LWRHUs or other possible spacecraft debris after an accidental reentry. No ablation burn‐through of the heat sources’ aeroshells was expected, as a result of earlier testing. Any predictions would be used in subsequent notification and recovery efforts. The Johns Hopkins University Applied Physics Laboratory, as part of a multi‐agency team, was responsible for prediction of the EIFs, and the time of reentry from a potential orbital decay. The tools used to predict the EIFs included a Three‐Degree‐of‐Freedom (3DOF) trajectory simulation code, a Six‐Degree‐of‐Freedom (6DOF) code, a database of aerodynamic coefficients for the LWRHUs and other spacecraft debris, secure links to obtain tracking data, and a high fidelity special perturbation orbit integrator code to predict time of spacecraft reentry from orbital decay. This paper will discuss the contingency plan and process, as well as highlight the improvements made to the analytical tools. Improvements to the 3DOF, aerodynamic database, and orbit integrator and inclusion of the 6DOF have significantly enhanced the prediction capabilities. In the days before launch, the trajectory simulation codes were exercised and predictions of hypothetical EIFs were produced. The contingency efforts, while not exercised for the two successful launches, still contributed to mission safety and demonstrated cooperation among multiple agencies. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649587

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

The extent to which, if any, full power ground nuclear testing of space reactors should be performed has been a point of discussion within the industry for decades. Do the benefits outweigh the risks? Are there equivalent alternatives? Can a test facility be constructed (or modified) in a reasonable amount of time? Is the test article an accurate representation of the flight system? Are the costs too restrictive? The obvious benefits of full power ground nuclear testing; obtaining systems integrated reliability data on a full‐scale, complete end‐to‐end system; come at some programmatic risk. Safety related information is not obtained from a full‐power ground nuclear test. This paper will discuss and assess these and other technical considerations essential in the decision to conduct full power ground nuclear‐or alternative‐tests. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649588

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

A methodology for the neutronics design of space power reactors is presented. This methodology involves balancing the competing requirements of having sufficient excess reactivity for the desired lifetime, keeping the reactor subcritical at launch and during submersion accidents, and providing sufficient control over the lifetime of the reactor. These requirements are addressed by three reactivity values for a given reactor design: the excess reactivity at beginning of mission, the negative reactivity at shutdown, and the negative reactivity margin in submersion accidents. These reactivity values define the control worth and the safety worth in submersion accidents, used for evaluating the merit of a proposed reactor type and design. The Heat Pipe‐Segmented Thermoelectric Module Converters space reactor core design is evaluated and modified based on the proposed methodology. The final reactor core design has sufficient excess reactivity for 10 years of nominal operation at 1.82 MW of fission power and is subcritical at launch and in all water submersion accidents. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649589

出版商:AIP    

年代:1904

数据来源: AIP

摘要:

This paper presents a logical approach to designing a safety test plan for a space nuclear system. It is pointed out that two important facts need to underlie the development of a test plan: first, that sequential insults and the accumulation of damage are the rule; and second that the response of the nuclear system is stochastic (i.e., for any given set of conditions a probabilistic range of outcomes will occur regardless of the state of our knowledge). Because of these facts a deterministic approach can only be a starting point. The substance of the approach consists of undertaking and documenting three basic efforts: (1) a description of the analysts view of the problem and how it fits into the safety analysis, (2) a formal documentation of the purpose and requirements of the test plan (or test), and (3) an assessment of the use or usefulness of existing test data. © 2004 American Institute of Physics

ISSN:0094-243X    

DOI:10.1063/1.1649590

出版商:AIP    

年代:1904

数据来源: AIP