Space Power Technological, Economic and Societal Issues in Space Systems Development Volume 8 Number 3 1989 —coifax— publi/hing company
SPACE POWER Published under the auspices of the SUNSAT Energy Council EDITOR Andrew Hall Cutler, Space Studies Institute Space Power is an international journal for the presentation, discussion and analysis of advanced concepts, initial treatments and ground-breaking basic research on the technical, economic and societal aspects of large-scale, space-based solar power, space resource utilization, space manufacturing, and other areas related to the development and use of space for the long-term benefit of humanity. Papers should be of general and lasting interest and should be written so as to make them accessible to technically educated professionals who may not have worked in the specific area discussed in the paper. Editorial and opinion pieces of approximately one journal page in length will occasionally be considered if they are well argued and pertinent to the content of the journal. Submissions should represent the original work of the authors and should not have appeared elsewhere in substantially the same form. Proposals for review papers are encouraged and will be considered by the Editor on an individual basis. Editorial Correspondence: Dr Andrew Hall Cutler can be reached by telephone at (619) 284-2779, and his address is 3030 Suncrest No. 214, San Diego, CA 92116, USA. Dr Cutler should be consulted to discuss the appropriateness of a given paper or topic for publication in the journal, or to submit papers to it. Questions and suggestions about editorial policy, scope and criteria should initially be directed to him, although they may be passed on to an Associate Editor. Details concerning the preparation and submission of manuscripts can be found on the inside back cover of each issue. Business correspondence, including orders and remittances to subscriptions, advertisements, back numbers and offprints, should be addressed to the publishers: Carfax Publishing Company, P.O. Box 25, Abingdon, Oxfordshire 0X14 3UE, United Kingdom. The journal is published in four issues which constitute one volume. An annual index and titlepage is bound in the December issue. ISSN 0951-5089 © 1989, SUNSAT Energy Council Cover: An artist's concept of the enhanced configuration of the permanently manned Space Station, produced by Rockwell International. The enhanced configuration includes an upper and lower keel for attaching external payloads, a 50 kilowatt solar dynamic system mounted on the ends of the transverse boom, a servicing bay and a co-orbiting platform (not pictured). Reproduced by Courtesy of NASA, Washington, DC, USA.
SPACE POWER Volume 8 Number 3 1989 IAF International Conference on Space Power, Cleveland, OH, USA, 5-7 June 1989 SPACE POWER MISSION APPLICATIONS I 9-1. Gary L. Bennett. Historical Overview of the US Use of Space Nuclear Power 259 9-4. Dana G. Andrews. Near-term Nuclear Space Missions 285 9-5. Joseph Appelbaum & Dennis J. Flood. The Mars Climate for a Photovoltaic System Operation 307 9-6. D. J. Bents. Preliminary Assessment of Rover Power Systems for the Mars Rover Sample Return Mission 319 9-8. Claude Poher. Candidate Space Missions for Nuclear Generators: Study Results and Implications 331 SPACE POWER BEAMING 10-2. William C. Brown. Status of Beamed Power Transmission Technology and Applications at 2.45 Gigahertz 339 10-4. E. J. Conway & R. J. De Young. Beamed Laser Power for Advanced Space Missions 357 10-5. K. Chang, J. C. McCleary & M. A. Pollock. Feasibility Study of 35 GHz Microwave Power Transmission in Space 365 NUCLEAR SAFETY 11 -4. Vladimir Georgevich, Frederick Best & Carl Erdman. Loss of Coolant Accident Mitigation for Liquid Metal Cooled Space Reactors 371 SPACE POWER MISSION APPLICATIONS II 12-2. Lyle M. Jenkins. Construction of Large Space Power Facilities 379 SPACE POWER ABSTRACTS 387
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9-1. Historical Overview of the US Use of Space Nuclear Power GARY L. BENNETT Summary Since 1961, the United States has successfully flown 35 space nuclear power sources on 20 space systems. These space systems have included the Apollo, Pioneer, Viking and Voyager spacecraft launched by the National Aeronautics and Space Administration and navigation and communications satellites launched by the Department of Defense. These power sources performed as planned and in many cases exceeded their power requirements and/or lifetimes. All of the power sources met their safety requirements. This paper surveys past uses of space nuclear power in the US and thus serves as a historical framework for other papers in this Conference dealing with future US applications of space nuclear power. Introduction The United States has used nuclear power on a number of technically sophisticated space systems which have greatly advanced our understanding of the solar system. In many cases, nuclear power was the only way to accomplish these missions. In the early 1950s, the US began studies of the use of nuclear power on spacecraft and by the late 1950s had active programs under way to develop both radioisotope and reactor power sources for spacecraft. The first actual use of a nuclear power source (NPS) on a spacecraft came in 1961 with the launch of the small SNAP-3B* radioisotope thermoelectric generator (RTG). In total, as shown in Table I, the US has launched 38 RTGs and one reactor to provide power for 23 space systems. (Thirty- five of these NPS on 20 space systems are still in space or on other planetary bodies.) The US has also used small radioisotope heater units (RHUs) on some of its RTG- powered science missions and on the Apollo 11 science package. All of the US RTGs have used 238Pu as the source of heat because of its long half-life (87.8 years) and its comparatively low level of radiation emission. The only US space reactor flown used 235U as the fuel [1,2]. Gary L. Bennett, National Aeronautics and Space Administration, Propulsion, Power and Energy Division, Washington, DC 20546, USA. Paper number IAF-ICOSP89-9-1. * SNAP is an acronym for Systems for Nuclear Auxiliary Power. All odd-numbered SNAP power sources used radioisotope fuel and all even-numbered SNAP power sources used nuclear fission reactors.
