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Low-thrust Propulsion Technologies, Mission Design, and Application 231 The targets for a hypothetical “Super-Dawn” mission were chosen from a list of high interest targets formulated by the scientific community Based on preliminary analysis of throughput requirements and delivered mass, a single spacecraft, with only a 5-kW array, could be used to rendezvous with four high interest near-Earth targets shown in table 1 The final delivered mass is comparable to the Dawn spacecraft The “Super-Dawn” mission illustrates the tremendous potential of electric propulsion for these types of missions Studies have looked at using a single spacecraft for tours of near-Earth objects, main-belt asteroids, and even Jupiter Trojans

Sample return missions are multi-body missions because they need to return to Earth Sample return missions are often considered high priority because of the higher fidelity science that can be performed terrestrially Mars sample return was under investigation for many years, but the large costs of such a mission has deterred its implementation Regolith from Phobos and Deimos are of high scientific value The mission options offer significantly lower cost with minimal technology development required

Segment Target Start Mass, kg Required, kg Propellant End Mass, kg

Table 1 Table of ΔV for a “Super-Dawn” type mission

Two concepts for a Phobos and Deimos sample return mission were evaluated using solar electric propulsion: a single spacecraft to both moons or twin spacecraft capable of returning samples from either moon The small bodies of Phobos and Deimos, with small gravity fields (especially Deimos), make electric propulsion rendezvous and sample return missions attractive Electric propulsion systems can be used for the transfer to Mars, and then to spiral into an orbit around the moons Chemical systems cannot easily leverage the Oberth effect for the sample return mission from Mars‘ moons because of the higher altitude orbit requirement So while the mission can be completed, it comes at a large mass penalty Figure 12 illustrates the benefits of using electric propulsion for a Phobos and Deimos sample return mission

Results show significant savings for using electric propulsion for Phobos and Deimos sample return missions The baseline case uses a NEXT thruster with one operating thruster, and a spare system for redundancy (1+1) A Delta II class launch vehicle is capable of delivering enough mass for a sample return from both targets For electric propulsion, the transfer between Phobos and Deimos has minimal mass implications The mass and technology requirements could potentially fit within the Mars Scout cost cap

Using an Evolved Expendable Launch Vehicle (EELV), twin electric propulsion vehicles can

be sent for a low-risk approach of collecting samples from Phobos and Deimos independently However, the use of an EELV enables a chemical solution for a sample return mission Going to a single moon chemically remains a significant challenge and results in a spacecraft that is greater than 70 percent propellant; a mass fraction more typical

of a launch vehicle stage Launching a single chemically propelled spacecraft to retrieve samples from both moons requires staging events adding risk and complexity

Fig 12 Comparison of required launch mass for chemical and EP Mars’ moons missions The use of electric propulsion was studied for various comet surface sample return (CSSR) missions The results are highly dependant on the targets of interest Electric propulsion compares favorably with chemical alternatives resulting in either higher performance or reduced trip times Studies for Temple 1 (Woo et al., 2006) determined the SOA NSTAR thruster to be inadequate due to its propellant throughput capability The mission required the use of a NEXT thruster Studies for the comet Wirtanen (Witzberger, 2006) were conducted and determined that the NSTAR could not deliver positive payload while both the NEXT and HiVHAC thrusters can complete the mission with sufficient margin The largest benefit is that electric propulsion enables a wide range of targets that cannot be reached using chemical propulsion systems

In 2008, NASA GRC completed a mission design study for a multiple near-Earth asteroid sample return mission (Oleson et al., 2009) The results indicated that it is feasible to use electric propulsion to collect multiple samples from two distinct targets in very different orbits An Earth fly-by was performed after leaving the primary target and before arriving at the second to releae the sample return capsule for a lower risk mission and mass savings to the secondary target This mission was not feasible using chemical propulsion The conceptual spacecraft for the multi-asteroid sample return mission is shown in figure 13

4.2 Inclined targets

Other missions enabled by electric propulsion are missions to highly inclined targets There are several Earth crossing targets that are thought to be old and inactive comets These asteroids typically have inclined orbits The ∆V requirement for a plane change is a function

of the spacecraft velocity and angle of the plane change as shown in equation 1 With the Earth’s heliocentric orbital speed near 30 km/s, a simple plane change of even 30 degrees will require a ∆V of at least 15 km/s to perform a fly-by, following equation 4

