福衛三號大氣遙測星系效能及部署技術挑戰與展望

國

立

交

通

大

學

光電工程研究所

博 士

論

文

福衛三號大氣遙測星系效能及部署技術挑戰與展望

FORMOSAT-3 Constellation Performance, Deployment

Challenges, and Prospect for Atmospheric Remote Sensing

研 究 生：方振洲

指導教授：祁 甡

福衛三號大氣遙測星系效能及部署技術挑戰與展望

FORMOSAT-3 Constellation Performance, Deployment Challenges,

and Prospect for Atmospheric Remote Sensing

研 究 生：方振洲 Student：Chen-Joe Fong

指導教授：祁 甡 Advisor：Sien Chi

國 立 交 通 大 學

光 電 工 程 研 究 所

博 士 論 文

Submitted to Institute of Electro-Optical Engineering College of Electrical Engineering

National Chiao Tung University in partial Fulfillment of the Requirements

for the Degree of Doctor of Philosophy

in

Electro-Optical Engineering March 2009

Hsinchu, Taiwan, Republic of China

福衛三號大氣遙測星系效能及部署技術挑戰與展望

學生：方振洲

指導教授：祁 甡

國立交通大學光電工程研究所

摘 要

全球導航衛星系統(GNSS)無線電掩星(Radio Occultation，簡稱 RO)技術有別

於傳統的衛星微波輻射計，是一個利用地球尺度的幾何光學折射原理

用於

大氣遙測的

先進邊緣探空太空遙測技術。此技術主要係接收經過地球遮掩

的 GNSS 衛星所傳送的電磁波折射信號，由電磁波訊號穿過電離層和大氣

層時受電子密度、溫度、壓力、及水氣等影響而改變信號的時間延遲，反

演推算行進路徑下的電離層和大氣層相關的資料。福爾摩沙衛星三號

(FORMOSAT-3，簡稱福衛三號)任務，又名「氣象、電離層及氣候之衛星

星系觀測系統」(Constellation Observing System for Meteorology, Ionosphere

and Climate，簡稱 COSMIC)任務，係由六顆同型實驗微衛星組成，是世界

上第一個進行全球氣象監測的近實時運作展示的 GPS RO 衛星星系觀測系

統。福衛三號於 2006 年 4 月中旬，在美國加州的范登堡空軍基地發射升空

到地表 516 公里的暫駐軌道上。六顆衛星本體完成入軌健康檢查之後，開

始進行三個衛星酬載包括 GPS 氣象量測儀(簡稱 GOX)、小型電離層光度計

及三頻段信標儀的一系列入軌儀器健康檢查、校正及實驗。隨後展開星系

部署工作，前後共歷經 19 個月，近 500 次軌道轉換，每一顆衛星分別升軌

到高度約 800 公里的全球均等分佈的六個軌道面上，福衛三號成為世界上

第一個利用先進的地球進動理論進行星系部署的系統。微衛星的質量參數

資料，將可供學術進行後續大地重力場量測及研究。目前每天觀測大約

1,800~2,200 個大氣層和電離層剖面資料點，提供給氣象操作中心和科學研

究團隊進行氣象預報及分析用。經過全球氣象單位的資料評估及驗證，福

衛三號對目前運作中的全球氣象預報模式及颱風及颶風軌跡路徑預測產生

正面的影響，並可用以監測全球氣候變遷。利用先進的開迴路技術，福衛

三號比之前的 CHAMP 任務所提供的 RO 資料，更深入穿透到對流層以下

以探測大氣層的變化。由於福衛三號的優異科學成就，後續任務將進一步

由實驗型轉換成作業型的任務，並計畫同時接收 GPS/GALILEO/GLONASS

系統的資料。本博士論文論述福衛三號星系任務的無線電掩星理論、星系

部署原理、升軌操作技術、星系操作結果及所面臨的操作挑戰、及如何利

用先進進動理論完成世界上第一個星系部署系統的寶貴操作經驗及成果，

並敘述後續任務的任務分析及攜帶 GNSS RO 量測儀酬載的衛星概念設計。

FORMOSAT-3 Constellation Performance, Deployment Challenges,

and Prospect for Atmospheric Remote Sensing

Student：Chen-Joe

Fong

Advisor：Sien Chi

Institute of Electro-Optical Engineering

National Chiao Tung University

ABSTRACT

The FORMOSAT-3/COSMIC (FORMOsa SATellite mission-3/Constellation

Observing System for Meteorology, Ionosphere, and Climate) satellites were

successfully launched in California on April 15, 2006 into a 516 km orbit plane.

The FORMOSAT-3 mission consisting of six low-earth-orbiting satellites is the

world’s first demonstration of near real-time operational Global Positioning

System (GPS) radio occultation (RO) mission for global weather monitoring.

After six spacecraft bus in-orbit checkout activities were completed, the mission

was started immediately at the parking orbit for in-orbit checkout, calibration,

and experiment of three onboard payload instruments: GPS occultation receiver

(GOX), Tiny Ionospheric Photometer (TIP), and Tri-Band Beacon (TBB).

Individual spacecraft was then maneuvered into six separate orbit planes of ~800

km with evenly distributed global coverage. FORMOSAT-3 mission has

verified a novel “proof-of-concept” way of performing constellation deployment

by taking the advantage of nodal precession. The received RO data have been

processed into 1,800 to 2,200 good atmospheric and ionospheric profiles per day,

respectively. The processed atmospheric RO data have been assimilated into

Numerical Weather Prediction (NWP) model for near real-time weather

prediction and typhoon/hurricane/cyclone forecasting by global weather centers

which have shown significant positive impact

With the advent of the open-loop

technique, the quality, the accuracy and the lowest penetration altitude of the RO

sounding profiles are better than CHAMP data. Due to the great success of this

innovative FORMOST-3 mission, the goal of the follow-on mission is to

transfer FORMOSAT-3 mission from research to operational with GPS, Galileo,

and GLONASS tracking capabilities. In this dissertation we

present the Global

Navigation Satellite Systems (GNSS) RO theory, the constellation deployment

theory, the constellation deployment results, the mission challenges, and the

lessons learned. We also present the spacecraft system performance, the

follow-on mission trade analysis results, and new spacecraft constellation

system conceptual design with a next-generation GNSS RO receiver onboard.

