A Satellite for Space Qualifying Electronics with in Van Allen Radiation Belts | M 385C, Papers of Probability and Statistics

Download A Satellite for Space Qualifying Electronics with in Van Allen Radiation Belts | M 385C and more Papers Probability and Statistics in PDF only on Docsity! Holt & Wagner - i A Satellite for Space Qualifying Electronics within the Van Allen Radiation Belts Greg Nate Holt & Ian Christopher Wagner The University of Texas at Austin Holt & Wagner - ii Abstract The VALOR (Van Allen Orbital Radiation) study involves developing an inexpensive mission plan to space qualify various electronics by conceiving processes and equipment that will determine the effects of the Van Allen Radiation Belts. The Van Allen Radiation Belts are bands of high-energy particles surrounding the earth, held by earth's magnetic field. Prolonged exposure to the Van Allen Belts can damage spacecraft electronics. The VALOR mission plan will serve as a preliminary study to aid in NASA’s design for the 2014 manned Mars mission, which will have high exposure to the radiation belts for a long duration. Four major mission areas have been developed: mission concepts, launch, orbit, and retrieval. Two concepts for mission operations were considered. The first concept depended on manufacturing one satellite to collect radiation belt data that would be used to recreate the space radiation environment in a particle accelerator. This concept was determined to be unfeasible. The second concept relied on manufacturing multiple, inexpensive satellites for repeated experiments in the Van Allen Belts. The Sea Launch commercial launch system was selected for the mission. Two orbit scenarios were evaluated: a dual circular orbit with a Hohmann transfer and a slowly decaying orbit scenario. The slowly decaying orbit scenario was selected. It has been decided to use an aerial capture to retrieve the satellite. This method was used by the Corona mission to retrieve film canisters deployed from spy satellites. Seven subsystems were designed for the satellite bus: electronics experiment, propulsion, communication, command and data handling, power, structures and mechanisms, and thermal. It was determined that the overall satellite structure would be a spherical, layered inflatable satellite with flexible solar panels and six omnidirectional antennas. In addition, retrorockets, parachutes, and GPS antennas allow for precise retrieval. The estimated satellite manufacture cost would be $6.34 million. Holt & Wagner - 1 A Satellite for Space Qualifying Electronics within the Van Allen Radiation Belts Greg Nate Holt & Ian Christopher Wagner The University of Texas at Austin 1.0 Introduction 1.1 Project Background The NASA Mars Exploration Study Team is in the process of designing a manned mission to Mars for the year 2014. The overall mission will consist of three launches to Mars. The first two launches will be performed in 2011. They will be unmanned and carry the Cargo Lander 1 and the Earth Return Vehicle. The third launch will be in 2014 and will be manned with six astronauts and carry the Crew Lander [6]. One aspect of the mission involves boosting cargo from Earth's gravity well to Mars. This transfer requires a large amount of energy, and both solar-electric and nuclear-electric propulsion systems are being considered for this problem. One proposal by the NASA Exploration Team is to use solar-electric propulsion to perform the bulk of the trans-Mars injection. A solar-electric space "tug" vehicle will perform a low-thrust transfer of the cargo from a circular, low-Earth orbit to an elliptic, high-apogee Earth orbit. This transfer period will take approximately nine months. At that time, the cargo will be injected into a Mars-bound orbit with a small chemical kick stage. The solar-electric space tug will again return to low-Earth orbit and repeat the process for the next cargo boost. For the 2014 manned mission, the process is the same except the crew will wait and take a high speed "taxi" to the final elliptical Earth Holt & Wagner - 2 orbit when the nine-month transfer is complete. This will save the astronauts from repeated exposure to Van Allen Belt radiation and excess time in space. [6] The solar-electric space tug will be crossing the Van Allen Radiation Belts many times during its transfer process from low to eccentric Earth orbit. The radiation belts could cause potential