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The electric propulsion transfer to Tantalus is also a challenging mission The low-thrust transfer is over 30 km/s over 4.5 years, but can still deliver over 800 kg of dry mass on a rendezvous mission using an Atlas V The mission would require two NEXT thrusters, and would not be viable with the NSTAR or Hall thruster based propulsion system Rather than going to high AU to perform the plane change, the low-thrust transfer gradually performs the plan change through several revolutions Figure 15 illustrates the low-thrust transfer to Tantalus Because of the advantages of electric propulsion, efficient use of propellant and low-thrust trajectory options, scientists can plan missions to high interest targets previously unattainable

Fig 15 Optimal low-thrust trajectory to Tantalus

4.3 Radioisotope electric propulsion

Another area of interest pushing the limits of propulsion technology is the use of a radioisotope power source with an electric propulsion thruster This achieves high post launch ∆V on deep-space missions with limited solar power Radioisotope electric propulsion systems (REPS) have significant potential for deep-space rendezvous that is not possible using conventional propulsion options

One example of mission that can benifit from REPS is a Centaur orbiter The Centaurs are of significant scientific interest, and recommended by the Decadal Survey Primitive Bodies Panel as a New Frontiers mission for reconnaissance of the Trojans and Centaurs The original recommendation was for a flyby of a Jupiter Trojan and Centaur While a flyby mission can use imaging, imaging spectroscopy, and radio science for a glimpse at these objects, a REP mission provides an opportunity to orbit and potentially land on a Centaur This greatly increases the science return An exhaustive search of Centair obiter missions concluded that a wide range of Trojan flybys with Centaur Rendezvous missions are pracitical with near-term electric propulsion technology and a Stirling radioisotope generator (Dankanich & Oleson, 2008) With near-term technology, flyby missions may no longer be scientifically acceptable Investigations are continuing using the enabling combination of electric propuslion and radioisotope power systems On-going and recent studies include multi-Trojan landers, Kuiper-belt object rendezvous, Titan-to-Enceldaus

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Fig 17 Trends of commercial satellite beginning-of-life (BOL) P/M ratio and average mass

Fig 18 GTO-to-GEO transfer times as a function of spacecraft specific power

Low-thrust Propulsion Technologies, Mission Design, and Application 237

Fig 19 Required ΔV from GTO-to-GEO as a function of spacecraft specific power

The GTO-to-GEO transfer time and ∆V is dependant on the launch site, or initial starting inclination Figure 20 illustrates the penalty of launch at inclined launch sites and the benefit

of near-equatorial launches

Fig 20 Effect of starting inclination on transfer time and ΔV from GTO-to-GEO

There were 32 commercial communication satellites launched in 2005 and 2006 as provided

by the Union of Concerned Scientists database These specific satellites were evaluated for potential to use an integrated electric propulsion system with a specific impulse of 1000 seconds, 1500 seconds, and 2100 seconds Integrated electric propulsion systems assume the use of 95% of the onboard solar array power of the spacecraft as launched

Using electric propulsion for the GEO insertion has significant mass benefits Typically this

is evaluated as a method to leverage the launch vehicle performance to deliver the greatest possible mass Another perspective is to evaluate the potential for existing launch vehicles to meet the demands of the COMSAT market Figure 21 illustrates that currently launch vehicles with GTO drop mass capabilities in excess of 7,500kg are required for a complete market capture However, using electric propulsion, a launch vehicle with a drop mass capability of 5,500kg can have complete market capture A low cost launcher with a capability to deliver 3,500kg to 5,500kg can create a paradigm shift in the commercial launch market This assumes the commercial entity is willing to endure the long transfer time, ranging from 66–238 days, depending on the spacecraft power-to-mass ratio and EP thruster selected

Fig 21 Capture fraction as a function of GTO drop mass for various propulsion options

6 Conclusion

Electric propulsion technology is widely used today, and multiple thrusters exist for primary electric propulsion application NASA and the U.S commercial market developed several thrusters suitable for primary electric propulsion on full scale spacecraft The