誌 謝

(Acknowledgements)

在這本論文完稿之際，首先我必須感謝我的指導教授－祁甡老師，若不

是他的悉心指導及論文建議方向，我想我不會走完最後這個階段。我也要

謝謝溫盛發博士，曾經協助祁老師來指導我通過博士讀書計畫資格考。我

也要謝謝國家太空中心的長官及指導並協助我以福衛三號的多項研究工作

為方向一起完成研究的同事們，他們是李羅權前主任、陳紹興副主任、吳

岸明副主任、陳正興副主任、陳正一前總計畫主持人、顏隆政總計畫主持

人、劉說安主任、陳彥升組長、劉肩吾博士、朱崇惠、楊善國、蕭文宗、

林辰宗、郭添全博士、黃成勇博士…等。福衛三號計畫是我國第三個衛星

計畫，是建立全球大氣即時觀測網之先進技術發展計畫。它的成功發射，

並且完成任務軌道部署，及提供了寶貴的衛星實時科學資料，建立了很多

個世界第一。

令人高興及感到欣慰的是國家太空中心三號計畫團隊於 2008 年參加國

家實驗研究院傑出科技貢獻獎的選拔，以「福爾摩沙衛星三號星系計畫的

科技成效與成就」研究主題拿下科技服務類的秀姑巒山獎殊榮。更證實此

計畫的成功背後是一個團隊合作的成功典範。2009 年 2 月所舉辦的「福衛

三號成效評估報告暨後續計畫規劃」審查，更是獲得國研院審查委員的一

致贊揚與好評。

我在完成所負責的華衛一號的整測及發射工作後，決定在十幾年前離開

校園後，又回到校園去追求更高的學術研究，實在是我一個人生旅程的另

一個轉捩點。整個論文研究從開始到更換題目到完成，證明了要在很短的

時間內，我要能在事業工作、學業研究、家庭經營、小孩教育及身體健康

上，都要能夠全部同時兼顧，是要我去完成一件不可能的任務。無論如何，

這將是我人生歷程至生難忘的一段回憶。

我要感謝國家太空中心和國家科學委員會對本任務的資助。謝謝國科

會、國研院及太空中心長官同事們對我的持續支持與關注。過去十年有很

多人對福衛三號任務的貢獻，特別是對我的論文研究工作，還有太多人以

至於我在此無法一一提及。我最後要謝謝我的家人及兩個可愛的小孩－俐

婷及偉丞，謝謝 Jo 提供我一個寫作論文的寧靜環境地方，他們一步一步陪

伴我走過這個艱辛的歷程。

Table of Contents

Chinese Abstract……….……….………..…………i

English Abstract……….………….….……….ii

Acknowledgements……...iv

Table of Contents……….……….v

List of Table………..………..……….vi

List of Figures………..………...…vii

Nomenclature………...………...….ix

Chapter 1 Introduction ... 1

1.1 History

of

Occultation... 1

1.2 GNSS

Radio

Occultation ... 2

1.3 FORMOSAT-3 Mission... 3

1.4 F3

System... 5

1.5 F3

Follow-on Mission... 6

Chapter 2 Radio Occultation Theory and Constellation Deployment Principle. 11

2.1 Introduction ... 11

2.2

The GNSS Radio Occultation Theory ... 11

2.3 Constellation

Deployment Principle ... 15

2.4 Conclusion ... 17

Chapter 3 Constellation Deployment ... 21

3.1 Introduction ... 21

3.2

Spacecraft System for Orbit Raising, and Flight Dynamics... 21

3.3 Constellation

Deployment Plan Evolution ... 26

3.4 Constellation

Deployment Results... 27

3.5 Conclusion ... 29

Chapter 4 Challenges of Constellation Mission Operations ... 42

4.1 Introduction ... 42

4.2 Constellation

Mission Operation ... 42

4.3 Constellation

Operations Challenges... 45

4.4 Payload

Operation Challenges ... 48

4.5 Constellation

Deployment Challenges... 51

Chapter 5 Constellation Spacecraft System Performance ... 61

5.1 Introduction ... 61

5.2

Constellation Spacecraft SystemPerformance Summary ... 61

5.3

Spacecraft Subsystem On-Orbit Performance Summary ... 62

5.4

GOX Payload Science Performance Results ... 64

5.5 Conclusion ... 66

Chapter 6 Follow-On Mission Trade Analysis and Design ... 77

6.1 Introduction ... 77

6.2 Follow-On

Mission

Definition Trade Analysis Results ... 77

6.3 Follow-On

Mission

System

Architecture and System Design ... 82

6.4 Conclusion ... 84

Chapter 7 Conclusions ... 94

Reference………..……….……..……95

Appendix Acronyms and Abbreviations………..………….…104

List of Tables

Table 1-1 The F3 Mission Characteristics... 8 Table 3-1 F3 Constellation Spacecraft Bus Key Design ... 30 Table 3-2 Constellation Deployment Status With Five Satellites (FM5, F M2, FM6, FM4,

and FM1) At Final Orbits as-of-2 Dec, 2007 ... 31 Table 3-3 Spacecraft Thrust-Burn Performance Statistics ... 32 Table 3-4 Spacecraft Mass Property and Moment of Inertia for Six Satellites as-of-2 Dec,

2007 ... 33 Table 5-1 Constellation Spacecraft Performance Summary (After Two Years in Orbit) ... 67 Table 5-2 Spacecraft Operation Status of Each Subsystem in All Six Spacecraft (After 2

years in Orbit) ... 68 Table 5-3 Spacecraft Subsystem Performance (After 2 years in Orbit) ... 69 Table 5-4 GOX Firmware Build (FB) Change History since Launch... 70 Table 6-1 Expected Atmospheric Profiles vs. Different Constellation and Different

Receiver Capability. ... 85 Table 6-2 Proposed Follow-On Mission Spacecraft Bus Design vs. F3 Design... 86

List of Figures

Figure 1-1 Schematic diagram illustrating radio occultation of GNSS signals... 9

Figure 1-2 F3 system architecture. ... 10

Figure 2-1 GNSS RO receiver operation concept. ... 18

Figure 2-2 Basic GNSS RO measurements and processing flow. ... 19

Figure 2-3 Ray path geometry from point G to point L in the plane of propagation. For a spherical symmetric medium a=aG =aL... 20

Figure 3-1 F3 spacecraft in deployed configuration and its major components. ...34

Figure 3-2 Spacecraft Reaction Control Subsystem block diagram... 35

Figure 3-3 Reaction Control Subsystem thruster geometry and torque. ... 36

Figure 3-4 Reaction Control Subsystem blowdown curve... 37

Figure 3-5 Functional block diagram of the spacecraft attitude control subsystem...38

Figure 3-6 Off-pulsing concept of ACS thrust mode. ...39

Figure 3-7 F3 as-is burn history and deployment timeline. ... 40

Figure 3-7. Spacecraft thrust-burn performance statistics... 41

Figure 4-2 Six spacecraft separation simulation result. ... 54

Figure 4-3 F3 final constellation. ... 55

Figure 4-4 GPS three-dimensional (3D) tracking coverage of all six spacecraft Bus GPSR.. 56

Figure 4-5 Number of GPS satellite vehicle tracked statistics for all six spacecraft bus GPSRs of one-year data after launch... 57

Figure 4-6 Geographic location of the spacecraft resets/reboots events two years since launch... 58

Figure 4-7 One-Year Trend of Solar Power and Battery SOC, ACS Mode, and Payload On-Off Status on Spacecraft FM2. ... 59

Figure 4-8 Payload (GOX/TBB/TIP) power-on/off Statistics. ...60

Figure 5-1 The six satellites attitude on-orbit performance with respect to the sun beta angle for one-year data since launch. ... 71

Figure 5-2 Trending plots of the tank pressures and temperatures for FM2, FM4, FM5, and FM6 (from 15 April 2006 to 15 April 2007) ... 72

Figure 5-4 Two Years Statistics of the Number of Daily Occultation Events for

Atmosphere Profiles since Launch. ... 74 Figure 5-5 Two Years Statistics of the Number of Daily Occultation Events for

Ionosphere Profiles of Electron Density since Launch. ... 75 Figure 5-6 Comparison of the lowest altitude penetration of RO event versus latitude for

F3/COSMIC and CHAMP...76 Figure 6-1 The relationship between total occultation number and inclination angle for one

satellite receiving GPS only. ... 87 Figure 6-2 The dependence of data distribution vs. latitude for a 72o inclination angle. The

“equivalent area covered by one occultation” is defined as the average area in square km associated with a single sounding. e.g., one sounding per N km (x N km)... 88 Figure 6-3 The dependence of data distribution with inclination angle. The “equivalent

area covered by one occultation” is defined as the average area in square km associated with a single sounding. e.g., one sounding per N km (x N km)... 89 Figure 6-4 The F3 follow-on constellation with 12 satellites. ... 90 Figure.6-5 6-hr Occultation Distribution with 12-satellite constellation for the F3

follow-on mission (the blue dots are from GPS, the green dots are from

GALILEO, and the purple dots are from GLONASS) ... 91 Figure 6-6 The F3 follow-on mission system architecture with constellation of 12 satellites.