problems for the on-board equipment, such as solar cells and semiconductors. The extent of this problem, though, has not yet been determined [13,6] 1.2 The Van Allen Radiation Belts The Van Allen Belts were discovered in 1958 by the physicist James Van Allen. The Earth’s magnetic field traps ions from cosmic and solar sources, creating a magnetosphere. The inner and outer Van Allen Belts are a part of Earth’s inner magnetosphere. The inner belt is found above the equator at an altitude of 1 to 2 Earth radii. This belt is composed of protons with energy ranging from 0.1 to 400 MeV (1.2×10-14 to 4.7×10-11 ft×lbf) and intensities up to 106 protons/(cm2×sec) (6.5×106 protons/(in2×sec)). The outer belt is composed of ions and electrons with energy ranging from 0.1 to 7 MeV (1.2×10-14 to 8.3×10-13 ft×lbf) and intensities up to 106 electrons/(cm2×sec) (6.5×106 electrons/(in2×sec)) [13]. Figure 1 illustrates the proton and electron intensities. The Van Allen Belts can severely damage an orbiting spacecraft by degrading solar arrays, sensors, and microelectronics devices. The radiation belts have two main effects on electronics: single event effects and total ionizing dose effects. Single event effects occur when a high- energy ion impacts electronics, often causing internal registers to change states. This includes random errors in sequential-logic circuits or memory. Total ionizing dose effects occur over a long duration as ions are trapped in the electronics. This will eventually cause permanent changes in the electronics. [13] Holt & Wagner - 3 1.3 Project Objectives The purpose of the VALOR project is to create a mission plan, orbit trajectory, and spacecraft to space-qualify various electronic components. The major focus of the project is to develop and analyze orbits to accomplish the mission. Details of the orbit, however, depend heavily on the design of the satellite used. To produce an accurate orbit simulation, the VALOR project also contains a conceptual satellite design. The design provides the parameters needed to develop a precise orbit model for analysis. 1.4 Summary of Sections Section two discusses the VALOR mission plan to space-qualify various electronics. This section provides details for mission concepts, launch, orbit, and satellite retrieval. Section three provides details on the satellite bus subsystems. This section discusses the electronics experiment, propulsion, communication, command and data handling, thermal, structures, and power. Section four presents our conclusions. The remainder of this report contains references, figures, tables, and an appendix. 2.0 Mission Areas In order to space-qualify various electronics, the VALOR team will develop orbits, mission plans, and equipment to determine the long-term exposure effects of the Van Allen Radiation Belts. VALOR researchers have established five areas of investigation. These areas are mission concepts, launch, orbit, retrieval, and satellite bus. This section of the report includes a discussion of the mission concepts, launch, orbit, and retrieval. Holt & Wagner - 6 The Kourou, French Guiana launch site has the capability to provide a 5.2o orbit inclination. This inclination is adequate for the proposed orbit of the VALOR satellite. However, since France controls the site, a French launch vehicle would have to be used. Therefore, either an Ariane-44L or the newer Ariane-5, shown in Figure 2, will be used. The Ariane-44L has the capability to send 4,520 kg (10,000 lb) into a geostationary transfer orbit (GTO), with a possible spacecraft diameter of 4 m (13.1 ft). The Ariane-5 has the capability to send 6,800 kg (15,000 lb) into GTO, with a possible diameter of 5.4 m (17.7 ft). Ariane launch vehicles are considered reliable. A typical dedicated launch costs of an Ariane-44L is approximately $110 million. [12] Another option is Sea Launch. Sea Launch is a newly developed commercial system for launching large satellites with a previously unrealized versatility. The Sea Launch system, shown in Figure 3, involves launching from a converted mobile oil-drilling platform, which has the ability to travel to most marine locations. This allows the satellites to be inserted into any orbital inclination. The VALOR project prefers an inclination as close to zero degrees as possible. A direct launch to a zero degree inclination can only be achieved presently by a system such as Sea Launch. The Sea