development of satellite remote sensing systems in...

Hydrological Applications of Remote Sensing and Remote Data Transmission (Proceedings of the Hamburg Symposium, August 1983). IAHS Publ. no. 145.

Development of satellite remote sensing systems in Japan

NATIONAL SPACE DEVELOPMENT AGENCY OF JAPAN (Yasushi Horikawa, 2-4-1 Hamamatsu-cho, Minatoku, Tokyo, Japan)

ABSTRACT In 1978, the Space Activities Commission of Japan, the policy making body of the Japanese space programme, produced an "Outline of Japan's Space Development Policy", which proposed that a marine and land observation satellite series should be developed in order to establish an earth observation operational system. The National Space Development Agency of Japan (NASDA) is carrying out the research and development of the satellite remote sensing system including satellites, sensors and ground facilities. Marine Observation Satellite-1, which is the first satellite of the series, is now under development with a launch targeted for 1986. The development of Earth Resources Satellite-1, which will carry a synthetic aperture radar, has been in progress since 1980. The research and development of active microwave sensors, which will be mounted on the follow-up satellite of the marine observation satellite series, are being carried out. This paper outlines the present status of development of satellite sensing systems in Japan.

Le développement des systèmes de télédétection par satellites au Japon RESUME En 1978 la Commission des Activités Spatiales du Japon, l'organisme responsable de la politique du programme spatial japonais a mis au point un "Résumé de la Politique de Développement Spatial du Japon" qui proposait qu'une série de satellites d'observations marines et terrestres soit mise au point pour réaliser un système opérationnel d'observations terrestres. L'Agence Nationale de Développement Spatial du Japon (NASDA) exécute les recherches et la réalisation d'un système de télédétection par satellites comportant les satellites, les senseurs et les aménagements au sol. Le Satellite-1 d'Observations marines, qui est le premier satellite de la série est actuellement en cours de réalisation, son lancement est programmé pour 1986. La mise au point du Satellite-1 de Resources terrestres, qui portera un radar à ouverture synthétique a progressé depuis 1980. Les recherches et la mise au point de senseurs actifs à micro-ondes qui seront montés sur le prochain satellite de la série des satellites d'observations marines, sont en cours. Cette communication donne un aperçu sur l'état actuel de développement du système de télédétection par satellites au Japon.

45

46 National Space Development Agency of Japan

MARINE OBSERVATION SATELLITE-1

Marine Observation Satellite-1 (MOS-1), Japan's first earth observation satellite, is an experimental satellite to establish the fundamental technologies which are common to both marine and land observation satellites and to collect information on the earth's surface.

The conceptual design and the preliminary design of MOS-1 were carried out in 1978 and 1979 respectively, and the basic design was completed at the end of July 1981. The detailed design was completed at the end of June 1983 and a prototype satellite is now being manufactured. MOS-1 will be launched by a N-II launch vehicle from the Tanegashima Space Centre in 1986.

MOS-1 system and objectives

The mission objectives of the MOS-1 programme are as follows; (a) Establishment of fundamental technologies which are common

to both the marine and land observation satellite. (b) Observation of the state of the sea surface and atmosphere

using visible, infrared and microwave radiometers, and verification of the performance of these sensors. In order to accomplish these objectives, MOS-1 carries three types of sensors: multispectral electronic self scanning radiometer (MESSR); visible and thermal-infrared radiometer (VTIR) and microwave scanning radiometer (MSR). Selected orbital parameters are as follows:

Altitude about 909 km Inclination about 99.1 degrees Recurrent period 17 days Local time of decending node 10-11 a.m.

The orbit will be adjusted during the two years after launch by an orbit control system to keep the cross track drift from the nominal orbit of ground track at the equator within 20 km. Satellite control will be made by using NASDA's satellite tracking and control system and data acquisition facilities will be installed at the Earth Observation Centre (EOC) located about 50 km northwest of Tokyo, presently receiving the Landsat data. The total MOS-1 system is illustrated in Fig.l.

Sensors

As noted, MOS-1 will mount three types of sensors to observe visible, near infrared, infrared and microwave regions. A brief overview of each sensor follows.