Nomenclature

α2 = Bending angle of L2 frequency

AC/A = Received power of the in-phase component of the L1 signal AP1 = Received power of the quadrature component of the L2 signal AP2 = Received power of the L2 signal

aSMA = Semi-Major Axis of the Orbit Altitude in km

ΔaSMA = Maximal difference in SMA (in meter) b = Impact Parameter

CA(t) = Clear acquisition (C/A) code-modulating the in-phase component of L1 signal at a rate of 1.023 MHz

E = Eccentricity

f = Frequency of Global Positioning System Carrier Signal in Hz

fD = Excess Doppler frequency shift measured by the GNSS receiver of LEO F = Thrust force

I = Inclination Isp = Specific Impulse

ΔL = Maximal deviation Argument of Latitude in degree

λ = Wavelength of the harmonic wave

M(t) = Amplitude modulation of L1 and L2 containing navigation data N = Refractivity

n = Index of Refraction

nG = Index of refraction at the occulted GNSS satellite nL = Index of refraction at the LEO satellite

ne = Electron Density in Number of Electrons per Cubic Meter

ΔΩ = Drift of the RAAN after a deployment time

Δφ = Phase delay P = Pressure in hPa Pm = Propellant Mass

Pw = Water Vapor Pressure in hPa

PY(t) = Precision (P) code-modulating the in-phase component of L1 and L2 signals at a rate of 10.23 MHz

r = Position along the raypath

rG = Geocentric position vector to the occulting GNSS satellite rL = Geocentric position vector to the LEO satellite

rLG = Geometric straight line distance between the LEO satellite and the occulted GNSS satellite s = Arc length along the ray path

σ = Standard Deviation T = Temperature in Kelvin

TG = Ray path tangent vectors of the occulted GNSS satellite TL = Ray path tangent vectors of the LEO satellite

t = Deployment Time Period in days VG = Velocity of the occulted GNSS satellite VL = Velocity of the LEO satellite

Chapter 1 Introduction

1.1 History of Occultation

The term “occultation” is widely used in astronomy when an object in the foreground occults (covers up) objects in the background, and it refers to a geometry involving the emitter, the planet and its atmosphere if any, and the receiver changes with times.1 The first scientific application of the occultation technique was introduced in the eighteen century when it was used for timing astronomical events. By observing scintillations, refraction, and variations in stellar brightness and spectra when a star is occulted by a planet or moon, the spectral intensity fading could be used to approximate the scale height of planetary atmosphere by using the geometric ray optics theory [1].

Radio occultation (RO) is a remote sensing sounding technique in which a microwave emitted from a spacecraft passes through an intervening planetary atmosphere before arriving at the receiver, and is used to study the physical properties of planetary atmosphere in the early days of interplanetary mission [2]. The atmospheric radio 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 [3]. Mariner-4, the first spacecraft to Mars (in 1964), flew along a spacecraft trajectory that passed behind Mars when viewed from Earth [4]. When Mariner-4 spacecraft 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 a very first valuable atmospheric and ionospheric density information by using the inversion techniques derived from basic geometric ray optics theory, Fourier optics theory, and Maxwell’s electromagnetic wave theory [5]. Mariner-4 opens an era of planetary RO

1

experiments. Since then a series of planetary experimental missions were undertaken to study the atmospheres and ionospheres of the planets and their moons, as well as certain physical properties of planetary surfaces and planetary rings [6].

1.2 GNSS Radio Occultation

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 [7]. A few early RO experiments from a satellite-to-satellite tracking link had been conducted. These included the occulted radio link between ATS-6 and GEOS-3 [8] and between the Mir station and a geostationary satellite [9].

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 [10]. 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 modeling [11]-[12]. However, one of the major limitations to the ground-based GNSS remote sensing is that it just only provides integrated PW with little useful vertical resolution, and it is restricted to land areas filled with GNSS networks. The space-based GNSS atmospheric limb sounding offers a complementary solution to these issues [13].

The space-based GNSS RO atmospheric remote sensing technique, which makes use of the 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 [14]-[17]. Figure 1-1 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 measurement of the atmosphere by satellites and became the first space-based “proof-of-concept” demonstration of GNSS RO mission to Earth [18]-[23]. For a more complete history of GNSS RO see Melbourne et al. in [5] and Yunck et al. in [6].

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 Gravity Recovery and Climate Experiment (GRACE) missions (in 2002). The GPS RO sounding data have been shown to be of high accuracy and high vertical resolution. 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 [19]-[24].2

1.3 FORMOSAT-3 Mission

The F3 mission is the world’s first demonstration of GPS radio occultation near real-time operational constellation mission for global weather monitoring. The primary scientific goal of the F3 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 F3 mission provides about 1,800 ~ 2,200 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 [25]-[33]. The mission results have shown that the RO data from F3 are of better quality than those from previous missions and penetrate much further down into the

2 _{In this dissertation we refer to the FORMOSAT-3/COSMIC mission as F3 mission for}

troposphere, the mission results could be referenced to Cheng et al. in [28], Liou et al. in [29], Anthes et al. in [30], Fong et al. in [31] and [32], and Huang et al. in [33]. In the near future, other GNSS, such as the Russian Global Navigation Satellite System (GLONASS), and the planned European Galileo system, will be used to extend the region of applications by the use of GPS RO technique [32], [34]-[36].

Table 1-1 shows the F3 mission characteristics. The F3 mission was launched successfully from Vandenberg Air Force Base in California 1:40 UTC on April 15, 2006 into the same orbit plane of the designated 516 km circular parking orbit altitude. The F3 mission is jointly developed by Taiwan’s National Space Organization (NSPO) and United State’ University Corporation for Atmospheric Research (UCAR) in collaboration with Orbital Sciences Corporation (OSC or Orbital) for the satellites, NASA’s Jet Propulsion Laboratory (JPL) and Naval Research Laboratory (NRL) for three onboard payloads including a GPS Occultation Receiver (GOX), a Tri-Band Beacon (TBB), and a Tiny Ionospheric Photometer (TIP). The TIP payload instrument is routinely collecting data at night, and observes the equatorial anomaly arcs and other density anomalies through measurements of 1356 Angstrom radiation. The nadir-pointing TBB enables observations of the line-of-sight total electron contents (TEC) and scintillations along the F3/COSMIC-TBB ground stations’ radio links. The data from these two instruments complement the ionospheric observations from the GOX and are used to improve the retrieval of electron density profiles at night and over TBB ground stations. These data are also valuable for evaluation of ionospheric models and use in space weather data assimilation systems [30].