Table I. Summary of Space Nuclear Power Systems Launched by the United States Power Source1 Spacecraft Mission type Launch date Status SNAP-3B7 TRANSIT 4A Navigational 29 June 1961 RTG operated for 15 years. Satellite now shutdown but operational. SNAP-3B8 TRANSIT 4B Navigational 15 Nov. 1961 RTG operated for 9 years. Satellite operation was intermittent after 1962 high-altitude nuclear test. Last reported signal in 1971. SNAP-9A TRANSIT 5-BN-l Navigational 28 Sept. 1963 RTG operated as planned. Non-RTG electrical problems on satellite caused satellite to fail after 9 months. SNAP-9A TRANSIT 5-BN-2 Navigational 5 Dec. 1963 RTG operated for over 6 years. Satellite lost navigational capability after 1.5 years. SNAP-9A TRANSIT 5-BN-3 Navigational 21 April 1964 RTG operated as planned. Mission was aborted because of launch vehicle failure. SNAP-10A (Reactor) Snapshot Experimental 3 April 1965 Reactor operated successfully as planned. Satellite shutdown reactor after 43 days. SNAP-19B2 Nimbus-B-1 Meteorological 18 May 1968 RTGs operated as planned. Mission was aborted because of range safety destruct. RTGs recovered. SNAP-19B3 Nimbus III Meteorological 14 April 1969 RTGs operated for over 2.5 years (no data taken after that). SNAP-27 Apollo 12 Lunar 14 Nov. 1969 RTG operated for about 8 years (until station was shutdown). SNAP-27 Apollo 13 Lunar 11 April 1970 Mission aborted on way to moon. Heat source returned to South Pacific Ocean. SNAP-27 Apollo 14 Lunar 31 Jan. 1971 RTG operated for about 6.5 years (until station was shutdown). SNAP-27 Apollo 15 Lunar 26 July 1971 RTG operated for over 6 years (until station was shutdown). SNAP-19 Pioneer 10 Planetary 2 Mar. 1972 RTGs still operating. Spacecraft successfully operated to Jupiter and is now beyond orbit of Pluto. SNAP-27 Apollo 16 Lunar 16 April 1972 RTG operated for about 5.5 years (until station was shutdown). Transit-RTG ‘TRANSIT' (TRIAD-01-1X) Navigational 2 Sept. 1972 RTG still operating. SNAP-27 Apollo 17 Lunar 7 Dec. 1972 RTG operated for almost 5 years (until station was shutdown). SNAP-19 Pioneer 11 Plantery 5 April 1973 RTGs still operating. Spacecraft successfully operated to Jupiter, Saturn, and beyond. SNAP-19 Viking 1 Mars 20 Aug. 1975 RTGs operated for over 6 years (until lander was shutdown). SNAP-19 Viking 2 Mars 9 Sept. 1975 RTGs operated for over 4 years until relay link was lost. MHW-RTG LES 8 Communications 14 Mar. 1976 RTGs still operating. MHW-RTG LES 9 Communications 14 Mar. 1976 RTGs still operating. MHW-RTG Voyager 2 Planetary 20 Aug. 1977 RTGs still operating. Spacecraft successfully operated to Jupiter, Saturn, Uranus, and beyond. MHW-RTG Voyager 1 Planetary 5 Sept. 1977 RTGs still operating. Spacecraft successfully operated to Jupiter, Saturn, and beyond. 1 SNAP stands for Systems for Nuclear Auxiliary Power. All odd-numbered SNAP power plants use radioisotope fuel. Even-numbered SNAP power plants have nuclear fission reactors as a source of heat. MHW-RTG stands for the Multihundred Watt Radioisotope Thermoelectric Generator.
Initially these NPS were used to supplement solar power sources but gradually with the improvement of NPS technology and with the ever increasing requirements of spacecraft power (particularly for outer planet missions) NPS became the sole source of power. In a sense this was inevitable given the compact size, self-sufficiency, reliability, survivability, long lifetimes and operational flexibility of NPS. The basic NPS consists of a heat source (either a naturally decaying radioisotope or a nuclear reactor) and a converter (e.g. thermoelectric, thermionic, Brayton, Rankine, Stirling, magnetohydrodynamic) to change the thermal power into electrical power. To date the US has only used thermoelectric converters because of their proven reliability and the lack of a requirement to provide powers high enough to warrant the use of more efficient conversion systems such as turbine/alternators. The following sections provide an overview of the NPS flown by the US. This overview will serve to provide the framework for understanding the current programs under way in the US. Throughout the evolution of the US space NPS program there has been a general technology trend to improve NPS performance, efficiency, and specific power. This trend has led to improvements in the fuel and in the technology of thermoelectric materials, from the lead telluride (PbTe) used in the first five RTG concepts flown to the silicon germanium (SiGe) used in the SNAP-10A reactor and in the multi-hundred watt (MHW) RTGs and planned for use in future NPS. The performance of these NPS has clearly demonstrated that they can be safely and reliably engineered to meet a variety of space-mission requirements [2]. Radioisotope Power Sources The first SNAP, known as SNAP-1, was to use a radioisotope heat source coupled to a mercury Rankine cycle turbine/alternator. However, evolving requirements led the US toward the use of thermoelectrics such as were used on the SNAP-3B RTGs shown in Fig. 1. For this paper the various RTGs have been grouped into six basic design concepts: SNAP-3B, SNAP-9 A, SNAP-19, SNAP-27, TRANSIT-RTG and MHW- RTG. Since the focus of this paper is on providing a general historical overview the detailed power performance, which has been summarized in Bennett et al. [2], will not be repeated here.