Multispectral electronic self-scanning radiometer (MESSR) A unique feature of this radiometer is that CCD is selected as the image detector to eliminate the moving portion in the sensor. The CCD is composed of 2048 photo-sensitive elements, and the size of one photosensitive element is about 14 yrad x 14 yrad, which corresponds to a measuring area of 50 m x 50 m on the ground. One CCD detector produces an earth image 100 km wide, such that a pair of sensors is required to cover the 180 km width which is the distance between adjacent orbits of MOS-1.

Satellite remote sensing systems in Japan 47

TLM,CMD

Tracking and Control Center (TACC)

I

Tracking and Control System

Operation Information

Operations ! Control j-System )

Operation Information!

Earth Observation Data Acquisition, Processing and Archiving Facilities

Data Distribution Office

Request =E

Earth observation Data Acquisition and Distribution System

Users FIG.l

I Data Users

Total system of MOS-1

Figure 2 indicates the image producing concept of MESSR. Separate or simultaneous operation of sensors can be made by ground command. Major characteristics of MESSR are shown in Table 1.

Each sensor is composed of two optical systems which provide an earth surface image of four bands in the visible and near infrared region given in Table 1. Image signals generated by CCD are fed to

e GHZ

Image

FIG.2 Image producing concept of MESSR.


TABLE 1 Characteristics of the MESSR

Item MESSR characteristics

Wavelength

IFOV Swath width (one optical

element) Scanning method Optics Detector S/N Quantization levels Data rate

Power Wei gh t

0.51-0.59 \im 0.61-0.69 yra 0.72-0.80 ]lm 0.80-1.10 \lm 54.7 ± 5 \xrad

100 km electrical Gauss type 2048 elements CCD 39-15 dB 7.6 ms 8.78 Mbits s~l (including VTIR data) 69 W 64 kg

signal processing unit and converted to six bit digital signals. Signal produced by the VTIR is also combined in the MESSR data stream, and these combined data are sent to the modulator.

A solid state transmitter on 8 GHz band is newly developed to transmit the high speed image data. Satisfactory data have been obtained during BBM phase and design efforts will be continued to meet temperature conditions and environmental conditions required in orbit.

Visible and thermal infrared radiometer (VTIR) The VTIR has one visible band and three infrared bands. Scanning of the earth surface is made by a rotating mirror which has an aperture diameter of about 15 cm. Si-PIN diode and HgCdTe are selected for the visible and infrared detectors, respectively. IFOV of this radiometer is 1 mrad for visible band and 3 mrad for infrared bands which correspond to about 1 km x 1 km and 3 km x 3 km on the ground, respectively. Arrangement of IFOV is illustrated in Fig.3. Characteristics of the VTIR are tabulated in Table 2.

Each detector developed for this radiometer has two photo-electric converting elements on the focal plane to increase reliability. Generated image signals are A/D converted and fed to the MESSR signal processing unit as previously mentioned.

Development effort is also continued on this VTIR, and radiation cooling characteristics for infrared detectors is being carefully examined.

Microwave scanning radiometer (MSR) The MSR is composed of two Dicke type radiometers with frequencies of 23.8 GHz and 31.4 GHz band to observe sea surface temperature and liquid water/water vapour in the atmosphere. Characteristics of MSR are given in Table 3.

Satellite remote sensing systems in Japan 49 IR

n + 1

n + 2

1 m rad n + 1 3 in rad

Direction of Scanning

n + 2

Direction of Satellite Track

FIG.3 Arrangement of IFOV of VTIR.