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 [30], [37]. The great success of the F3 mission expected to operate through 2011, has initiated a new era for near real-time operational GNSS RO soundings [35]-[38].

1.4 F3 System

The F3 constellation system architecture consists of the six identical on-orbit micro-satellites, Spacecraft Operations Control Center (SOCC) in Taiwan, several TT&C (telemetry, tracking and command) Ground Stations, and two data receiving and processing centers, and the fiducial network. There are two TT&C local tracking stations (LTS), one located in Chungli and the other in Tainan of Taiwan, respectively. There are two remote tracking stations (RTS) to support the passes. Originally one is located at Fairbanks, Alaska and the other one is located at Kiruna, Sweden. After two years in orbit operation, the F3 program switches from these two ground stations to two new ground stations in Fairbanks (FBK), Alaska, and Tromso (TRO), Norway, plus a third RTS located in McMurdo, Antarctica. This McMurdo ground station is expected to reduce the data latency of some RO products. These three RTS are currently set as primary stations for the F3 mission. Figure 1-2 shows the F3 system architecture [32], [39],

The SOCC uses the real-time telemetry and the back orbit telemetry to monitor, control, and manage the spacecraft state-of-health (SOH). The downlinked science RO data is transmitted from the RTS via National Oceanic and Atmospheric Administration (NOAA) to the two Data Receiving and Processing Centers: (1) CDAAC (COSMIC Data Analysis and Archive Center) which is located at Boulder, Colorado, USA; and (2) 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 through double differencing. All collected science data is processed by CDAAC and then transferred to TACC and other facilities for science and data archival [40].

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 (JSCDA), 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), Meteo France, etc. And they are made ready for assimilation into weather prediction models. The data is currently provided to weather centers within 90 minutes (data latency requirement is 180 minutes) after satellite on-orbit science data collection in order to be ingested by the operational weather forecast model [36].

1.5 F3 Follow-on Mission

As addressed in the Final Report of “Workshop on the Redesign and Optimization of the Space Based Global Observing System,” the World Meteorological Organization (WMO) had recommended continuing RO observations operationally and the scientific community had urged continuation of the current mission and planning for a follow-on operational mission [41]. The proposed follow-on mission is a greatly improved operational and research mission with redundancy and robustness and consisting of a new constellation of 12 satellites. The need mission will seek to establish international standards so that future RO missions deployed by any country can be assimilated into the same systems. The primary payload of the follow-on satellite will be equipped with the GNSS RO receiver and will collect more soundings per receiver by adding European GALILEO system and Russian’s Global Navigation Satellite System (GLONASS) tracking capability, which will produce a significantly higher spatial and temporal density of profiles. These will be much more useful for weather prediction models and also severe weather forecasting including typhoons and hurricane, as well as for a research [36].

In this dissertation we provide an overview of the radio occultation theory, new constellation deployment theory, the constellation spacecraft design, the constellation mission

operations, the orbit-raising challenges, and the lessons learned during the orbit-raising operations. We also present the F3 satellite constellation system performance, and the prospect of a future follow-on mission with the performance enhancements we have accomplished.

TABLE 1-1 THE F3MISSION CHARACTERISTICS

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

Power ~ 81 W orbit average

Communication S-band uplink (32 kbps) and downlink (2 Mbps) Sounding ~2000 soundings per day

Data Latency 15 minutes to 3 hours Design and Mission life 5 years

Figure 1-1. Schematic diagram illustrating radio occultation of GNSS signals. GNSS Atmosphere Ionosphere LEO GNSS Signal Tangent point

a

a

r

α

a

a

r

α

Figure 1-2. F3 system architecture. T_B B ch ain_s Satellite Operations Control Center (SOCC), Taiwan Fiducial Network FORMOSAT-3 L1&L2 GPS TT&C, Taiwan Users TBB signal Command & telemetry

NOAA

CDAAC/UCAR Boulder,CO

GPS

Payload science data

L1&L2 Remote Terminal Stations (RTS) at Fairbanks/Tromso/McMurdo Wallops/Svalbard TACC/CWB Taipei Users T_B B ch ain_s Satellite Operations Control Center (SOCC), Taiwan Fiducial Network FORMOSAT-3 L1&L2 GPS TT&C, Taiwan Users Users TBB signal Command & telemetry

NOAA NOAA

CDAAC/UCAR Boulder,CO

GPS

Payload science data

L1&L2 Remote Terminal Stations (RTS) at Fairbanks/Tromso/McMurdo Wallops/Svalbard TACC/CWB Taipei Users Users

Chapter 2 Radio Occultation Theory and Constellation

Deployment Principle

2.1 Introduction

This Chapter begins with an overview of the GNSS radio occultation theory (in Section 2.2) and followed by the constellation deployment principle (in Section 2.3). In Section 2.2 we present the GNSS, GNSS radio occultation theory and operation concept; and radio occultation data retrieval theory. As for Section 2.3, we present earth oblateness right ascension ascending node phasing, argument of latitude, final phasing, contact conflict avoidance, and dispersion operation to maximize science data downloads, followed by the conclusion.

2.2 The GNSS Radio Occultation Theory

2.2.1 The Global Navigation Satellite System

The GPS developed by United States, is the only fully functional GNSS in the world. It consists of 24 satellites, with a few more satellites for backup, distributed in six circular orbit planes about the globe with an inclination angle of ~55o, a period of 12 hours and an altitude of 20,200 km. Although originally designed as a navigation aid by the U.S. Air Forces, the ground-based and the space-based applications of the GNSS remote sensing have shown positive impacts on climate monitoring, global and regional weather prediction, ionospheric research, and space weather forecasting.

Each GPS satellite continuously transmits right-hand circularly polarized signals at L1 and L2 band frequencies. The L1 and L2 signals received from each GPS satellite can be written as [3]: ) 2 sin( ) ( ) ( 2 ) 2 cos( ) ( ) ( 2 ) (t = A M t P t πf t+θ + A M t C t πft+θ S , (1)

) 2 cos( ) ( ) ( 2 ) ( 2 2 1 2 t = A M t P t πf t+θ S p Y . (2)

2.2.2 GNSS Radio Occultation Retrieval Theory

In Figure 2-1 a GNSS RO operation concept and data set for an occultation event are shown. By measuring the phase delay of radio waves from GNSS satellites as they are occulted by the Earth’s atmosphere, accurate and precise vertical profiles of the bending angles of radio wave trajectories in the ionosphere, stratosphere and troposphere are obtained. A complete GNSS RO data set for an RO event includes (1) Occultation data: signal from an occulting GNSS satellite to occulting LEO satellite with 20 msec data rate (see link 1 marked in Figure 2-1); (2) Referencing data: signal from a non-occulted GNSS satellite with 20 msec data rate (see link 2 marked in Figure 2-1); (3) Precision orbit determination (POD) data: signals from other three non-occulted GNSS satellites with 10 sec data rate; and (4) Fiducial IGS (International GNSS Service) data: GNSS navigation data from ground fiducial network sites with 1sec data rate from occulting GNSS satellite (see link 3 and link 4 marked in Figure 2-1) [39]-[40].