Low-thrust Propulsion Technologies, Mission Design, and Application 239 technology drivers for new electric propulsion thrusters include: ability to use available power (i.e high maximum power with large throttle range), increased total throughput capability, and lower cost systems and integration The optimal specific impulse is limited

by thrust required to minimize propulsive inefficiencies and available power Due to power constraints, the optimal specific impulse is typically less than 5,000s and closer to 2,000s for near-Earth application Electric propulsion is an enabling technology for a large suite of interplanetary missions Several targets are infeasible with advanced chemical propulsion technologies, while practical with today’s electric propulsion options Electric propulsion is well suited for missions with very high post-launch ∆Vs including multi-target missions, sample return missions, deep-space rendezvous, and highly inclined targets Electric propulsion has tremendous capability to impact the commercial launch market by leveraging on-board available power Today’s commercial satellites have mass-to-power ratios for practical GTO-to-GEO low-thrust transfer As available power and performance demand continues to rise, electric propulsion technologies will continue to supplant chemical alternatives for a wide range of missions The technology will continue to focus on developing lower cost propulsion systems with higher power and longer lifetime capabilities

7 References

Brophy, J R (2007) Propellant Throughput Capability of the Dawn ion Thrusters,

IEPC-2007-279, 30th International Electric Propulsion Conference, Florence, Italy, September 2007

Brophy, J., Rayman, M D., & Pavri, B (2008) Dawn: An Ion-propellanted Jounrey to the

Beginning of the Solar System, IEEE Aerospace Conference, Big Sky, MT, March

2008

Byers, D., & Dankanich, J W (2008) Geosynchronous-Earth-Orbit Communication Satellite

Deliveries with Integrated Electric Propulsion Journal of Power and Propulsion, Vol

24, No 6, November–December 2008, pp 1369–1375

Dankanich, J W & Oleson, S R (2008) Radioisotope Electric Propulsion (REP) Centaur

Orbiter Mission Design, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Hartford, CT, July 2008

Dankanich, J W., & Woodcock, G R (2007) Electric Propulsion Performance from

GEO-Transfer to Geosynchronous Orbits, International Electric Propulsion Conference, Florence, Italy, September 2007

Kamhawi, H., Manzella, D., Pinero, L., & Mathers, A (2009) Overview of the High Voltage

Hall Accelerator Project, AIAA 2009-5282, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Denver, CO, August 2009

Manzella, D (2007) Low Cost Electric Propulsion Thruster for Deep Space Robotic

Missions, 2007 NASA Science Technology Conference, University of Maryland,

MD, June 2007

Oh, D (2007) Evaluation of Solar Electric Propulsion Technologies for Discovery-Class

Missions Journal of Spacecraft and Rockets, Vol 44, No 2., March-April 2007, pp 399–

411

Oh, D., Witzberger, K., & Cupples, M 2004) Deep Space Mission Applications for NEXT:

NASA’s Evolutionary Xenon Thruster, AIAA-2004-3806, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Fort Laudersale, FL, July

2004

Oleson, S et al (2009) Near-Earth Asteroid Sample Return Mission, 31st International

Electric Propulsion Conference, Ann Arbor, MI, September 2009 (to be publsihed) Patterson, M & Benson, S (2007) NEXT Ion Propulsion System Development Status and

Performance, AIAA-2007-5199, 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Cincinatti, OH, July 2007

Welander, B., Carpenter, C., de Grys, K., Hofer, R R., Randolph, T M., & Manzella, D H

(2006) Life and Operating Range Extension of the BPT-4000 Qualification Model Hall Thruster, AIAA 2006-5263, 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Sacramento, CA, July 2006

Wilson, F., & Smith, B (2006) Hall Thruster System Qualification Provides Major Satellite

Benefits, IAC-06-C4.4.09, 57th International Astronautical Congress, Valencia, Spain, October 2006

Witzberger, K E (2006) Solar Electric Propulsion for Primitive Body Science Missions,

NASA TM 2006-214236, March 2006

Woo, B., Coverstone, V L., & Cupples, M (2006) Application of Solar Electric Propulsion to

a Comet Surface Sample Return Missions Journal of Spacecraft and Rockets, Vol 43,

No 6, November–December 2006, pp 1225–1230

13

Global GNSS Radio Occultation Mission for

Meteorology, Ionosphere & Climate

Nick L Yen1, Chen-Joe Fong1, Chung-Huei Chu1, Jiun-Jih Miau1,

Yuei-An Liou2, and Ying-Hwa Kuo3

1National Space Organization

2Center for Space and Remote Sensing Research, National Central University

3University Corporation for Atmospheric Research

a spacecraft passes through an intervening planetary atmosphere before arriving at the receiver, and is used to study the planetary atmosphere properties in the interplanetary mission (Fjeldbo & Eshleman, 1965) The atmospheric RO observations represent a planetary scale geometric optics experiment in which the atmosphere acts as a big optical lens and refracts the paths and propagation velocity of electromagnetic wave signals passing through

it (Kursinski et al., 2000) The first RO experiment started with the Mars flyby by