TABLE 2 Characteristics of the VTIR

Item

Wavelength

IFOV Swath width (km) Scanning method Scan period Detector Optics S/N

NEAT

Quantization level Total power Total weight

VTIR characteristi Visible

0.5-0. 7 \xm

1 mrad 500 km Mechanical 1/7.3 s Si-PIN Diode Ritchey-Chreti 55 dB (Alb. = 80%) ~~

256 (8 bits) 35 20

en

W

kg

CS:

Thermal infrared

6.0-7.0 \im 10.5-11.5 ym 11.5-12.5 ]im 3 mrad 500 km Mechanical 1/7.3 s Eg Cd Te Ritchey-Chretien

_ 0.5 K (at 300K)

256 (8 bits)

EARTH RESOURCE SATELLITE-1

System requirements and configuration

The total system of the ERS-1 programme is shown in Fig.4. This ERS-1 programme is under study. Image data from the synthetic aperture radar (SAR) and visible and near infrared radiometer (VNR) will be transmitted to data receiving stations of the earth observation station in Japan and foreign data receiving stations. ERS-1 tracking will be done by NASDA tracking stations and the TDRS net-


TABLE 3 Characteristics of MSR

Item MSR characteristics

Frequency Beam width Integration time Swath width Scanning method

Dynamic range Antenna type Receiver type Accuracy Scan period Quantization level Data rate Total power Total weight

23.8 ± 0.2 GHz 1.99° 10, 47 ms 317 km Mechanical (conical scan) 30 K-330 K Offset casegrain Dicke 1.5 K (at 300 K)

3.2 1024 (10 2 K bits

31.4 ± 0.25 GHz 1.45° 10, 47 ms 317 km Mechanical (conical scan) 30 K-330 K Offset casegrain Dicke 1.5 K (at 300 K)

bits) s " 1

60 W 54 kg

1024 (10 bits) 2 K bits s _ 1

work of the United States. ERS-1 will be launched by an H-I launch vehicle (two stages) of NASDA from Tanegashima Space Centre in Japan. Launch capability of H-I is approximately 1400 kg with about 570 km altitude circular orbit and about 98° inclination.

ERS-1 system design was performed based on the following criteria. System design is conducted with such technologies as design, test and integration techniques developed in the past space programme.

Since each subsystem will be developed separately as a module from the procurement policy, system requirements and interface conditions must be cleared. Some components will be, however, imported from foreign suppliers.

Each subsystem utilizes existing technology as much as possible. However, if new technology is needed, this must be started in an

TDRS. G.S. (White Sands)

NASDA TACC

Earth Observation Center

FIG.4 Total system of ERS-1.


early phase. SAR, VNR and bus equipment will be designed by domestic technology from this standpoint. Spacecraft bus equipment will achieve high reliability, weight reduction and low power consumption. System and subsystem design of ERS-1 will take into account the expansion of following operational and large scaled satellites. Also the design must consider the test facility and test method.

Design of mission equipment

Mission equipment of ERS-1 will be described for SAR, VNR, MDR and MDT.

Synthetic aperture radar (SAR) SAR is a main observation equipment to establish the technology of an active sensing satellite. L-band radar frequency was selected from the developing feasibility of antenna flatness and high power transmitter. Off nadir angle of 33 of antenna was selected from the point of view feasibility of pulse repetition frequency, signal to noise ratio and signal to ambiguity ratio. SAR characteristics are shown in Table 4.

TABLE 4 SAR characteristics

Item Characteristics

Swath width Spatial resolution Off nadir angle Transmitting frequency Polarization RF band width S/N S/A Da ta ra te Transmitting power Pulse width Pulse compression ratio Pulse repetition frequency Antenna size Weight antenna Weight electronics

74 km 25 m x 25 m 33 degrees 1275 MHz H-H linear 12 MHz 7 dB 20 dB 60 MHz 1 kW peak 35 us 450 1550-1690 pps 2.4 m x 12 m 134 kg 120 kg

Visible and near infrared radiometer (VNR) VNR is an improvement of MESSR in the area of resolution and swath width installed in MOS-1 which was the first remote sensing satellite in Japan. VNR data will be used not only for optical observation but also complement SAR data. Characteristics of VNR are shown in Table 5.

Mission data recorder (MDR) Observation of ERS-1 will be done mainly by the existing Landsat station. However, as a backup to the Landsat station and for the area where the Landsat station is not available, a high density data recorder, called Mission Data Recorder


TABLE 5 VNR characteristics


Swath width 150 km Spatial resolution 25 x 25 m Wavelength (1) 0.45-0.52 ym

(2) 0.52-0.60 \im (3) 0.63-0.69 ym (4) 0.76-0.95 ym

IFOV 44 \irad FOV 15.4° Image acquisition time 3.6 ms Weight 40 kg

(MDR), is installed in ERS-1. This data recorder will be procured from the USA.