A basic GNSS RO measurements and processing flow is presented in Figure 2-2. We derive the single path GNSS RO theory in this Section. From the calculus of variation the ray path from the GNSS satellite to the LEO satellite, in a geometric optics context, is by definition a path of stationary path and satisfies Fermat’s principle globally and Snell’s law locally [5], [42]. Figure 2-2 we show a ray path geometry from a occulted GNSS satellite (point G) to a LEO satellite (point L) in the plane of propagation and illustrating radio occultation of GNSS signals. This ray must satisfy the requirement

= − ′ + = − − = ∆ = ∆

∫

∫

whereΔρis the ray delay, Δφ is the phase delay, n(r) is the real part of the refractive index, r is the geocentric position vector of any point on the ray, s is the arc length along the

ray path, rL is the geocentric position vector to the LEO satellite, r is the geocentric G

position vector to the occulting GNSS satellite, and r is the geometric straight line distance _LG

between the LEO satellite and the occulted GNSS satellite.

From Figure 2-2, the excess Doppler from the intervening medium can be derived as

LG G L G L G G G L L L D r r r V V V T n V T n f = ⋅ − ⋅ −( − )⋅( − ) λ (4)

where fD =(dϕ/dt)/2π is the excess Doppler frequency shift measured by the GNSS

receiver of LEO; λ is the wavelength of the harmonic wave; n_L and n are the index of _G

refraction at the LEO and occulted GNSS satellites and is equal to unity, respectively; TL

and T_G are the ray path tangent vectors of the LEO and occulted GNSS satellites, respectively; and VL and VG are the velocity of the LEO and occulted GNSS satellites,

respectively. The triangle OLG defines the instantaneous plane of propagation of the ray from the occulted GNSS satellite to the LEO satellite. The interior angles of this triangle OLG and its sides are completely determined from the precision orbit determination (POD) information about the orbits of the LEO and occulted GNSS satellite. The refraction-related quantities, which are the bending angleα =δ_L +δ_G, can be determined from the excess Doppler measurement of Eq. (4) by applying a=nr×T =constant, which is Bouguer’s law, essentially a Snell’s law for a spherical symmetric medium.

As the ionosphere is considered as a source of concentration of electrons and the frequency of electromagnetic wave, the L1 and L2 GNSS signals can be combined to significantly reduce the effect of the ionosphere. The atmospheric bending angle can be calculated using Eq. (5) below 2 2 2 2 2 2 1 2 1 ( ) ( ) ) ( f f r f r f r L − − = α α α (5)

From the bending angles, profiles of atmospheric index of refraction are obtained through the equation of Abel transformation as [3], [42]:

⎥ ⎥ ⎦ ⎤ ⎢ ⎢ ⎣ ⎡ − =

_∫

where n(ap)is the refractive index at ap, ap =nrp is the impact parameter for the ray at

perigee, and rp is the altitude of perigee,α(a)is the bending angle at a.

In the atmosphere, the index of refraction, n, is very close to unity such that it is usually discussed in terms of the refractivity, N. By using Eq. (7) N is a function of temperature (T in K), pressure (P in hPa), water vapor pressure (Pw in hPa), electron density (ne in number of

electrons per cubic meter), and frequency of the GPS carrier signal (f in Hz) as

2 6 2 5 6 10 3 . 40 10 73 . 3 6 . 77 10 ) 1 ( f n T P T P n N = − × = + × w − × e . (7)

The refractivity profiles can be used to derive profiles of electron density in the ionosphere, temperature in the stratosphere, and temperature and water vapor in the troposphere by using Eq. (7).

For problems from multipath, there have been several data processing methods for RO data inversion to retrieve atmospheric parameters from a wave optics theory treatment [5], As for the F3 mission, Kuo et al. develop a RO data processing procedures used to obtain stratospheric and tropospheric bending angle and refractivity profiles from the raw phase and amplitude data [23], [37]. The Phase Lock Loop (PLL) technique employed in earlier RO missions was replaced by a novel open loop technique for the F3 mission [43]-[45]. There are other data processing procedures or algorithms developed by other methods [5], such as the geometrical optics method (GOM) [46]-[47], the back-propagation method (BPM) [48]-[49], the radio holographic method (RHM) [50]-[51], the amplitude-retrieval method (ARM) [52], the full-spectrum-inversion method (FSIM) [53], the canonical transformation

method (CTM) [54], the sliding spectral (or radio optics) method (SSM) [44]-[45] and National Central University Radio Occultation (NCURO) algorithms [55]-[56].

The F3 RO processing includes four radio holographic algorithms: BPM, SSM, CTM, and FSIM. Detailed description and derivations of F3 RO data processing procedure could refer to Kuo et al. in [23]. The RO data processing procedure and steps currently used for F3 mission are listed as follows:

1. Input (Phase, amplitude, LEO/GPS position and velocity);

2. Open-loop data processing GNSS navigation data messages (NDM) removal and phase correction;

3 Detection of L1 phase locked loop tracking errors and truncation of the signal; 4: Filtering of raw L1 and L2 Doppler;

5. Estimation of the “occultation point”

6. Transfer of the reference frame to the local center of Earth’s curvature; 7. Calculation of L1 and L2 bending angles from the filtered Doppler; 8. Calculation of the bending angles from L1 raw complex signal; 9. Combining (sewing) L1 bending angle profiles from steps 7 and 8; 10. Ionospheric calibration of the bending angle;

11. Optimal estimation of the bending angle; 12. Retrieval of refractivity by Abel inversion; 13. Retrieval of pressure and temperature;

14. Output (bending angle, refractivity, pressure, temperature, moisture).

2.3 Constellation Deployment Principle

2.3.1 Earth Oblateness Right Ascension Ascending Node (RAAN) Phasing

The total mass of a F3 satellite is 61.05 kg, including the dry mass of 54.4 kg and the propellant mass of 6.65 kg. And the overall altitude increase from injection orbit to mission

orbit is 285 km. The estimated total delta-V required is 147 m/s, and the estimated propellant required is 4.6 kg. Fuel margin is 2.05 kg [57]-[58].

Due to the oblateness of the Earth gravity, the RAAN (Ω) of a LEO satellite will drift away at a rate. The drift rate of RAAN (∆Ω/∆t), also called “orbit precession rate,” which is a function of the Semi-Major Axis (SMA), inclination, and eccentricity of the orbit. For the F3 near-circular orbit with an inclination of 72o and eccentricity of 0, the orbit precession rate is modeled as an equation below [59]:

t a_SMA ⋅∆ ∆ × − ≅ ∆Ω − ) ( 10 3804 . 6 13 7/2 (8) where

∆Ω the drift of the RAAN after a deployment time of ∆t;

SMA

a the SMA of the orbit altitude in km;

t

∆ the deployment time period in day.

The deployment strategy is to use the first raised spacecraft (FM5) as a reference point. The second spacecraft is then raised to its mission orbit when the difference of the RAAN between the first and the second spacecraft reaches the desired separation angle, and so forth.