SNAP-3B The SNAP-3B RTGs, which were developed out of the SNAP-3 program, were used to provide 2.7-We of power to radio transmitters and other electronic equipment aboard the US Navy's Transit 4A and Transit 4B navigation satellites. The SNAP-3B RTGs also were flown to prove the practicability of radioisotope power sources in space [2, 3], Prior to the use of NPS, continuous electrical power had been obtained by solar arrays and nickel-cadmium (NiCd) batteries. Concern over possible degradation of solar cells in the inner Van Allen belt and battery breakdown from repeated chargedischarge cycles had led the Navy to fly RTGs [3]. Each 2.1-kg SNAP-3B RTG contained 27 spring-loaded, series-connected pairs of PbTe thermoelectric elements operating at a hot junction temperature of about 783 K and a cold-junction temperature of about 366 K. Each radioisotope heat source provided about 52.5 Wt. The design life was 5 years. Figure 2 shows an assembled SNAP-3B and Fig. 3 shows the first mounting of NPS to a spacecraft in 1961. At the time Transit 4A, which is shown in Fig. 4, had the longest operating life of any satellite launched by the US—over 15 years. The RTG on Transit 4B was still operating 10 years after launch when the last signals were received [2-5]. SNAP-9A The SNAP-9A RTGs, as shown in Figs 5 and 6, were built to provide all of the electrical power for the Navy Transit 5BN navigation satellites. In fact, Transit 5BN-1, which was launched in 1963 and is shown in Fig. 7, was the first satellite to get all of its power from an RTG. The RTG approach was selected because RTGs are
inherently radiation resistant, whereas the solar-cell power system of Transit 4B had been adversely affected by a 1962 high-altitude nuclear explosion [6]. Each 12.3-kg SNAP-9A was designed to provide 25 We at a nominal 6 V for 5 years in space after 1 year of storage on Earth [7]. One of the objectives of the Transit 5BN program was to demonstrate the satisfactory operation and long-life potential of the SNAP-9A power supply. The Applied Physics Laboratory, which built the satellites, reported that the objective was fully satisfied. In fact, Transit ‘5BN-1 demonstrated the extreme simplicity with which thermoelectric generators may be integrated into the design, not only to provide the electrical power but also to aid in thermal control' [4]. Some waste heat from the RTG was used to maintain electronic instruments within the satellite at a temperature near 293 K. SNAP-19 The SNAP-19 technology-improvement program built on the SNAP-9A development program, with the SNAP-19B power source specifically designed for use on NASA's
Nimbus weather satellites. The Nimbus SNAP-19 program was the first demonstration of RTG technology aboard a NASA spacecraft, and, as such, it developed the data and experience to support interplanetary missions using RTGs. Subsequent modifications were made in the SNAP-19B design to power NASA's Pioneer and Viking missions. The Viking SNAP-19 is shown schematically in Fig. 8. For Nimbus III, two 13.4-kg SNAP-19B RTGs were mounted on the spacecraft platform as shown in Fig. 9 to provide a total of 56.4 We at beginning of mission (BOM) to augment the solar power source. During the design lifetime of one year, nuclear power comprised about 20% of the total power delivered to the regulated power bus, allowing a number of extremely important atmospheric-sounder experiments to operate in a full-time duty cycle. Without the RTGs the total delivered power would have fallen below the load line about 2 weeks into the mission [8,9]. Additional improvements were made leading to the SNAP-19s which were built for the Pioneer 10 and 11 spacecraft, the first to fly by Jupiter and Saturn. Figure 10 is an artist's rendition of a Pioneer spacecraft flying past Jupiter. The four RTGs on each
Pioneer spacecraft provided over 160 We at BOM. The Pioneer RTGs performed so well that Pioneer 11 was retargeted for a flyby of Saturn [10]. Both spacecraft are still operating 16-17 years after their launches, well beyond their 3-year design life requirement, and are providing valuable information about the heliosphere. Pioneer 10 is presently the most distant man-made object, having traveled beyond the orbit of Pluto, the outermost known planet [11]. The spacecraft should have sufficient power to provide useful data through at least 1996 [12]. The SNAP-19 design was further modified for the Viking Mars Landers to accommodate high-temperature (400 K.) sterlization, storage during the spacecraft's cruise to Mars, and, on the surface of Mars, the thermal cycling caused by the rapid and extreme temperature changes of the Martian day-night cycle. As shown in Fig. 11 each Viking Lander carried two of the 15.2-kg RTGs which produced a total power of over 85 We at BOM. The RTGs were to produce a total of 70 We for the primary mission of 90 days on the surface of Mars. All four RTGs met the 90-day requirement and they were still operating 4-6 years later when the Landers were separately and inadvertently shutdown on commands from Earth [13,14]. Based on their power performance, it had been estimated that the RTGs on Viking Lander 1 were capable of providing sufficient power for operation until 1994—18 years beyond the mission requirement [15]. Both the Pioneer and Viking RTGs demonstrated the operability and usefulness of RTGs in interplanetary spacecraft. All of these RTGs performed beyond their mission requirements.
SNAP-27 The SNAP-27 RTGs were developed to power the experiments of NASA's Apollo Lunar Surface Experiments Package (ALSEP). The RTG design requirement was to provide at least 63.5 We at 16 V DC 1 year after lunar emplacement. (In the case of Apollo 17, the requirement was 69 We 2 years after emplacement.) The use of RTGs to power the ALSEPs was a natural choice because of their light weight, reliability, and ability to produce full electrical power during the long lunar night-day cycle. Since the ALSEPs were to be manually positioned by the astronauts, the RTG designers took advantage of this assembly capability. The converter and the sealed-fuel-capsule assembly were kept separately in the Lunar Module and integrated on the Moon as shown in Fig. 12. This approach allowed optimization of the electrical, mechanical, and thermal interfaces of the two major hardware subsystems of the RTG [16], Figure 13 is a schematic of the SNAP-27 RTG. A total of five RTG-powered ALSEPs were placed on the Moon. In each case the RTGs exceeded their mission requirements in both power and lifetime (all were still operating when NASA shut down the stations on 30 September 1977). Through this performance beyond mission requirements, the SNAP-27 RTGs enabled the ALSEP stations to gather long-term scientific data on the internal structure and composition of the Moon, the composition of the lunar atmosphere, the state of the lunar interior, and the genesis of lunar features [17].