Launch program is able to send a payload of up to 5,000 kg (11,000 lb) of spacecraft mass to GTO. The maximum diameter that the system can support is 4.3 m (14 ft). These weight and size limits are more than adequate to accommodate the VALOR satellite. In addition to the VALOR satellite, Sea Launch can potentially carry several more satellites comparable in size or a few satellites of larger size. This will allow for piggyback or dual-launch opportunities, considering the large demand for small satellites to be placed in equatorial orbits. Since Sea Launch is a commercial enterprise owned by Boeing and several foreign companies, costs are expected to be very competitive with if not less expensive than other options. However, Sea Launch is a very new system, having had their first commercial launch Holt & Wagner - 7 on October 9, 1999. The system requires the use of Boeing's own customized launch vehicle. The VALOR team estimates the total satellite launch cost to be $2.96 million. The cost was based on mass ratios. Our satellite will be approximately 159 kg (350 lb) and will be a piggyback launch with a geostationary satellite. The total cost of a launch can be approximated to be $90 million [19]. Our satellite will be 3.2% of the total mass. Therefore, VALOR estimates the launch cost will be 3.2% of the total launch cost with an additional $100,000 for satellite integration. Although Boeing maintains that the Sea Launch system is composed of proven components, the system as a whole has not been thoroughly tested [3]. However, considering Boeing’s experience in spacecraft systems, the VALOR team selected Sea Launch for our launch scenario. 2.3 Orbit The VALOR researchers have considered two main orbit scenarios for use with the VALOR spacecraft project. These are the dual circular and slowly decaying scenarios. Both need to be of low inclination to maximize the exposure to the Van Allen Radiation Belts. The type and duration of exposure is very different within each scenario. 2.3.1 Dual Circular Orbit The goal of the dual circular orbit scenario, shown in Figure 4, is to maximize the satellite's exposure to both the inner and outer radiation belts in the most direct way. Therefore, the satellite will travel in a circular orbit through the middle of the inner radiation belt for 4½ months. The satellite then performs a Hohmann transfer to a circular orbit through the middle of the outer radiation belt where it will stay for another 4½ months until returning to the Earth. The obvious disadvantage to this option is the need to carry fuel and a propulsion system to perform the orbit change maneuvers. This increases the launch weight and reduces the available space Holt & Wagner - 8 for electronics experiments. In addition, this scenario does not accurately model the radiation exposure that will be experienced by the NASA space tug. 2.3.2 Slowly Decaying Orbit The goal of the slowly decaying orbit scenario, shown in Figure 5 and Figure 6, is to imitate the planned orbit of the NASA trans-Mars Injection Vehicle. NASA has specified that the space tug will continuously collect energy while traveling from a low-Earth orbit to an elliptical Earth orbit. This collection period will take approximately nine months to complete. For this mission scenario, the satellite apogee change will be opposite in order to use atmospheric drag instead of fuel to imitate the space tug orbital path. Therefore, the satellite will go from an elliptical Earth orbit to a low-Earth orbit. Because atmospheric drag is used to adjust the orbit during the mission, only a single, small de-orbit propulsion system is needed. This frees up extra space and mass for electronics experiments. In addition, the type of radiation received during the mission will be the same as that experienced by the space tug. The first iteration for the orbit called for an initial apogee radius equal to the radius of the outer Van Allen Belt. After examination of piggyback options, however, the VALOR team determined that a geostationary satellite would most likely be the host vehicle. For this reason, a second orbit was calculated using an initial apogee radius equal to geostationary orbit radius. Because the slowly decaying orbit scenario depends on atmospheric drag to decrease the radius of the orbit, the exact rate of descent cannot be predicted before the mission. Using the 1976 atmosphere model (the most recently