2.3.2 Argument of Latitude (AOL) Final Phasing and Contact Conflict Avoidance

As one ground station can support one pass from elevation angle 10o to 10o, if there are two satellites flying over the same ground station at the same time frame, the ground station could support only one satellite unless there were special arrangements. Therefore, a 52.5o phasing on AOL must be implemented to ensure that one orbit’s worth of occultation science data are sent to the receiving stations. The maximal difference in SMA (ΔaSMA in meter)

and the maximal deviation (ΔL in degree) of the AOL from its nominal value are deployed to fulfill the following equation

ΔaSMA +5*ΔL < 50. (9)

The differentiation of the AOL of the other five satellites against the reference orbit is achieved by controlling the altitude deployment profile in the final stage of the “maneuvering window.” When the orbit altitude is different from the reference orbit (FM5), the AOL change rate is also different from the reference orbit. The different AOL change rate differentiates the AOL of the satellite against the reference orbit along with time. By manipulating the altitude deployment profile in the final stage, the AOL difference is targeted at the same time to maneuver the satellite into the mission orbit altitude. Then both the RAAN and AOL differences are frozen and kept constant simultaneously.

2.3.3 Dispersion Operation to Maximize Science Data Downloads

The dispersion operation is very similar to the AOL phasing. In order to increase the number of GOX data downlink, a spacecraft dispersion operation plan was executed to differentiate the AOL of FM4, FM3, FM1 and FM6 in parking orbits. These four satellites were maneuvered to the same altitude around 519 km with an AOL difference around 80o so that they can contact a ground station in turn to increase GOX science data downlink with no contact conflicts [57]-[58].

2.4 Conclusion

In this Chapter we have given an overview of the GNSS radio occultation theory and the constellation deployment theory. The constellation deployment theory is used for unique F3 constellation deployment and the results are presented in Chapter 3.

Figure 2-1. GNSS RO receiver operation concept. POD Ref. Occ. Occulting LEO Occulting GNSS Calibrating GNSS Ground receiver 1-se c d ata _(L IN_K 4) 20 m sec da ta (L INK 2) 1-se c d ata (LIN K 3 ) 20 msec d_ata Ionosphere Neutral atmosphere

Earth

Earth

Earth

Figure 2-2. Basic GNSS RO measurements and processing flow. L1 & L2 Phase and Amplitude L1 & L2 Bending Angle

LEO Satellites Orbits & Spherical Symmetry

Radio Holographic Methods, Multipath

Iono-free Bending Angle

L1 & L2 Phase Geometric Optics Method,

Single Path Ionospheric Correction Refractivity (N) Pressure, Temperature, Moisture

High Attitude Climatology & Abel Inversion

Auxiliary Meteorological Data L1 & L2

Phase and Amplitude

L1 & L2 Bending Angle

LEO Satellites Orbits & Spherical Symmetry

Radio Holographic Methods, Multipath

Iono-free Bending Angle

L1 & L2 Phase Geometric Optics Method,

Single Path Ionospheric Correction Refractivity (N) Pressure, Temperature, Moisture

High Attitude Climatology & Abel Inversion

Figure 2-3. Ray path geometry from point G to point L in the plane of propagation. For a spherical symmetric mediuma=a_G =a_L.

G L Electromagnetic Signal α O b

rˆ

θ

ˆ

χ

δ

χ

δ

θ

θ

π

θ

ˆ

rˆ

rˆ

θ

ˆ

χ

δ

χ

δ

θ

θ

π

θ

ˆ

rˆ

Chapter 3 Constellation Deployment

3.1 Introduction

The F3 mission operation concept is to launch the entire cluster of satellites by a single launch vehicle. All six satellites are delivered to the same injection orbit plane of a designated 516-km circular parking orbit altitude, and the six satellites are in a cluster formation fly configuration after separation from the launch vehicle. They are then deployed into six different orbit planes at specific time intervals using the constellation deployment principle[57]-[58], [61].

The F3 mission takes advantage of nodal precession to conduct orbit-raising maneuvers at the appropriate times so that the effect of different altitudes makes the orbital planes drift [62]. It is well-known that the nodal precession is a gravity phenomenon where the orbital plane drifts due to the Earth’s oblateness. The approach using the natural physics of the Earth’s oblateness, as well as time, allows the spacecraft to drift instead of requiring complex propulsion systems or even depending on individual launch vehicle to arrive at their orbit planes directly. Although this approach requires a lengthy orbit-deployment time, it significantly reduces the size of the propulsion subsystem design needed [31].

The F3 spacecraft systems for orbit raising and ground flight dynamics design are presented in Section 3.2 below. We present the evolution of the constellation plan in Section 3.3, the constellation deployment results in Section 3.4, and followed by the conclusion in Section 3.5.

3.2 Spacecraft System for Orbit Raising and Flight Dynamics

3.2.1 F3 Spacecraft System

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 [32], [61]. Table 3-1 shows the F3 constellation spacecraft bus key design features.

3.2.2 Spacecraft Propulsion for Thrust Burn

The spacecraft propulsion subsystem (also named the RCS) is a blowdown monopropellant Hydrazine (N2H4) Propulsion Subsystem with gas-helium (GHe) as the

pressurant. And the designed blowdown ratio is 5:1 with the MEOP (Maximum Expected Operating Pressure) of 400 psia at 50ºC. The initial tank pressure is pressurized to about 330 psia at 20oC. We utilize the RCS to provide impulses for attitude control during orbit-raising and to transfer the satellite from the injection orbit to an intermediate orbit if required, and finally to the mission orbit of the constellation. Figure 3-2 shows the block diagram of the RCS. For F3 spacecraft system the RCS consists of a propellant tank, gaseous helium and Hydrazine service valves, a latching valve, a filter, an orifice, four thrusters, pressure transducer, and a set of pipelines. The spacecraft RCS characteristics are summarized as follows [57]-[58]:

–Thrust Force: 1.1 [Beginning of Life (BOL)]－0.2 N [End of Life (EOL)]; –Specific Impulse: 217－194 s;

–Propellant Mass: ~6.65 kg;

–Thrust Type: OFF pulsing (Duty Cycle ≦ 50%).

the four quadrants of the x-z plane of the satellites. These four thrusters are canted by 10° to enable three-axis control capability. By modulating the off-pulsing duration of the four thrusters, control torque is generated for the attitude control around X, Y, and Z axis of the satellite. The estimated thrust and specific impulse over the entire blowdown pressure range are shown in Figure 3-4.

3.2.3 Spacecraft Attitude Control for Orbit Raising

The function of the spacecraft ACS is to control the attitude of the satellite in the Safe Mode, the Stabilization Mode, the Nadir Mode, the Nadir-Yaw Mode, and the Thrust Mode. And the ACS sensors for attitude estimation include Earth horizon sensors, coarse sun sensors, and a magnetometer. The ACS actuators for attitude control include magnetic torquers, a reaction wheel and thrusters [57], [61].

Figure 3-5 shows the functional block diagram of the spacecraft ACS where FC stands for Flight Computer and ACE means the Attitude Control Electronics. In Figure 3-5 the Attitude Reference System (ARS) includes attitude and rate estimators using a Kalman filter algorithm with measurements from the sensors. The ACS Controller processes the attitude and rate estimation from ARS through the control gains/algorithm, and distributes the torque commands to the actuators. The ACS also receives the satellite position and velocity data from the bus GPS receiver (GPSR). Based on this information it then propagates and computes necessary information for the navigation purpose, the ARS and the commanded angles for the Solar Array Drive (SAD).