TRANSIT RTG The TRANSIT RTG was developed specifically as the primary power for the TRIAD navigational satellite, with auxiliary power to be provided by four solar-cell panels and one 6-Ah NiCd battery. The 13.6-kg TRANSIT RTG, shown in Fig. 14, was a modular RTG with a 12-sided converter surrounding the radioisotope heat source. The low hot side temperature (673 K) allowed operation of the PbTe thermoelectric elements in a vacuum [18]. To accomplish its mission of improving the accuracy of orbital determinations the TRIAD spacecraft was designed with three (‘triad') main units as shown in Fig. 15. These units are the power unit, the disturbance compensation system (DISCOS), and the main electronics unit. The TRANSIT RTG was the primary power source in the power unit. DISCOS, which was located at the satellite's center of mass, was designed to minimize the effects of aerodynamic drag forces and solar radiation pressure experienced in lower altitude orbits. DISCOS performed very successfully leading to the provision of excellent navigational capabilities to a wider variety of users. In addition, TRANSIT TRIAD provided very important measurements of the Earth's magnetic field. TRANSIT TRIAD operated for over 13 years—well beyond the design requirement of 5 years. Multihundred Watt (MHW) RTG The designs of the Lincoln Experimental Satellites 8 and 9 (LES 8/9) and NASA's
Voyager 1 and 2 spacecraft led to a doubling of the power requirement compared to the SNAP-27 RTGs. The MHW-RTG, which is illustrated in Fig. 16, was designed to produce over 150 We at BOM. Two MHW-RTGs were flown on each LES as shown in Fig. 17 and three MHW-RTGs were flown on each Voyager as shown in Fig. 18. Originally, Voyagers 1 and 2 were to fly past Jupiter and Saturn. The MHW-RTGs were the first US space RTGs to use SiGe as the thermoelectric material (see Fig. 19). The use of SiGe permitted higher operating temperatures and higher specific powers all within a space vacuum operating environment [19]. The MHW-RTGs on LES 8/9 continue to operate beyond the prelaunch required 5-year operational life. Similarly, the MHW-RTGs on Voyagers 1/2 continue to operate well beyond the prelaunch required 4-year operational life. Because of the outstanding performance of the Voyager RTGs NASA was able to extend the Voyager mission to include flybys of Uranus and Neptune [20]. The RTGs are performing so well that scientific data will be received into the early 21st century [12], The successful performance of the MHW-RTGs has led to the use of the SiGe technology for the high-power (285 We) general-purpose heat source RTG (GPHS- RTG), shown in Fig. 20, which is to provide power for NASA's Galileo spacecraft and the European Space Agency's Ulysses spacecraft [21]. Table II illustrates the trends in RTG technology from SNAP-3B to GPHS-RTG, showing the overall steady progress to date [2]. Reactor Power Sources By the early 1950s it was apparent that nuclear reactors offered the potential to power
some of the space satellite concepts then being considered. By the mid-1950s the US had developed the basic design of a compact space reactor with hydrided zirconiumuranium alloy fuel elements coupled with liquid metal coolant for efficient heat transfer. The SNAP-2, SNAP-8 and SNAP-10A reactor power sources were developed from this basic design [1, 22, 23, 24]. Table III lists the major US space reactor programs, including both power and propulsion [23,24]. SNAP-10A, which was the first and so far only space reactor flown by the US, evolved from the SNAP-2 sodium-potassium (NaK)-cooled Rankine converter reactor and the SNAP-10 conduction-cooled thermoelectric converter reactor. In 1960, the US Air Force (USAF) and the US Atomic Energy Commission (AEC) initiated the Space System Abbreviated Development Plan for Nuclear Auxiliary Power Orbital Test (SNAPSHOT) Program. Under the program, the USAF was to furnish the launch and satellite vehicles and the AEC was to furnish the SNAP-10A reactor units. The reactor was to provide not less than 500 We with a 1-year operating lifetime [22]. Included among the objectives of the SNAP-10A/SNAPSHOT program were to: • Demonstrate, proof test, and flight qualify SNAP-10A for subsequent operational use; • Demonstrate the adequacy and safety of ground handling and launch procedures; and • Demonstrate the adequacy and safety of automatic reactor startup in orbit. As shown in Fig. 21, the completed SNAP-10A system had the shape of a truncated cone with an overall length of 3.48 m and a mounting base diameter of 1.27 m. This
configuration was dictated by minimum mass shield requirements, especially the requirement to eliminate neutron scattering around the steel-reinforced lithium hydride shadow shield. The base diameter was established by the Agena vehicle payload and the upper diameter was determined by the effective area of the reactor. The length was determined by the total radiator area requirement. The total system mass of the final flight unit (known as FS-4) was 435 kg including the shield [22]. The reactor is shown in Fig. 22. The power conversion system basically consisted of 2880 SiGe thermoelectric elements mounted in groups of 72 along 40 stainless steel tubes through which the NaK coolant flowed. Figure 23 shows the overall thermodynamic cycle including a thermoelectric module. Despite its lower figure of merit at the SNAP-10A operating temperatures SiGe was chosen over PbTe because of (1) its stability to higher temperatures; (2) its potential for future performance growth; (3) its ease of manufacture; and (4) its mechanical properties. The converter hot side operating temperature was about 765 K and the mean radiator temperature was about 590 K [22]. On 3 April 1965, SNAP-10A was placed into a 1288 km by 1307 km orbit by an Atlas/Agena launch vehicle. Once it was confirmed that SNAP-10A was in a very long-lived orbit, the AEC authorized startup of the reactor [22]. Figure 24 is an artist's concept of SNAP-10A in space with the Agena. The automatic startup of SNAP-10A was accomplished flawlessly. The response of the FS-4 flight system was in excellent agreement with predictions based on analog computer studies and ground test results obtained from the FS-3 reactor. Net power
Table IL Trends in RTG Technology Parameter SNAP-3B SNAP-9A SNAP-27 TRANSIT-RTG SNAP-19 MHW-RTG GPHS-RTG Mission TRANSIT 4 TRANSIT 5BN Apollo Triad Pioneer Voyager Galileo/Ulysses BOM Power per RTG (We) 2.7 26.8 73.4 35.6 40.3 158.0 300“ Thermoelectric material PbTe 2N/2P PbTe 2N/2P PbTe 3N/3P PbTe 2N/3P PbTe 2N/TAGS-85 SiGe SiGe Pu-238 Fuel form Metal Metal Oxide (microspheres) PMC* PMC* Pressed oxide Pressed oxide Conversion efficiency (%) 5.1 5.1 5.0 4.2 6.2 6.6 6.8 Specific power (We/kg) 1.29 2.2 2.3“ 2.6 3.0 4.2 5.3 “Approximate power for originally planned 1986 launches—see [21]. *Plutonia Molybdenum Cermet. “The SNAP-27 specific power is shown with the fuel-cask mass included.