published by NOAA), preliminary estimates of orbit decay for the first two orbits are shown in Figure 7. These estimates were obtained with the Analytical Graphics Satellite Toolkit Long Term Orbit Propagator. Orbit perturbations due to solar radiation pressure, earth gravity field variations, atmospheric drag, lunar gravity, and solar gravity were included. The estimate of orbit decay for the final orbit is shown in Figure 8. This Holt & Wagner - 11 3.1 Electronics Experiments The electronics experiments on the VALOR satellite will consist of components used in the NASA space tug and other electronics that need to be qualified for a high-radiation space environment. These include but are not limited to microprocessors, electronic memory devices, magnetic memory devices, circuit boards, data converters, transceivers, and solar panel parts [14, 16]. In addition, various personal electronics manufacturers would be used for sponsorship purposes and for samples to be included in the electronics experiments. The electronics experiments will be housed in the less shielded outer core area of the spacecraft. The electronics test area is approximately 1.74 m3 (61.34 ft3). The electronics will be exposed to radiation levels similar to what they will experience aboard the space tug. Additionally, there will be three samples of each electronic device. The experimental control device will remain unpowered throughout the duration of the mission. The first experimental device will remain powered throughout the mission. The final experimental device will be powered on and off throughout the mission. Additionally, there will be several Geiger counters to measure radiation levels during the experiment. The measured amount of radiation will be included as part of the spacecraft health status. The ultimate goal of the electronics experiments is to recover, modify, and retest the components until the adequate methods of radiation- hardening and shielding are found. According to VALOR consultant Dr. Glenn Lightsey, in the Van Allen Belts the electronics will be exposed to both modes of radiation failure: total dose and single event effects. The minimal shielding of the secondary core will only provide a slightly lower total dose than free exposure. Single event effects will not be affected by the minimal shielding, but will be partially reduced by the radiation hardening of the electronics. In any case, the heavily- shielded main command computer will provide power to and log the status of the electronics Holt & Wagner - 12 throughout the mission. A lockup or failure will be recorded and the time, radiation dose, and other pertinent information will be correlated with the event. 3.2 Propulsion The propulsion system for the VALOR satellite will be simple, lightweight, and inexpensive. No liquid chemical or electric propulsion will be needed during the nominal orbit since atmospheric drag near perigee will be used for the on-orbit changes. The primary propulsion system under consideration is the use of small outgas vents installed on the inflatable expansion chambers. These vents will be oriented along three axes as shown in Figure 13 so that arbitrary attitude adjustments can be made. Orbital corrections will be made by inflating or deflating the expansion chambers to achieve the cross-sectional area needed for proper drag. The VALOR satellite will also launch directly into an equatorial orbit to avoid the need for an upper-stage inclination-change motor. For de-orbit, the VALOR satellite will use a solid-fuel retrorocket similar to the Thiokol TE-M-385 engines used aboard the Gemini spacecraft. With 11,000 N (2,500 lbf) of thrust, this engine will accomplish the de-orbit burn in approximately 0.7 seconds. [22] 3.3 Communication The communication system will need to provide two-way communication between the satellite and ground control. This will allow the satellite to receive instructions and to transmit collected data and spacecraft health. The satellite will have six omnidirectional antennas evenly spaced along its surface. These antennas will allow the satellite to transmit and receive data without any concern of the attitude of the satellite. Because of the small scale of this mission and the number of orbits that the satellite will make, the VALOR project needs an inexpensive and geographically distributed system to track the satellite. A network composed of satellite dish Holt & Wagner - 13 receivers owned by universities around the world can be used to receive the signals transmitted by the VALOR satellite. VALOR proposes the network include but not be limited to stations in the following locations: 1. United States 2. Mexico 3. Spain 4. Australia 5. Columbia 6. French Guiana 7. Brazil 8. Algeria 9. Egypt 10. Israel 11. India 12. Philippines This network can be easily established with pre-existing facilities for very little cost. The VALOR team predicts that command data will not need to be sent on a regular basis. Therefore, a single transmission station could be located in the United States. Additionally, the satellite will need seven GPS receivers and a radio transponder to aid in tracking the orbital position and post- entry retrieval. 3.4 Command and Data Hand ling Command and data handling will be located in the heavily-shielded core area of the satellite. This system will be responsible for the following six items: 1. General command of satellite (antennas, valves, expansion chambers) 2. Store information gathered from electronics experiments 3. Send periodic health, status, and electronic radiation data 4. Fire retrorockets for de-orbit maneuver 5. Fire drogue chute 6. Deploy descent chute Radiation-hardened computer and memory will be used for the main command functions. The VALOR team has investigated radiation-hardened command systems already in use with geostationary satellites. Several fabrication facilities produce these components, including Holt & Wagner - 16 antennas will be retracted while the satellite is encapsulated in the launch vehicle. On the outer skin next to the omnidirectional antennas will be six GPS patch antennas. These antennas will be used for precise attitude determination directly before re-entry. Also next to the omnidirectional antennas will be outgas vents, one for each expansion chamber. These outgas vents will be used for minor attitude adjustment and depressurization. Several mechanisms will be placed along the outer core. Imbedded in a cavity in the heat shield will be a partially spherical retrorocket with the nozzle protruding from the shell. After the retrorocket has fired, it will be jettisoned leaving a rounded cavity in the front of the heat shield as shown in Figure 14. Recent research by Dr. David Goldstein at the University of Texas suggests this re-entry configuration actually reduces heat generation at hypersonic velocities [19]. On the opposite side of the outer core will sit the parachute assembly. The assembly includes a drogue chute, ring chute, and main chute with top mounted radio transponder and GPS antenna. Inside the outer core will be six pressure vessels, mounted close to the attachment point of the omnidirectional antennas. These will be used to pressurize the inner mantle for the aerogel foaming process and to inflate the expansion chambers. There will also be an aerogel foamer strategically located opposite the batteries to not significantly disturb the center of mass of the satellite. 3.6 Power The VALOR satellite will be powered by solar panels while in sunlight, storing power within batteries for use in periods of darkness. Since the satellite structure needs to remain a sphere for atmospheric drag, the solar panels must be wrapped along the outer skin. This requires the solar panels to be flexible in order for the outer expansion chambers to operate properly. Currently, no solar panels have been developed that are fully flexible, however, several ductile solar panels have been created. The solar panels will be configured as to not Holt & Wagner - 17 hinder the completely unexpanded satellite with the needless additional volume that would occur if they where to be folded and layered. Several small solar arrays will be placed around the outer skin, situated in a manner that would cause the panels to be exactly adjacent while the satellite is unexpanded as in Figure 13a. When the satellite is expanded, the solar panels will be oriented in an evenly distributed configuration as in Figure 13c. This arrangement will maximize the area of the solar panels while minimizing the volume and complexity of the panels while the satellite is within the launch vehicle. When the satellite is unexpanded, the satellite will have a diameter of 2 m (7 ft) (1.8 m (6 ft) for the outer core with 15.2 cm (6 in) on either side for the unexpanded expansion chambers). Therefore, the