Mariner-IV in 1964 (Kliore et al., 1965) When Mariner-Mariner-IV satellite passed behind and emerged from the other site of Mars, the extra carrier phase delay and amplitude variation of the microwave signals were observed These observed data provided very first valuable atmospheric and ionospheric density information by using the inversion techniques (Melbourne et al., 2005) Since then a series of planetary experimental missions were planned to study the atmospheres and ionospheres of the planets and their moons (Yunck et al., 2000)

The limb sounding of the earth’s atmosphere and ionosphere using the RO technique can be performed with any two cooperating satellites before the United States’ Global Positioning System (GPS), the first Global Navigation Satellite Systems (GNSS), becoming operational (Lusignan et al., 1969) A few early RO experiments from a satellite-to-satellite tracking link had been conducted These included the occulted radio link between ATS-6 (Applications Technology Satellite-6) and GEOS-3 (Geodetic and Earth Orbiting Satellite-3) and between the Mir station and a geostationary satellite (Liu et al., 1978; Yakovlev et al., 1996)

1 http://en.wikipedia.org/wiki/Occultation [cited 1 July 2009]

2 GNSS radio occultation mission

After GNSS becomes operational, substantial and significant progress has been made in the science and technology of ground-based and space-based GNSS atmospheric remote sensing over the past decade (Davis et al., 1985) The ground-based GNSS atmospheric remote sensing with upward-looking observations arose in the 1980s from GNSS geodesy As the rapid increase of the GNSS geodetic ground networks around the world, great quantity of atmospheric integrated perceptible water (PW) were used in numerical weather prediction (NWP) for weather and climate modelling (Liou et al., 2000 & 2001; Elgered et al., 2003; Ha

et al., 2003) However, one of the major limitations to the ground-based GNSS remote sensing is that it only provides integrated PW without vertical resolution, and it is restricted

to land areas distributed with GNSS networks The space-based GNSS atmospheric limb sounding offers a complementary solution to these issues (Yunck et al., 2003)

The space-based GNSS RO atmospheric remote sensing technique, which makes use of the L-band radio signals transmitted by the GNSS satellites, has emerged as a powerful approach for sounding the global atmosphere in all weather over both lands and oceans (Yunck et al., 1990 & 2003; Wu et al., 1993; & Liou et al., 2002) Figure 1 shows a schematic diagram illustrating radio occultation of GNSS signals received by a low-earth-orbit satellite The GPS/Meteorology (GPS/MET) experiment (1995-1997) showed that the GNSS RO technique offers great advantages over the traditional passive microwave measurements of the atmosphere by satellites and became the first space-based “proof-of-concept” demonstration of GNSS RO mission to earth (Ware et., 1996; Kursinski et al., 1996; Rius et al., 1998; Anthes et al., 2000; Hajj et al., 2000; Kuo et al., 2000) For a more complete history

of GNSS RO see Yunck et al (2000) and Melbourne et al (2005)

Fig 1 Schematic diagram illustrating radio occultation of GNSS signals

The extraordinary success of GPS/MET mission had inspired a series of other RO missions, e.g., the Ørsted (in 1999), the SUNSAT (in 1999), the Satellite de Aplicaciones Cientificas-C (SAC-C) (in 2001), the Challenging Minisatellite Payload (CHAMP) (in 2001), and the twin

Global GNSS Radio Occultation Mission for Meteorology, Ionosphere & Climate 243 Gravity Recovery and Climate Experiment (GRACE) missions (in 2002) The GNSS RO sounding data have been shown to be of high accuracy and high vertical resolution Table 1 lists GNSS RO sounding data characteristics All these missions set the stage for the birth of the FORMOSA SATellite mission-3/Constellation Observing Systems for Meteorology, Ionosphere, and Climate mission, also known as FORMOSAT-3/COSMIC mission Kursinski et al., 1996; Rius et al., 1998; Anthes et al., 2000; Hajj et al., 2000; Kuo et al., 2000; Lee et al., 2001)

Characteristics of GNSS Radio Occultation Data

• Limb sounding geometry complementary to ground and space nadir viewing instruments

• Global 3-D atmospheric weather coverage from 40 km to sea level surface

• High accuracy temperature measurement (equivalent to <1 K; average

accuracy <0.1 K)

• High precision temperature measurement (0.02-0.05 K)

• High vertical resolution (0.1 km surface – 0.1 km tropopause)