The Thrust Mode is dedicated to the orbit-raising operation. When the orbit-raising operation is performed, the satellite first maneuvers itself to a yaw angle of 90° to align the thrust direction with the velocity direction. Then, as soon as the ACS enters Thrust Mode the thruster ignition starts up, the attitude is controlled by thrusters while orbit-raising proceeds. When the operation is terminated or finished, the ACS enters the Nadir-Yaw

Mode and maneuvers itself to a pre-set yaw angle.

A proportional-integral-derivative (PID) controller is designed for the Thrust Mode to compute the desired 3-axis control torque. Four thrusters are commanded off-pulsing in each control cycle to provide both the impulse for orbit raising and the 3-axis control torque to diminish the attitude errors. Figure 3-6 shows the concept of the “off-pulsing” in each control cycle. In orbit-raising operations, the thrust turn-on time in each control cycle is either kept constant as the “InitialThurstPower” value, or increased by “AddThrustIncrement” seconds in every “AddThrustInterval” control cycles. The Thrust Mode control gains are adjusted in order to compensate for changes in thrust level during the RCS blowdown process.

The PID controller will minimize the attitude control error and improve the orbit-raising performance, but it suffers from the relative instability issue. This is because the control system may diverge with a large thruster turn on time when the PID integral terms are not yet converged to their steady-state values. Therefore, during orbit-raising operations, the PID controller requires a series of “calibration burns” in order to converge the attitude integral terms and to ramp up the thruster turn-on time to a larger value. Calibration burn is usually a smaller burn than the full-thrust burn. During the calibration process, the final values of the thrust turn on time and the integral terms of a previous burn are used as the initial values for the next burn. In this way, it takes about 6~8 calibration burns to reach the so-called full-thrust burn.

3.2.4 Flight Dynamics and Orbit Dynamics

The main function of ground-based Flight Dynamics Facility (FDF) is to conduct various orbit dynamics analyses including orbit determination, orbit-ephemeris propagation, orbit-maneuver planning, orbit-parameter trending, and orbit-event prediction. In the F3 mission, we use the commercial off-the-shelf software package called “Orbit Analysis System

(OASYS)” in FDF for orbit analysis. The OASYS database includes the thrusting model of the onboard RCS and ACS, such as the thruster number, location and direction; propellant mass and pressure; pressurant mass; blowdown curves for thrust and specific impulse; and thrust type, thruster duty cycle and efficiency [57], [61].

The blowdown curves for thrust force (F) and specific impulse (Isp) as shown in Figure

3-4 are modeled as the equations:

F = (0.001141+0.0006*P)* 4.448221 (in newtons). (1)

Isp = 222.84 - 2268.4/Pm(in seconds). (2)

where

F the thrust force;

Isp the specific impulse; Pm the Propellant Mass.

and used in the OASYS database for F3 orbit raising. Both equations are functions of the propellant tank pressure in the unit of psia.

The thrust power in each ACS control cycle is modeled as the duty cycle of the thruster and listed as Duty Cycle = Thrust Power/Control Cycle. In full-thrust orbit-raising burns, the thrust power in each control cycle is kept constant, as the duty cycle is in the OASYS model. However, in calibration burns, the thrust power in each control cycle is linearly ramped up to the end of the burn. In other words, the duty cycle in each control cycle also increases in the same way as the thrust power does. Unfortunately, there is no way in OASYS to correctly model the calibration burns with increasing thrust powers. Instead, an averaged thrust power (duty cycle) using the initial and final thrust powers of the burn is used in the OASYS database to model the thrusting of a calibration burn.

The OASYS is also used to conduct an orbit determination to compare the actual post-burn orbit and the OASYS-planned post-burn orbit after a thrust-burn is completed. Based on the actual and OASYS-planned orbit altitude, a thrusting efficiency is recalculated,

which in turn provides another input for the next orbit-raising planning.

3.3 Constellation Deployment Plan Evolution

3.3.1 Original Constellation Deployment Plan

The F3 mission operation plan changes as time passes following launch. Originally the F3 constellation deployment plan included a tandem flight design during the deployment phase. The tandem flight satellites would maintain an along-track distance of 200~400 km. Two pairs (FM1&FM2, FM3&FM4) of satellites would fly in tandem in an intermediate orbit altitude (525km and 576 km) for the geodesy research. However, spacecraft FM3 and FM4 have been very close together since launch of the satellites. The data from April to October were able to provide adequate data for geodesy research at the parking orbit of 516 km. The constellation plan was thus changed to meet the need for more science dumps for Intensive Operation Period (IOP) campaign and tropical cyclone (typhoon and hurricane, etc.) prediction forecast studies [29], [31].

The constellation plan at an 800-km orbit with 24o separation planes was for a shorter deployment time consideration (13 months after launch) and based on the assumption that spacecraft attitude control performance in lower altitude is worse than that in the mission orbit. However, this plan is not favorable for the ionospheric monitoring and climate seasonal variability studies, due to non-uniform coverage globally. Shorter duration to complete the constellation deployment has become less of a concern since the spacecraft attitude performance is better than expected and the data of the early phase (mostly at lower orbit) are much better than anticipated [61].

3.3.2 New Constellation Deployment Plan

Scientists from Taiwan and the US coherently favor 30o separation with ~6 months longer constellation deployment duration over 24o separation for global uniform coverage in

manpower was reallocated, and the orbit-raising schedule was rearranged to accommodate the science team’s request. This change in new constellation plan reflects integral teamwork among the operations team and data users and leads to greater mission success. The constellation deployment plan change from the 24o separation to 30o separation was made in September 2006 after the completion of FM5 orbit transfer and during FM2 orbit raising. The decision was made to put the FM2 orbit transfer on hold in October 2006 and to allow its separation from FM5 further. The decision postponed the completion of the final constellation to December 2007 [29],[57],[61].

3.4 Constellation Deployment Results

3.4.1 As-Burn Constellation Results

The current constellation configuration as of December 2007 is five satellites (FM5, FM2, FM6, FM4, and FM1) successfully reaching the 800-km mission orbits. On August 3, 2007 FM3 encountered the solar array drive mechanism malfunction when reaching the 711 km orbit. This anomaly blocks the FM3 thrust burn activity to be deployed at the 800 km mission orbit. The reasons for this anomaly are still under investigation. The constellation deployment status as-of- December 2007 is shown in Figure 3-7. The dash line is the newly planned schedule and the dots recorded the execution results of the thrusting. The relative orbital separation angle, the relative AOL, and the relative altitudes of these four satellites are shown in Table 3-2 [57].

3.4.2 Spacecraft Thrust-Burn Performance Statistics

Figure 3-8 and Table 3-3 show the spacecraft thrust-burn performance statistic results in strip chart and table formats, respectively [57]. Starting from FM4 orbit transfer, the NSPO operations team uses the autopilot scheme to increase the burn success rate and reduce the burn working days. The data show that the FM5 burn working days number 39. However,

seven burns per day for FM4 and FM1 compared to three burns per day for FM5 as deployed earlier. The better spacecraft burn performance indicates that more successful rate has been achieved. The operations team has decreased the planned burn duration from 456 minutes for FM5 to 382.8 minutes for FM1 and also decreased the executed burn duration from 326.1 minutes for FM5 to 329.8 minutes for FM1. These results show that the thrust-burn success rate (= executed burn/planned burn) has been increased by the operations team from 71.5% for FM5 to 86.2% for FM1. Total burn number has increased from 53 times in FM5 to 71 times in FM1. From Table 3-3 it can be seen that the average orbit transfer height per burn has decreased from 5.4 km/burn for FM5 to 3.4 km/burn for FM1. Additionally, the average burn duration per burn has decreased from 369.4 sec/burn for FM5 to 238.4 sec/burn for FM1.