Characteristics of MDR are shown in Table 6.

TABLE 6 MDR characteristics


Capacity 272 Gbits Data rate (input/output) 30 M pbs x 2 ch Recording/reproducing time 20 min Weight approx. SO kg

Mission data transmitter (MDT) Observation data will be transmitted through the Mission Data Transmitter (MDT). In order to receive the ERS-1 data at the Landsat station, a 20 W TWTA transmitter will be used. A difficult problem with this transmitter is the on/off cycle of the transmitter. Reliability in this field will be studied further.

Characteristics of MDT are shown in Table 7.

GMS-2 SYSTEM

Mission objectives

As a member of WMO, Japan responded to the needs of the Global Atmospheric Research Programme (GARP) and WWW by developing a Geostationary Meteorological Satellite, known as GMS. In July 1977, GMS was launched into geosynchronous orbit, approximately 36 000 km above the equator at 140°E longitude. Epitomizing the spirit of international cooperation manifested by the Global Observing System, Japan's GMS is joined in its celestial watch by the United States'


TABLE 7 MDT characteristics


Frequency Data rate Modulation EIRP

Radiation pattern Polarization RF band width Weight

8025-8400, 2 frequencies 60 M bps/1 frequency QPSK EL 90° 2 dBW EL 5° 17 dBW Shaped broad beam RHCL 60 MHz/1 frequency 40 kg

GOES satellites positioned at 75 and 135°W; Europe's METEOSAT at 0°; and Russia's GOMS at 70°E (GOMS replaced by GOES at 57°E during the First GARP Global Experiment (FGGE)).

GMS-2 (see Fig.5), the successor of GMS, was developed to continue

FIG.5 Geostationary Meteorological Satellite "GMS-2".


this meteorological satellite service. The GMS-2 was launched by Japanese N-II rocket from Tanegashima

Space Centre, Mission objectives of GMS-2 are fundamentally the same as GMS: weather watch by VISSR; collection of weather data; distribution of weather data; monitoring of solar particles.

Progress of GMS-2 programme

As shown in Fig.6, the basic and detailed design of GMS-2 was performed in 1978. During 1979 and 1980, two spacecraft (proto-flight model and flight model) were assembled, integrated and tested by Hughes Aircraft Company. After the system integration test, two spacecraft were shipped to Japan. The protoflight model was stored at the Tsukuba Space Centre as a back-up. The flight model spacecraft was checked out and prepared for launch at Tanegashima Space Centre of NASDA. After launch, an in-orbit check of GMS-2 was performed.

Basic Design

Detailed Design

Assembly & Manufacture

Subsystem Test

Integration QT/PFM

Integration AT/FM

Ship to Japan

Launch Base Test

On Orbit Check Station Change

Operation

1978

A A PDR CDR

1979 1980 1981

A A PSR Launch

C3

C=3

=

<

I

FIG.6 Progress of the "GMS-2" programme.

GMS-2 configuration

The GMS-2 is a spin-stabilized geostationary meteorological satellite with mechanical despun antennas. The configuration and characteristics of the spacecraft are improved over those of the predecessor, GMS and most subsystems are flight-proven. The configuration is shown in Fig.7.


USB OMNI ANTENNA DESPUN ANTENNA ASSEMBLY

DESPIN BEARING ASSEMBLY VISSR SUNSHADE

DYNAMIC BALANCE

FIG.7 Overall view of GMS-2.

Overall size, weight and shape of the spacecraft are designed to be compatible with the N-II launch vehicles. The spacecraft length is 444 cm at launch and 345 cm on station, and the diameter is 215 cm. The weight when GMS-2 is separated from N-II third stage is 653 kg, and the spacecraft end of life (EOL) weight is 285 kg.