solar panels will have a surface area of 14.3 m2 (154 ft²). However, as seen in Figure 12, only around 4 ½ of the panels on the sphere will be exposed to the sun at any one time. This results in an effective solar panel area of 3.7 m2 (40 ft²). According to Ref. 15, standard flexible solar panels produce around 149.6 Watts/m2 (.002 hp/ft2), yielding around 550 Watts (0.74 hp) of power for the satellite. This is more than adequate for the satellite operations, the test electronics and recharging of the batteries. The solar panels will be silicon based [1] and will thereby cost $1,800,000 for the entire array [15]. There were three battery types being considered by the VALOR team. They were nickel- cadmium, nickel-hydrogen, and lithium-based batteries. Nickel-hydrogen batteries have generally replaced nickel-cadmium batteries in spacecraft. This is due to nickel-hydrogen battery efficiency and less memory associated with charging and recharging. Although nickel- cadmium batteries can be less expensive, the added mass due to their lower efficiency is not acceptable due to the mass constraints for the proper orbit. This mass consideration led VALOR researchers to explore the possibility of using lithium-based batteries. Lithium batteries can provide 3 to 4 times the amount of energy per unit mass of even nickel-hydrogen batteries. These batteries are also 30% smaller and 35% lighter than nickel-cadmium batteries [21]. The Holt & Wagner - 18 drawback is that lithium-ion batteries do not have the life span that is usually required for satellites. However, the VALOR satellite will only be in orbit for approximately nine months. This is considerably less than the 7 to 15 years (GEO) and the 3 to 6 years (LEO) of most satellites. The VALOR team has decided to use lithium batteries. The battery will be required to operate the satellite while it is in the shadow of the Earth or the moon. The satellite will be launched into an orbit that will have the minimal amount of time in darkness by aligning the apogee toward the sun. Analysis using the Satellite Toolkit High Precision Orbit Propagation module shows the period of the orbit is at maximum around 10 hours, thereby requiring at most 50 minutes of battery power each orbit and during lunar eclipses. The minimum amount of power that the satellite will need during this period of darkness is approximately 200 Watts (0.27 hp). Therefore, the battery will be required to provide 167 Watt×hours (443,000 ft×lb) of energy between recharges. A primary and backup set of lithium-ion batteries that satisfy this constraint will cost approximately $72,000. The batteries will have a mass of 6 kg (13.2 lb) [8]. 3.7 Thermal The VALOR team is investigating a passive thermal-control system for this satellite project. First, the aerogel mantle layer has a thermal conductivity value of R20/inch. This layer has the potential of overheating the electronics. The Two-Phase Heat Transfer Laboratory has designed a heat pipe transfer system that is extremely light and flexible [23]. This system will be used to transfer the heat through the aerogel mantle layer. The initial stage will begin with heated liquid converting to vapor. The vapor will be carried to a radiator at the surface of the spacecraft. The heat will then be dissipated into the space environment. Capillary action will then cause the vapor to return to a liquid phase and begin the cycle again. Fluid flow and surface tension move the vapor/liquid fluid within the heat pipe. The excess heat can additionally be dissipated using the tubing connecting the pressure tanks to the expansion chambers. As gas is Holt & Wagner - 21 5.0 References 1. ALF Enterprises. “Flexible Solar Panels” ALF Enterprises, November 12, 1999. http://www.alfenterprises.com/UnisolarFlex.html 2. Barth, J., LaBel, K., "Core Technologies for Space Systems Presentation", Space Radiation Environments, GSFC, Nov. 1998. 3. Boeing. “Sea Launch” Boeing, October 25, 1999. http://www.boeing.com/defense- space/space/sealaunch/ 4. CORONA Imagery Library. “CORONA Imagery Library” Corona, October 25, 1999. http://www.nro.odci.gov/corona/imagery.htm 5. Diamond, John. “Declassified spy satellite photos help arms reduction effort.” The Associated Press, November 12, 1999. http://www.fas.org/eye/V000135-021699-idx.htm 6. Drake, Bret. "Reference Mission Version 3.0 Addendum to the Human Exploration of Mars", The Reference Mission of the NASA Mars Exploration Study Team, June 1998. 