3.4.3 Spacecraft Mass Property and Moment of Inertia Results

We found that the propellant mass remains in the propellant tank are about 2.0 kg after the orbit-transfer operations are completed for each satellite. It is also expected that the spacecraft mass property (weight and center of gravity) and moment of inertia (MOI) are changed accordingly when propellant mass is changed. It was observed that the spacecraft center of gravity (CG) has a change of -0.7 cm shift in Z-axis before and after orbit-transfer activities, and has a CG shift in -Y and -X axes too. These changes will have a significant impact on the geodesy and earth gravity research [63]-[64]. Table 3-4 shows the spacecraft mass property and moment of inertia results of the six satellites. The spacecraft remaining propellant mass was estimated and provided by Propulsion subsystem. The error of the mass was estimated in the range of ±0.1 kg. Based on computation results, a very minor impact on MOI and CG results was observed due to this error range [57].

In the F3 satellites case, the TBB Boom and the Solar Panels are two portions that are deployed after satellite separation from the launch vehicle. The propellant fuel is also changed after orbit transfer. For the MOI computation, we assume the SAD is at 0o position.

The CG is valid for any SAD position, and therefore applies to the ACS Nadir and Nadir-Yaw Modes. The MOI and CG for six spacecraft were re-computed based on the above propellant mass.

3.5 Conclusion

We have presented a new fundamental operation concept for the F3 spacecraft constellation deployment, orbit-raising results, operations challenges and lessons learned. With five satellites (FM5, FM2, FM6, FM4, and FM1) successfully reaching the 800-km mission orbits as of December 2007, the F3 mission has verified the “proof-of-concept” of a novel way of performing constellation deployment by taking the advantage of nodal precession. This novel approach has dramatically reduced the spacecraft propellant mass and the complexity of the spacecraft RCS and ACS subsystem design. The success of the constellation deployment of the F3 mission has also provided a powerful demonstration of RO scheme in particular and for the remote-sensing applications of micro-satellite constellations in general. All these technical principles have paved the way for the design of future GNSS RO remote-sensing systems

TABLE 3-1 F3CONSTELLATION SPACECRAFT BUS KEY DESIGN

Mass ~ 54 kg (Dry Weight)

Power: ~ 81 Watts (bus and payload)

Shape Disc-shape of 116cm diameter, 18cm in height 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 3-2 CONSTELLATION DEPLOYMENT STATUS WITH FIVE SATELLITES (FM5,FM2,FM6,

FM4, AND FM1)AT FINAL ORBITS AS-OF-2DEC,2007

Items SMA Eccentricity Inclination 　 RAAN (Ωi/5) 　 AOL (Li/5)

SC No. (km) (deg) (deg) (deg)

FM5 799.475 0.0046 71.973 0 0 FM2 799.449 0.0041 72.037 29.9 50.7 FM6 799.444 0.0051 71.982 62.0 104.4 FM4 799.471 0.0072 72.009 90.0 158.2 FM3* 711.047 0.0054 72.012 129.9 Time Variant FM1 799.475 0.0046 71.973 145.9 262.53

*Note: On 3 Aug. 2007 the FM3 encountered solar array drive mechanism malfunction when reached 711 km orbit.

TABLE 3-3 SPACECRAFT THRUST-BURN PERFORMANCE STATISTICS Items Total Burn Days Total Burn Number Planned Burn Executed Burn Successfu l Rate Total Fuel Used Total Fuel Mass Average SMA/burn Average Duration/burn

SC No. (Days) (no.) (Minutes) (Minutes) (%) (kg) (kg) (km/burn) (sec/burn)

FM5 39 53 456 326.1 71.5 4.634 6.671 5.4 369.4 FM2 50 80 646.5 321.7 49.8 4.686 6.651 3.6 241 FM6 36 65 390 294.7 75.6 4.332 6.635 4.4 279.9 FM4 41 90 390.5 307.8 78.8 4.644 6.627 3.2 205.4 FM3 39 74 265.7 190.3 71.6 3.345 6.665 2.7 154.3 FM1 40 71 382.8 329.8 86.2 4.993 6.697 3.4 238.4

TABLE 3-4 SPACECRAFT MASS PROPERTY AND MOMENT OF INERTIA FOR SIX SATELLITES AS-OF-2 DEC, 2007 1

Items Total Mass (Full Tank) Remaining SC Total Mass Remaining Propellant +/- 0.1 kg Center of Gravity (CG) Moment of Inertia (MOI) Assume SAD = 0 deg

SC No. (kg) (kg) (kg) (m) kg m2 FM1 61.097 56.104 1.704 (94 psi/ 13.2 oC) x= 0.0035084 y=-0.0043757 z=-0.0334029 Ixx= 7.1677273 Iyx= 0.0288131 Izx=-0.0071984 Ixy= 0.0288131 Iyy=10.0887230 Izy=-0.4359628 Ixz=-0.0071984 Iyz=-0.4359628 Izz= 5.2806052 FM2 61.295 56.609 1.965 (100 psi/ 12.68 oC) x=-0.0034182 y=-0.0041841 z=-0.0364667 Ixx= 6.9711402 Iyx= 0.0292363 Izx=-0.0096030 Ixy= 0.0292363 Iyy= 9.8405863 Izy=-0.4376625 Ixz=-0.0096030 Iyz=-0.4376625 Izz= 5.2101918 FM3 61.295 57.950 3.320 (129 psi/ 27.86 oC) x= -0.0015454 y=-0.0070990 z=-0.0367495 Ixx= 7.0538797 Iyx= 0.3262446 Izx= 0.1441285 Ixy= 0.3262446 Iyy= 9.8458681 Izy= -0.2834290 Ixz= 0.1441285 Iyz= -0.2834290 Izz= 5.1711034 FM4 61.020 56.376 1.983 (105 psi / 29.10 oC) x = -0.0037843 y = -0.0073189 z = -0.0371947 Ixx= 6.8193710 Iyx= 0.0317362 Izx= 0.0744942 Ixy= 0.0317362 Iyy= 9.7484668 Izy=-0.4389625 Ixz= 0.0744942 Iyz=-0.4389625 Izz= 4.8734748 FM5 61.167 56.533 2.037 (98 psi/ 13.68 oC) x=-0.0036067 y=-0.0045262 z=-0.037113 Ixx= 6.9437632 Iyx= 0.0275360 Izx=-0.0087138 Ixy= 0.0275360 Iyy= 9.8007081 Izy=-0.4379625 Ixz=-0.0087138 Iyz=-0.4379625 Izz= 5.2086237 FM6 61.315 56.983 2.303 (106 psi/ 18.40 oC) x = -0.0032281 y = -0.0044101 z = -0.0360353 Ixx= 6.9827399 Iyx= 0.0289346 Izx=-0.0115537 Ixy= 0.0289346 Iyy= 9.8596525 Izy=-0.4397625 Ixz=-0.0115537 Iyz=-0.4397625 Izz= 5.2408835