Table III. Principal US Space Nuclear Reactor Programs Power plant Purpose Power level Operating temp. (K) Period Type reactor Fuel Converter Development level Rover (includes NERVA) Propulsion 365-500 MWt 2450 1955-1973 Epithermal UC Twenty reactors tested. Demonstrated all components of flight engine >2 h. Ready for flight engine development. Fluidized bed reactor Propulsion 1000 MWt 3000 1958-1973 Thermal UC-ZrC Cold flow, bed dynamics experiments successful. Gaseous core reactors Propulsion and electricity 4600 MWt 10,000 1500 1959-1978 Fast Uranium plasma UF6 Brayton Successful critical assembly of UF6. SNAP-2 Electricity 3 kWe 920 1957-1963 Thermal Uranium zirconium hydride Mercury Rankine Development level. Tested two reactors with longest test reactor operated 10,500 h. Precursor for SNAP-8 and — 10A. SNAP-10A Electricity 0.5 kWe 810 1960-1966 Thermal Uranium zirconium hydride Thermoelectric Flight tested reactor 43 days. Tested reactor with thermoelectrics in 417-day ground test. SNAP-8 Electricity 30-60 kWe 975 1960-1970 Thermal Uranium zirconium hydride Mercury Rankine Tested two reactors. Demonstrated 1 year operation. Non-nuclear components operated 10,000 h and breadboard 8700 h. Advanced hydride reactors Electricity 5 kWe 920 1970-1973 Thermal Uranium zirconium hydride Thermoelectric and Brayton PbTe thermoelectrics tested to 42,000 h. SNAP-50 Electricity 300-1200 kWe 1365 1962-1965 Fast UN, UC Potassium Rankine Fuels tested to 6000 h. Advanced metal- cooled reactor Electricity 300 kWe 1480 1965-1973 Fast UN Brayton and potassium Rankine Non-nuclear potassium Rankine cycle components demonstrated to 10,000 h. Ready for breadboard loop. 710 gas reactor Electricity and propulsion 200 kWe 1445 1962-1968 Fast uo2 Brayton Fuel element tested to 7000 h. In-core thermionic reactor Electricity 5-250 kWe 2000 1959-1973 Fast or thermal driver uo2 UC-ZrC In-core thermionics Integral fuel element, thermionic diode demonstrated > 1 year operation. Nuclear electric propulsion Electricity 400 kWe 1675 1974-1981 Fast UO2 Out-of-core thermionics Limited testing on thermionic elements. SPAR/SP-100 Electricity 100 kWe 1500 1979-present Fast UN Thermoelectric Limited testing on core heat pipes and advanced thermoelectric materials.
output ranged from a transient high of 650 We in the early part of the mission to a low of 527 We in the Sun after 43 days. The overall system efficiency was about 1.3%. In general, the system operated exactly as intended [22], On 16 May 1965, after 43 days of successful operation, the reactor was shut down by a spurious command caused by a failure of a voltage regulator on the Agena unregulated bus. There was no evidence of any malfunction in the SNAP-10A system. The FS-3 ground test twin to FS-4 successfully operated at full power for 10000 hours
thereby demonstrating the capability of SNAP-10A to operate unattended for a year [22], The SNAP-10A reactor successfully completed most of its objectives, including the following significant achievements [22]: • First application of a nuclear reactor in space; • First development of a reactor thermoelectric power system and the first use of such a system in space;
• First remote automatic startup of a nuclear reactor in space; • First application of a high-temperature (810 K) liquid metal transfer system in space and the first application of a high-temperature spacecraft in space; • First use of a nuclear shadow shield in space; • Development and application of the highest powered thermoelectric power system to that time and the first use of a thermoelectric power system of that size in space; and • First thermoelectric powered liquid metal pump and the first use of such a pump in space. Space Nuclear Safety From the beginning, the US space nuclear power program has placed great emphasis on the safety of people and the protection of the environment. For RTGs, the safety philosophy is to contain or immobilize the radioisotope fuel to the maximum extent possible during all mission phases and postulated credible accidents. In the case of space nuclear reactor power systems, the current safety philosophy includes the launch of a nonoperating system so there is no buildup of radioactive fission products [25], The earlier NPS through SNAP-9A were designed to contain the fuel if the mission were aborted on the launch pad or during early ascent but to permit complete burnup of the fuel in the stratosphere. Worldwide dispersion and dilution of fine nuclear fuel particles would preclude local contamination. Transit 5BN-3, with a SNAP-9A power source, was launched on 21 April 1964 but failed to achieve orbit because of computer problems that affected the operation of the launch vehicle. The
satellite reentered the atmosphere over the ocean east of Africa. The RTG burned up on reentry, as it was designed to do. The burnup of SNAP-9A added about 4% to the total amount of plutonium in the environment. Subsequent studies by Italy, Japan, the UK, and the US have shown no measurable health effects from this reentry [25-30]. All US RTGs following SNAP-9A were designed to contain or immobilize the fuel through all credible accident conditions, including reentry and impact on Earth. The implementation of the new reentry philosophy was verified in two subsequent reentries: • Abort of the launch of the Nimbus-Bl satellite on 18 May 1968 by the range safety officer because of a guidance error. The two SNAP-19B RTGs were recovered intact as designed. • Damage of the Apollo 13 spacecraft from an oxygen tank explosion after a successful launch on 11 April 1970 leading to the intact reentry (as designed) of the SNAP-27 fuel cask over the South Pacific Ocean on 17 April 1970. The US Government employs an independent, formal multi-agency safety and environmental review of all NPS designs before the first launch. This process is illustrated in Fig. 25. The overall US approach is consistent with a UN working group report [31, 32]. In fact, the US has been an active participant in UN discussions on the safe use of NPS in outer space [33]. The US has supported the conclusion reached by the UN technical experts: ‘The Working Group reaffirmed its previous conclusion that NPS can be used safely in outer space, provided that all necessary safety requirements are met' [31]. Conclusion Space nuclear power sources have proved to be reliable, long-lived sources of electrical power that have enabled the conduct of a number of important US space missions, including the first long-term study of the surfaces of the Moon and Mars and the first exploratory visits to Jupiter, Saturn, and Uranus. In general, the NPS, from SNAP-3B to the MHW-RTG, met or exceeded their design requirements by providing power at or above that required and beyond the planned lifetime. All of the power sources met their safety requirements. This successful performance has laid a secure foundation for future US missions that will use nuclear power. ACKNOWLEDGEMENTS The author acknowledges with thanks the contributions made by members of the staffs of Teledyne Energy Systems, Rockwell International, General Electric Company, TRW Space and Defense, Fairchild Space Company, NUS Corporation, Applied Physics Laboratory, Battelle Columbus Laboratories, 3M Company, Sandia National Laboratories, Los Alamos National Laboratory, Savannah River Plant and Laboratory, the Mound Plant, and Oak Ridge National Laboratory. In particular, the author would like to thank John Dassoulas, Paul J. Dick, James C. Hagan, Richard B. Harty, C. E. Kelly, Frank D. Postula, E. A. Skrabek, and C. W. Whitmore for supplying information over the years.