7. Dudley, G. and Verniolle, J. “Secondary Lithium Batteries for Spacecraft” ESA, October 24, 1999. http://esapub.esrin.esa.it/bulletin/bullet90/b90dudle.htm 8. Goddard Space Flight Center Academy. "Easy Low-cost Lunar Explorer (ELLE)." Goddard Space Flight Center. November 11, 1999. http://academy.gsfc.nasa.gov/1998/html/project.html 9. Goddard Space Flight Center Radiation Effects Facility homepage. “GSFC.” Goddard Space Flight Center. October 1, 1999. http://flick.gsfc.nasa.gov/radhome/ref/GSFC_REF.html 10. IMEX homepage. “IMEX.” IMEX homepage @ the University of Colorado, September 22, 1999. http://lasp.colorado.edu/stp/imex/imex_main.html 11. Indiana University Cyclotron Facility homepage. “IUCF.” Indiana University Cyclotron Facility. October 1, 1999. http://www.iucf.indiana.edu 12. Isakowitz, Steven J., International Reference Guide to Space Launch Systems, Second Edition, AIAA, 1995. Holt & Wagner - 22 13. Long, Stacia M. Thesis: Van Allen Orbital Radiation Belts Experimental Satellite Missions. The University of Texas at Austin, Aerospace Engineering and Engineering Mechanics, 1999. 14. Maxwell Space Electronics, Inc. “Product List.” Maxwell Space Electronics, Inc., October 25, 1999. http://www.spaceelectronics.com/pdf/Space%20Products%20pdf/RHList.PDF 15. Mostert, Sias and Milne, Garth W. “SUNSAT: Solutions for Remote Sensing.” University of Stellenbosch, Dept of Electrical and Electronic Engineering, November 13, 1999. http://sunsat.ee.sun.ac.za/sspapers/sats.95/sats.html 16. Research Triangle Institute. “Radiation Hardened Electronics.” Research Triangle Institute, October 25, 1999. http://www.rti.org/units/es/csr/simons.cfm 17. Science@NASA. “The NASA Aerogel Web Page.” Marshall Space Flight Center. October 22, 1999. http://www.aerogels.com 18. Schewe, Phillip F. and Stein, Ben. “Physics News Update Number 30 – THE COST OF THE SUPERCONDUCTING SUPERCOLLIDER.” American Institute of Physics, November 12, 1999. http://www.aip.org/enews/physnews/1991/split/pnu030-4.htm 19. Scott, Terrance L., “Student Design Project.” Sea Launch, November 17, 1999. Email response. 20. Silton, Sidra Idelle. “Numerical and Experimental Investigation to Reduce Hypersonic Nose Tip Ablation.” The University of Texas at Austin, Aerospace Engineering and Engineering Mechanics, November 13, 1999. http://www.ae.utexas.edu/~sidra/research.html 21. Space Daily. “Com Dev Markets Lithium Space Battery” Space Daily, November 13, http://www.spacedaily.com/spacecast/news/battery-99b.html 22. Thiokol Propulsion. “The Thiokol/Gemini Connection.” Thiokol Propulsion, November 13, 1999. http://www.thiokol.com/Gemini.htm 23. Two-Phase Heat Transfer Laboratory., “Capillary Pumped Loop” Texas A&M University, November 17, 1999. http://two-phaseheat.tamu.edu/cpl.html Holt & Wagner - 23 6.0 Figures and Tables Figure 1 - Proton and Electron Intensities [2] Figure 2 - Ariane-5 Launch Vehicle Holt & Wagner - 26 0 5000 10000 15000 20000 25000 30000 35000 40000 1/1/00 2/4/00 3/9/00 4/12/00 5/16/00 6/19/00 7/23/00 8/26/00 9/29/00 Date (UTCG) A lt it u d e (k m ) Apogee Altitude (Orbit 2) Apogee Altitude (Orbit 1) Perigee Altitude (Orbit 2) Perigee Altitude (Orbit 1) Figure 7 - Estimated Orbit Decay Using 1976 Standard Atmosphere 0 5000 10000 15000 20000 25000 30000 35000 40000 1/1/00 2/4/00 3/9/00 4/12/00 5/16/00 6/19/00 7/23/00 8/26/00 9/29/00 Date (UTCG) A lt it u d e (k m ) Apogee Altitude (Final) Perigee Altitude (Final) Figure 8 - Estimated Orbit Decay Using High Precision Propagation Holt & Wagner - 27 0 30 60 90 120 150 180 210 240 270 300 330 360 1/1/00 2/4/00 3/9/00 4/12/00 5/16/00 6/19/00 7/23/00 8/26/00 9/29/00 Date (UTCG) A rg u m en t o f P er ig ee ( d eg ) Figure 9 - Argument of Perigee Drift for a Nine-Month Mission 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 1/1/00 2/4/00 3/9/00 4/12/00 5/16/00 6/19/00 7/23/00 8/26/00 9/29/00 Date (UTCG) E cl ip se T im e (s ec ) Lunar Eclipse Figure 10 - Estimated Eclipse Times for Nine-Month, Equatorial Mission Holt & Wagner - 28 Figure 11 - Aerial Capture [4] Figure 12 - Preliminary Proposal for the VALOR Satellite Inner Core: Satellite Operation Systems Outer Core: Test Electronics Heat Shield Gas Channels Inner Mantle: Aerogel Chamber Outer Mantle: Expansion Chambers Crust Flexible Solar Panel GPS Patch Antenna Retraction Cable: Corner Cube Reflector Tape Outgas Jet Omnidirectional antenna