• Only system from space to resolve atmospheric boundary layer

• All weather-minimally affected by aerosols, clouds or precipitation

• Independent height and pressure

• No first guess sounding requirement

• Independent of radiosonde calibration

• No instrument drift

• No satellite-to-satellite bias

• Compact sensor, low power, low cost

• A typical RO sounding showing very sharp tropopause

• No other instrument from space provides such high vertical resolution profile Table 1 Characteristics of GNSS radio occultation data

3 FORMOSAT-3/COSMIC mission

3.1 Mission

The FORMOSAT-32 satellite constellation was launched successfully from Vandenberg Air Force Base in California 1:40 UTC on April 15, 2006 into the designated 516 km circular parking orbit altitude Table 2 shows the mission characteristics The FORMOSAT-3 mission is the world’s first demonstration of GPS RO occultation near real-time operational constellation mission for global weather monitoring The primary scientific goal of the mission is to demonstrate the value of near-real-time GPS RO observation in operational numerical weather prediction With the ability of performing both rising and setting occultation, the mission provides about 1,600~2,400 atmospheric and ionospheric soundings per day in near real-time that give vertical profiles of temperature, pressure, refractivity, and water vapor in neutral atmosphere, and electron density in the ionosphere with global coverage (Anthes et al., 2000 & 2008; Liou et al., 2006a, 2006b, & 2007; Fong et al., 2008a & 2009a)

2 In this chapter the FORMOSAT-3/COSMIC mission was referred to as the FORMOSAT-3 mission for simplicity

The retrieved RO weather data are being assimilated into the NWP models by many major weather forecast centers and research institutes for real-time weather predictions and cyclone/typhoon/hurricane forecasts (Kuo et al., 2004; Anthes et al., 2008) The mission results have shown that the RO data from FORMOSAT-3 are of better quality than those from previous missions and penetrate much further down into the troposphere, mission results could be referred to Liou et al (2007), Anthes et al (2008), Fong et al (2008a, 2008b, 2008c & 2009a), and Huang et al (2009) In the near future, other GNSS, such as the Russian Global Navigation Satellite System (GLONASS), and the planned European Galileo system, could be used to extend the applications by the use of RO technique (Chu et al., 2008; Fong

et al., 2009a & 2009b) The great success of the FORMOSAT-3 mission expected to operate through 2011, has initiated a new era for near real-time operational GNSS RO soundings (Fong et al., 2009b; Kuo et al., 2004, 2008a & 2008b)

Number Six identical satellites

Weight ~ 61 kg (with payload and fuel)

Shape Disc-shape of 116 cm diameter, 18 cm in height

Orbit 800 km altitude, circular

Inclination Angle 72o

Argument of latitude 52.5o apart

Communication S-band uplink (32 kbps) and downlink (2 Mbps)

Sounding 1,600 ~ 2,400 soundings per day

Data Latency 15 minutes to 3 hours

Design and Mission life 5 years

Launch date April 15, 2006

Table 2 The FORMOSAT-3 mission characteristics

3.2 System architecture

Figure 2 shows the FORMOSAT-3 system architecture After two years’ in orbit operations, starting from mid-April 2008, the FORMOSAT-3 program switched and changed from two commercially operated ground stations at Fairbanks, Alaska and Kiruna, Sweden, operated

by United Service Network (USN), to two new ground stations in Fairbanks and Tromso, Norway, operated by National Oceanic and Atmospheric Administration (NOAA) The constellation operation plans to use the new stations for the remaining five-year mission The FORMOSAT-3 constellation system consists of the six microsatellites, a Satellite Operations Control Center (SOCC) in Taiwan, several tracking, telemetry and command (TT&C) ground stations, two data receiving and processing centers, and a fiducial network There are two TT&C Local Tracking Stations (LTS), one located in Chungli, Taiwan and the other in Tainan, Taiwan, respectively Currently there are four Remote Terminal Stations (RTS) to support the passes: Fairbanks Command and Data Acquisition Station (FCDAS), and Kongsberg Satellite Services Ground Station (KSAT), which are currently set as primary stations for the FORMOSAT-3 mission, and the Wallops station at Virginia, USA and the McMurdo station located in McMurdo, Antarctica The latter two RTS stations provide partial support for the mission (Fong et al., 2009a; Rocken et al., 2000)