REFERENCES [1] Bennett, G.L., Lombardo, J.J. & Rock, B.J. (1981) Development and use of nuclear power sources for space applications, The Journal of the Astronautical Sciences, XXIX, pp. 321-342. [2] Bennett, G.L., Lombardo, J.J. & Rock, B.J. (1983) US radioisotope thermoelectric generator space operating experience, Paper 839171, in: Proceedings of the 18th Intersociety Energy Conversion Engineering Conference, Orlando, Florida, 21-26 August. [Reprinted (1984) as: US radioisotope thermoelectric generators in space, The Nuclear Engineer, 25 (2), pp. 49-58.] [3] Dick, P.J. & Davis, R.E. (1962) Radioisotope power system operation in the transit satellite, presentation at the American Institute of Electrical Engineers Summer General Meeting, Paper no. CP 62-1173, Denver, Colorado, 17-22 June. [4] Johns Hopkins University Applied Physics Laboratory (1980) Artificial Earth Satellites Designed and Fabricated by the Johns Hopkins University Applied Physics Laboratory, JHU/APL Report SDO-1600 [revised] (August). [5] Harvey, D.G., Dick, P.J. & Fink, C.R. (1963) Isotope-generator reliability and safety, Nucleonics, 21, No. 4. [6] Hittman Associates, Inc. (1963) Radioisotope powered space systems, US Atomic Energy Commission report NYO-3165-11, 23 August. [Additional discussion of survivability issues may be found in: Bennett, G.L. (1988) Survivability considerations in the design of space power systems. Proceedings of the 23rd Intersociety Energy Conversion Engineering Conference, Paper 889159, Denver, Colorado, 31 July-5 August. [7] Marietta, M. (1964) SNAP-9A Pu-238 fueled thermoelectric power supply for auxiliary space power, Final report, Volume 1, Martin Nuclear Division report MND-P-3098-1 (October). [8] Fihelly, A.W. & Baxter, C.F. (1971) Orbital performance of the SNAP-19 radioisotopic thermoelectric generator experiment, Proceedings of the 6th Intersociety Energy Conversion Engineering Conference, Paper 719152, Boston, MA, 3-5 August. [9] Jaffe, H. & O'Riordan, P. (1972) Isotope power systems for unmanned spacecraft applications, Proceedings of the 7th Intersociety Energy Conversion Engineering Conference, Paper 729088, San Diego, CA, 25-29 September. [10] Goebel, C.J. (1975) SNAP-19 Pioneer 10 and 11 RTG deep space performance, Record of the 10th Intersociety Energy Conversion Engineering Conference, Paper 759130, Newark, Delaware, 18-22 August. [11] Skrabek, E.A. & McGrew, J.W. (1988) Pioneer 10 and 11 RTG performance update, in: M. S. El-Genk & M. D. Hoover (Eds) Space Nuclear Power Systems 1987 (Malabar, FL, Orbit Book Co.). [12] Cesarone, R.J., Sergeyevsky, A.B. & Kerridge, S.J. (1984) Prospects for the Voyager extra-planetary and interstellar mission. Journal of the British Interplanetary Society, 37, No. 3. [13] Brittain, W.M. (1976) SNAP-19 Viking RTG mission performance, Proceedings of the 11th Intersociety Energy Conversion Engineering Conference, Paper 769255, State Line, Nevada, 12-17 September. [14] Brittain, W.M. & Skrabek, E.A. (1983) SNAP-19 RTG performance update for the Pioneer and Viking missions, Proceedings of the 18th Intersociety Energy Conversion Engineering Conference, Paper 839172, Orlando, FL, 21-26 August. [15] Britting, A.O. Jr (1981) Viking lander battery performance, degradation, and
reconditioning, Proceedings of the 16th Intersociety Energy Conversion Engineering Conference, Paper 819110, Atlanta, GA, 9-14 August. [16] Pitrolo, A.A., Rock, B.J., Remini, W.C. & Leonard, J.A. (1969) SNAP-27 program review, Proceedings of the 4th Intersociety Energy Conversion Engineering Conference, Paper 699023, Washington, DC, 22-26 September. [17] Bates, J.R., Lauderdale, W.W. & Kernaghan, H. (1979) ALSEP termination report, NASA Reference Publication 1036 (April). [18] Bradshaw, G.B. & Postula, F.D. (1973) Beginning of mission flight data on the TRANSIT RTG, Proceedings of the 8th Intersociety Energy Conversion Engineering Conference, Paper 739091, Philadelphia, PA, 13-16 August. [Additional information on the TRIAD spacecraft may be found in: Dassoulas, J. (1973) The TRIAD spacecraft, APL Technical Digest, 12, No. 2.] [19] Kelly, C.E. The MHW converter (RTG), Record of the 10th Intersociety Energy Conversion Engineering Conference, Paper 759132, Newark, Delaware, 18-22 August. [20] Bennett, G.L. (1985) Voyager 2: Encounter with Uranus, Astronomy, 13, No. 9. [21] Bennett, G.L., Lombardo, J.J., Hemler, R.J. & Peterson, J.R. (1986) The general-purpose heat source radioisotope thermoelectric generator: power for the Galileo and Ulysses missions, Proceedings of the 21st Intersociety Energy Conversion Engineering Conference, Paper 869458, San Diego, CA, 25-29 August. [22] Staub, D.W. (1967) SNAP 10A summary report, Atomics International Report NAA-SR-12073 (25 March). [23] Bennett, G.L. & Buden, D. (1983) Use of nuclear reactors in space, The Nuclear Engineer, 24, No. 4. [24] Bennett, G.L. (1989) On the application of nuclear fission to space power, Proceedings of Fifty Years with Nuclear Fission Conference, Washington, DC and Gaithersburg, MD, 26-28 April. [25] Bennett, G.L. (1981) Overview of the US flight safety process for space nuclear power, Nuclear Safety, 22, No. 4. [26] DeBortoli, M.C. & Gaglione, P. (1969) SNAP Plutonium-238 fallout at Ispra, Italy, Health Physics, 16, pp. 197-204. [27] United Nations Scientific Committee on the Effects of Atomic Radiation (1982) Ionizing Radiation: Sources and Biological Effects: UNSCEAR 1982 Report to the General Assembly (New York, United Nations). [28] Japan (1979) Studies on technical aspects and safety measures of nuclear power sources in outer space, submitted to: Working Group on the Use of Nuclear Power Sources in Outer Space of the Scientific and Technical Subcommittee of the UN Committee on the Peaceful Uses of Outer Space, UN Document A/AC.105/C.1/WG.V/L.5 (28 December). [29] United Kingdom (1980) Studies on technical aspects and safety measures of nuclear power sources in outer space, submitted to: Working Group on the Use of Nuclear Power Sources in Outer Space of the Scientific and Technical Subcommittee of the UN Committee on the Peaceful Uses of Outer Space, UN Document A/AC.105/C.1/WG.V/L.11 (28 January). [30] Bartram, B.W. & Englehart, R.W. (1988) A pre-boost risk assessment of Cosmos 1900, NUS Corporation Report NUS-5148 (17 October). [31] Working Group on the Use of Nuclear Power Sources in Outer Space, Work of third session, Annex II Report of the Scientific and Technical Subcommittee [of the UN Committee on the Peaceful Uses of Outer Space], 18th Session, UN Document A/AC. 105/287 (13 February).