The satellite mission life is three years due to the limited amount of on-board hydrazine fuel; however, the design life is five years. Redundancy of mission-critical functions is provided to ensure electronic lifetimes significantly in excess of five years. The solar panel power of 264 W includes an approximately 30 W margin at the end of five years (summer solstice).

Future plan

In spite of several anomalies which were observed after launch, the GMS-2 is now performing well and should provide excellent meteorological services until 1985, its expected mission life. Accordingly, the next weather watch satellite, GMS-3, is now under consideration. The GMS-3 programme will consist of two spacecraft, designated 3a and 3b. The GMS-3a will be the GMS-2 proto-flight spacecraft refurbished to provide more capability and modified to improve its reliability. The GMS-3b, will be a back-up for 3a, and will be almost identical in design to the GMS-3a. The GMS-3a was scheduled to be launched in 1984.


ACTIVE MICROWAVE SENSORS

As the baseline of designing active microwave sensors, orbit parameters and system performance requirements are tentatively settled as shown in Tables 8 and 9, respectively.

TABLE 8 Orbit parameters

Altimeter and scatterometer

SAR

Height Eccentricity Orbit

800 km 0.004 max.

570 km

Sun-synchronous

TABLE 9 System performance requirements

Altimeter Seatterometer

Geodetic accuracy

Topographic accuracy

Wave height range

accuracy Reflection

coefficient Acquisition

time

50 cm

20 cm rss

1-20 m max.(0.5 m, 10%)

Wind velocity range accuracy

Wind direction range accuracy

Swath width Grid spacing

±1 dB

Less than 6

4-25 m s max.(2ms , 10%)

0-360 degrees ±20 degrees 200-700 km 50 km

SAR

Resolution Swath width

25 m 75 km

Radar parameters are summarized in Table 10. and DC power requirements are not fixed yet.

Total weights, sizes,

Microwave altimeter

In this section, the major functional elements of the altimeter are described. A block diagram of the altimeter is shown in Fig.8. This system can be divided mainly into three sections: RF section, signal processor section and tracking processor section.

The functions of RF section are to transmit and receive radar pulses, and to process received signals with a full-deramp technique. This basic design is similar to the SEASAT-1 altimeter.

An analogue-to-digital conversion of I and Q video signals from


TABLE 10 Preliminary radar parameters

Frequency Transmitted

band width Uncompressed

pulse width Compression

ratio Transmitted RF

peak power Pulse, repetition

frequency Noise figure Antenna beam

width

Antenna beam centre gain

Antenna pointing angle

Signal to noise power ratio

Surface resolution

Altimeter

13-14 GHz

320 MHz

3.2 ys

1024

2 kW

1000 Hz 5.5 dB

1.6°

40 dB

nadir

>10 dB

25 km

Scatterometer

13-14 GHz

4 KHz

5 ms

-

100 W

40 Hz 5.3 dB

0.5°x24° (orthogonal) 0.5°x20° (3rd antenna)

32 dB

43°(orthogonal) 37°(3rd antenna)

Ï-15 dB

25 km

SAR

1.2-1.3 GHz

13 MHz

35 s

450

1.5 kW

1600 Hz 4.5 dB

6.2°(range

1.0°(azimuth)

34 dB

33°(off nadir)

>7 dB

25 m (range) 25 m (azimuth,

4 looks)

Antennaj

S/C-

RF section

Full-deramp processing

Grid modulator| ZE HVPS

Up-converter s frequency multiplier

SAW-DDL

X Frequency synthesizer

S/ Power system

HSWS

Signal processor

Synchronizer

I I

4] Digital filter

iJ Averaging J

Tracking processor interface

ASG

^_P Filtering

I

S/C

FIG.8 A block diagram of the altimeter.


RF section is accomplished in high speed waveform sampler (HSWS), which converts and stores 64 samples of these signals at a 20 MHz rate during 3.2 s. Digital filter transforms 64 samples into ones in frequency domain and averages them to provide smoothed waveform samples (50-100).