Global GNSS Radio Occultation Mission for Meteorology, Ionosphere & Climate 245 The SOCC uses the real-time telemetry and the back orbit telemetry to monitor, control, and manage the spacecraft state-of-health The downlinked science RO data is transmitted from the RTS via National Oceanic and Atmospheric Administration (NOAA) to CDAAC (COSMIC Data Analysis and Archive Center) located at Boulder, Colorado, USA, and TACC (Taiwan Analysis Center for COSMIC) located at Central Weather Bureau (CWB) in Taiwan The fiducial GNSS data is combined with the occulted and referencing GNSS data from the GOX payload to remove the clock errors All collected science data is processed by CDAAC and then transferred to TACC and other facilities for science and data archive (Wu et al., 2006)

Fig 2 The FORMOSAT-3 system architecture

The processed results are then passed to the National Environmental Satellite, Data, and Information Service (NESDIS) at NOAA These data are further routed to the weather centers in the world including the Joint Center for Satellite Data Assimilation (JCSDA), National Centers for Environment Prediction (NCEP), European Centre for Medium-range Weather Forecast (ECMWF), Taiwan CWB, UK Meteorological Office (UKMO), Japan Meteorological Agency (JMA), Air Force Weather Agency (AFWA), Canadian Meteorological Centre (Canada Met), French National Meteorological Service (Météo France), etc They are made ready for assimilation into weather prediction models The data

is currently provided to weather centers within 180 minutes data latency requirement in order to be ingested by the operational weather forecast model (Fong et al., 2009b)

4 FORMOSAT-3 satellite design

Figure 3 illustrates the FORMOSAT-3 satellite designed by Orbital Science Corporation in a deployed configuration and its major components The FORMOSAT-3 satellite avionic block

diagram is shown in Figure 4 The major subsystem elements of the spacecraft system are Payload Subsystem, Structure and Mechanisms Subsystem (SMS), Thermal Control Subsystem (TCS), Electrical Power Subsystem (EPS), Command and Data Handling Subsystem (C&DH), Radio Frequency Subsystem (RFS), Reaction Control Subsystem (RCS), Attitude Control Subsystem (ACS) and Flight Software Subsystem (FSW) The spacecraft bus provides structure, RF power, electrical power, thermal control, attitude control, orbit raising, and data support to the instrument (Fong et al, 2008a, 2008b & 2009a) Table 3 shows the spacecraft bus key design features

GPS Occultation Experiment (GOX)

Tiny Ionospheric Photometer

(TIP)

Tri-Band Beacon (TBB)

Electronics

Flight Computer (FC - upper) Attitude Control Elec

(ACE - middle) GPS Receiver (bottom)

Mission Interface Unit (MIU)/

Power Control Module (PCM)

Battery

Battery Charge Regulator

(BCR)

TBB Antenna S-band Antenna

GOX OCC2 Antenna

GOX POD1 Antenna

Earth Horizon Sensor (EHS)

GOX POD2 Antenna

GPS Occultation Experiment (GOX)

Tiny Ionospheric Photometer

(TIP)

Tri-Band Beacon (TBB)

Electronics

Flight Computer (FC - upper) Attitude Control Elec

(ACE - middle) GPS Receiver (bottom)

Mission Interface Unit (MIU)/

Power Control Module (PCM)

Battery

Battery Charge Regulator

(BCR)

TBB Antenna S-band Antenna

GOX OCC2 Antenna

GOX POD1 Antenna

Earth Horizon Sensor (EHS)

GOX POD2 Antenna

GPS Occultation Experiment (GOX)

Tiny Ionospheric Photometer

(TIP)

Tri-Band Beacon (TBB)

Electronics

Flight Computer (FC - upper) Attitude Control Elec

(ACE - middle) GPS Receiver (bottom)

Mission Interface Unit (MIU)/

Power Control Module (PCM)

Battery

Battery Charge Regulator

(BCR)

TBB Antenna S-band Antenna

GOX OCC2 Antenna

GOX POD1 Antenna

Earth Horizon Sensor (EHS) GOX POD2 Antenna

Fig 3 The FORMOSAT-3 satellite in deployed configuration and its major components

Structure Metal Matrix (AlBeMet)

Science Data Storage 128 MB

Distributed Architecture Motorola 68302 Microprocessor

Attitude Control Magnetic 3-axis Control Pointing Control = 5° Roll & Yaw, 2 ° Pitch

Propulsion Hydrazine Propulsion Subsystem

S-Band Communications HDLC Command Uplink (32 kbps) CCSDS Telemetry Downlink (2 Mbps)

Single String Bus Constellation Redundancy

Table 3 The FORMOSAT-3 spacecraft bus key design