[32] Bennett, G.L. (1987) Flight safety review process for space nuclear power sources, Proceedings of the 22nd Intersociety Energy Conversion Engineering Conference, Paper 879046, Philadelphia, PA, 10-14 August. [33] Bennett, G.L., Sholtis, J.A. & Rashkow, B.C. (1988) United Nations deliberations on the use of nuclear power sources in space: 1978-1987, in: M. S. El- Genk & M. D. Hoover (Eds) Space Nuclear Power Systems 1988 (Malabar, FL, Orbit Book Co) [in press].
9-4. Near-term Nuclear Space Missions DANA G. ANDREWS Summary This paper summarizes the current status of proposed NASA Human Exploration Missions (HEMs), presenting detail with respect to schedule and system features. Using the Lunar exploration and Mars exploration missions as design drivers, new vehicle elements have been configured and optimized to meet proposed mission requirements. Both conventional chemical propulsion and alternative vehicle designs incorporating advanced nuclear propulsion systems are included, and the total mass required in Low Earth Orbit (LEO) is compared for each propulsion system. 1 Introduction Successful completion of the Civil Space Leadership Initiatives, and in particular, the Human Exploration of the Solar System as currently being defined by the Office of Exploration (OEXP) will require system capabilities to evolve beyond current capabilities, and humans to learn to live and work in space for periods measured in years instead of days. This will require new and innovative designs for spacecraft and surface elements, and thoughtful approaches to the problems of long-term space power. Space-based nuclear reactors are prime candidates to provide this power because of their low specific mass at the multimegawatt power levels and their ability to operate in all orbits and on planetary surfaces. Accordingly, we have concentrated on nuclear reactors as alternative propulsion and power elements in this paper. The overall objective of this paper is to summarize the human exploration case studies, define reference spacecraft systems and complementary advanced technology alternatives, performance-size the various elements to meet mission requirements, and determine what mix of vehicle types and technology developments results in the most cost effective system capable of completing the candidate mission cases. 2 Mission Characteristics Definition Baseline mission scenarios were established using the case studies in OEXP's Exploration Studies Technical Report, Scenarios Requirements Document, and the FY'89 Focused Case Studies White Paper [1-3] as guidelines. A key issue brought forth in [3], with respect to the baseline scenarios, and addressed in this paper, is the need to assemble very massive spacecraft in LEO using Block I technologies, or alternativey, the need to develop Nuclear Thermal Rockets (NTRs) and Nuclear Electric Propulsion (NEP) as Block II Technologies. As we shall show later, development of NEP reduces the need for the massive chemical stages assembled in LEO, making it an enabling technology for several of the case studies and cost effective for the others. Dana G. Andrews, Boeing Aerospace, M/S 8K-53, PO Box 3999, Seattle, Washington 98124, USA. Paper number IAF-ICOSP89-9-4.
Lunar Evolution Mission The lunar evolution strategy places emphasis on development of an early human- tended outpost which is expanded to a self-sufficient facility. Key sizing groundrules include initial crew size of 4 growing to 8, maximum reusability of system elements, aerobraking at Earth return, and baseline chemical propulsion for all stages. Key costing groundrules include in-situ propellant production as a Block I reference, and limiting the operational phase of the exploration and evolution of lunar facilities so it can be accomplished for constant annual investment. This means the life-cycle development of major components (both spacecraft and surface systems) will be phased with their first use date to minimize yearly funding profiles. The lunar facility evolution is carried out through a series of unmanned cargo flights and manned sorties. The first series of missions establishes the human-tended basecamp and early science outpost. This facility would be assembled in advance of the manned landing using teleoperated equipment or must be capable of rapid deployment/assembly by a small unassisted crew. A combination of the two assembly strategies is most likely. The initial basecamp should be capable of housing a crew of at least four, who will remain at the base for as long as the system will allow, given early lunar lander delivery capability, with an initial goal of two plus lunar days (60 Earth days). Following inital manned operations and systems checkout, an oxygen production facility will be delivered, assembled, checked out, and operations initiated by a second crew of four. Early emphasis during this stage will be on deploying science instrumentation in the vicinity of the base, on gaining operating experience with the production facility, and on demonstrating use of in situ materials to support closure of the Life Support System (LSS). When operational, the oxygen facility should produce enough LOX to support one lunar lander round-trip to orbit per year (roughly 2 tons/month). Eventually, replication of facility elements should bring this capability up to 10 tons of LOX per month to support a permanent lunar base. At the end of this human-tended phase, the first crew of 4-6 will be delivered to begin permanent occupancy with six month tours of duty. The objective during the early permanently manned phase is to establish and test the systems required to support additional crew for extended durations. This involves expansion of habitat and laboratory facilities, plus increases in the degree of closure of the LSS using in situ materials to decrease the dependance on earth resupply. During this period, the permanent crew will increase to 8 or 12 serving 1 year tours of duty, and the lunar oxygen facility will be increased in capability to provide LOX for all lunar lander traffic (10 tons/month). During lulls between resupply/construction missions, substantial scientific exploration and establishment of a lunar observatory will be undertaken. At the end of this phase, the capability to support a crew of 30 will exist, and the first crew to serve a 2-year tour of duty will arrive. The lunar base would now be fully operational and be used to exploit lunar resources for all mankind. First, the production of in situ propellants would be enhanced to include liquid hydrogen (LH2) in quantities necessary to fuel all lunar based transportation systems. Then, the quantity of LOX produced could be increased to allow LOX to be ‘exported' to LEO in aerobraked tankers, allowing more useful materials to be launched into LEO and trans-shipped to low lunar orbit (LLO). During this phase, advanced propulsion systems such as NTRs and NEPs could provide very large benefits in reducing the cost of LEO to LLO transportation, allowing more rapid expansion of lunar capabilities for the same system operating