In the tracking processor section, oceanic parameters (altitude, wave height and signal-to-noise ratio) are estimated and predicted by using smoothed waveform samples. Quality of waveform samples depends on prediction accuracy, because predicted parameters control receiver trigger and AGC (automatic gain control). To improve prediction accuracy, maximum likelihood estimation (MLE) (including a-g filter) is used as an algorithm of estimation. MLE can simultaneously estimate three parameters and reduce variances of them in conparison with the adaptive split gate (ASG) a-3 tracker used in the SEASAT-1 altimeter.

Some problems about MLE have been considered, as follows; (a) MLE needs a large amount of calculation; (b) MLE needs an accurate waveform model; (c) MLE has not yet been used on a satellite.

The first and second problems are settled by using a recent advanced microprocessor, which has features of 16 bits/word and high speed calculation. Simulation for MLE operation has been executed in various conditions, and this system has the capability of selecting MLE or ASG. The third problem seems not to be important, considering simulation results and this configuration. Moreover, this configuration makes it easy to compare this system with the SEASAT-1 altimeter.

Microwave wind scatterometer

Microwave wind scatterometer is a pulse radar system with fan beam scanning and doppler filtering. The principle of design is basically similar to that of SASS on SEASAT-1. However, our system has a third antenna on each side of the satellite in order to remove the alias solutions of wind directions, and also has the dual independent polarity systems for transmitting and receiving signals.

Features of the system The features of the system are as follows:

(a) Tri-directional antennas. As mentioned above, this scatterometer has basically a three beam system. In order to determine the beam direction of the third antenna, wind vector inferring simulation was executed. Simulation results show that the most preferable direction of the third antenna is 75° in the case of Doppler radar system and the capability for wind vector determination will fairly upgrade as compared with SASS.

(b) Footprint editing. The observation area is located along both sides of the sub-satellite track. Each swath width is 500 km and divided into 20 unit cells (25 km x 25 km) by Doppler filter bank. In order to observe the sea closely by using the spacecraft movement, the six antennas are switched sequentially in about a 2-s interval. In an interval, four pulses and 16 pulses are assigned to the third and orthogonal antennas respectively, so that the data taken by each antenna have the same accuracy. Consequently, a 50 km x 50 km cell


consists of eight unit cells in each beam (Fig.9(a)), and is observed from three different directions in as many times (Fig.9(b)). Eight unit cells are averaged in normal operation. According to the various demands for sea phenomena, closer cell constitution is also possible.

FIG.9 Footprint overview: (a) unit cells; (b) three beam pattern.

(c) Dual polarity. This system can measure the ocean in dual polarity (vertical and horizontal).

Synthetic aperture radar

In this section, outline of the SAR research and development model is described. Tables 9 and 10 indicate the basic SAR parameters. The SAR system consists of transmitter, receiver controller and digital section.

Spacecraft

SAR Antenna

Center Arm

1st stage 2nd stage ~^\~ 3rd stage

FIG.10 The deployment sequence of SAR antenna.

KEVLAR Radiation Element _/_

T

6mm Radiation Panel

Support Panel

15mm

Aluminum Core

FIG.11 The configuration of SAR antenna.


The received echo is detected synchronously by the transmitted signal to get "hologram" data. The hologram data are converted into digital format to be recorded on-board or to be transmitted with an X-band data link.

The antenna is one of the components that we are taking effort in developing now. The SAR antenna consists of micro-strip array of 1024 elements and its dimension has 2.1 m x 12 m. Due to the constraints of 2.2 m diameter of H-I launch vehicle's payload section, the antenna must be déployable.

A solar panel deployment mechanism is applied to the antenna expansion. The deployment sequence consists of three stages, as shown in Fig.10. In the first stage, the antenna package is released to the right angle from the sidewall of spacecraft. In the second stage, two halves of the antenna deploy on either side of the centre arm simultaneously. In the third stage, the deployed antenna tilts to the off-nadir angle. In each stage, a spring-operated latch-up mechanism is employed for locking.

A support panel, as shown in Fig.11, consists of a honeycomb sandwich structure in order to keep the antenna panel flat in the space environment.

These are major differences from the SAR system of SEASAT-1.

development of satellite remote sensing systems in...

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