costs. As a byproduct of the intensive mining effort, large quantities of iron and titanium would be produced which could be used to enlarge pressurized living and working volumes. Finally, serious mining of the lunar regolith to produce 3He could be started in an effort to make the moon economically independent. If demand for clean fusion power materialized on Earth then the crew of the lunar base could expand a thousandfold in order to operate and maintain thousands of automated 3He mining machines. Mars Evolution Mission Definition The Mars evolutionary strategy grows from an initial exploration of Mars to establishment of propellant facilities and ultimately a permanent manned base on Mars. Key sizing groundrules include: initial crew size of 4 growing to 8, reusability of selected vehicle elements, aerobraking at Mars and at Earth, baseline chemical propulsion for all stages, and accomplishment of program objectives with constant annual investment during the operational phase. The evolutionary program is carried out through a series of unmanned cargo flights and manned exploration missions described in Craig [3] and summarized briefly here. FY89 NASA strategy also imposed an annual limit on the allowable mass to LEO in support of the Mars evolution mission. The design groundrule was to use Block I technologies (chemical propulsion) until the exploration objectives, especially schedules, could no longer be met within the yearly mass-to-LEO constraint. At that point, advanced propulsion technologies would be introduced as a Block II update. Evolutionary strategy determines the mission characteristics for the Mars case. The initial manned mission would be to explore the Martian moons with no manned landing on Mars. This would be followed by development of a Martian moon gateway on Phobos which functions as a transportation node, a propellant source, and a propellant depot to reduce the cost of establishing an outpost and eventually a base on Mars. Both manned and unmanned missions are planned, with the cargo often ‘sent ahead' using minimum fuel Hohmann type trajectories and the crew using faster opposition class 600+ day trajectories. The first manned mission to Mars will aerobrake in the Martian atmosphere and then rendezvous with Phobos. The Mars Piloted Vehicle (MPV) can remain at Phobos and explore Diemos either robotically or with a small chemical Orbit Transfer Vehicle (OTV). Use of this exploration option allows the reusable MPV to be sized for its primary mission, which is to ferry humans to the Mars gateway station at Phobos. If it must rendezvous with both Phobos and Diemos in its maiden flight, the MPV design will be compromised with oversized tankage and heat shield, and therefore be nonoptimum for later flights. The second Mars mission departs earth in September 2007 in the refurbished MPV with a cargo of a Mars Crew Sortie Vehicle (MCSV) and a Mars Cargo Lander (MCL) containing a surface habitat and 1 year's supplies. All three vehicle aerobrake in the Martian atmosphere in August 2008 with the MCL landing on Mars while the MPV and the MCSV perform rendezvous with Phobos. The crew secures the MPV at Phobos, transfers into the MCSV, and descends to the Martian surface. They assemble the habitat facility and spend 1 year collecting extensive geophysical and environmental data and performing selected science experiments. In July 2009, the crew secures the habitat facility and ascends from Mars using the MCSV. They rendezvous with Phobos and transfer to the MPV for the trip back to earth. In May 2010, the MPV aerobrakes at Earth and rendezvous with the space station.
The third Mars mission is a one-way cargo flight which departs Earth in November 2009 for Phobos. The cargo is an automated propellant production plant designed to operate on Phobos and a ‘dry' MCL loaded with 45 tons of equipment to upgrade the Mars outpost to human-tended base status. The flight arrives in August 2010, the vehicles aerobrake separately in the Martian atmosphere and rendezvous separately with Phobos. The propellant plant begins telerobotic operation to produce water and then propellants from the carbonaceous chondrite material of Phobos, so that the Mars gateway station will take on operational status the following year. The MCL ‘docks' with Phobos and awaits the arrival of the next manned mission. The fourth mission departs Earth in November 2011 using the same refurbished MPV and arrives at Mars in September 2012. The MPV aerobrakes at Mars and proceeds to Phobos where the crew controls fueling of the MCSV, the MCL, and the MPV. After propellant transfer the MCL descends to the vicinity of the Mars Outpost and is followed shortly by the MSCSV containing the entire crew. During the year long stay on the surface the outpost is upgraded to base status with the capability to grow some food and utilize in situ resources to increase stored supplies of air and water. In August 2013, the crew secures the Mars base and returns to the Phobos gateway in the MCSV. They return to Earth in the gateway refueled MPV, which is then used in subsequent cycles. 3 Description of Potential Nuclear Elements Since this paper addresses the benefits of space nuclear power we will only describe, in depth, those transportation and surface elements for which acceptable nuclear alternatives exist. At this moment, the acceptable nuclear alternatives are all orbit-to-orbit transfer vehicles. Other nuclear alternatives have been proposed (e.g. for surface landers), but they present serious environmental concerns and are beyond the scope of this paper. Lunar Transfer Vehicles The lunar transfer vehicle is capable of delivery of cargo or crew modules to Low Lunar Orbit (LLO), and return of cargo or crew to LEO. The baseline propulsion system is LO2/LH2, with other options including nuclear thermal rockets, and nuclear electric propulsion. For the purposes of this study, it was assumed the baseline LEO- to-LLO transportation vehicle would be the LO2/LH2 Space Transportation Vehicle (STV), which is the subject of current Phase A studies. Support of the lunar evolutionary mission cases is a primary STV design requirement. We will assume the manned missions will always use chemical propulsion, and concentrate on the lunar cargo missions as possible applications for space nuclear power. For cargo missions, the Lunar Cargo Vehicle assembled in LEO includes the transfer stage with droptanks, and the fully fueled lunar cargo lander with payload. Upon arrival at lunar orbit, the cargo lander detaches and performs its lunar descent and cargo delivery. The lunar transfer stage departs lunar orbit and splits into tankset and propulsion module prior to Earth aeromaneuver. The tankset is expended and the propulsion module performs an aeromaneuver and returns to LEO for reuse. A sequential performance summary is shown for the chemical lunar cargo mission in Table I. The nuclear thermal rocket propulsion system could also be a partially reusable system, with the propulsion elements and avionics packaged in an aerobraked module and the hydrogen propellant tankage discarded after propellant depletion. Trip time for the NTR system is comparable to the chemical system, as both systems require similar delta-V's for lunar transfer. A sequential performance summary for the NTR
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