The IRS-P3 remote sensing mission

The IRS-P3 remote sensing mission

Acta Asrronaurica. Vol. 39, No. 9-12, PP. 711-716, 1996 91997 Published by Elsevier Science Ltd Printed in Great Britain PII:SOO94-5765(97)53-2 0149-1...

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Acta Asrronaurica. Vol. 39, No. 9-12, PP. 711-716, 1996 91997 Published by Elsevier Science Ltd Printed in Great Britain PII:SOO94-5765(97)53-2 0149-1970/96 $17.00+0.00


THE IRS-P3 REMOTE SENSING MISSION K. Thyagarajan ISRO Space Application Center, ISAC Airport Road, 560 0 17 Bangalore, India A. Neumann, G. Zimmermann DLR, Institute of Space Sensor Technology 12489 Berlin, Rudower Chaussee 5, Germany ABSTRACT ISRO has developed the PSLV rocket (Polar Spacecraft Launch Vehicle) for polar orbiting satellites up to 1000 kg and is conducting a series of test missions. One of this is the IRS-P3, an remote sensing satellite with German participation. The payload consists of 3 scientific instruments: The wide field sensor WiFS for vegetation monitoring (ISRO), the imaging spectrometer MOS (DLRGermany) for coastal zone and ocean studies an the X-ray astronomy payload (ISRO). The paper gives technical details and parameters on the launch vehicle, the satellite, the instruments and scientific goals and data utilization. 01997 Published by Elsevier Science Ltd. of mass 283 t for launching remote sensing satellites up to a mass of about 1000 kg. It is powered by solid propellant first and third stage and liquid propellant second and forth stage. A 2.8 m diameter core motor and six 1.O m diameter strap on motors constitute the first stage. The satellite heat shield has 3.2m diameter. The vehicle is provided with S-Band PCM-TM and C-band transponder for performance monitoring of vehicle subsystems, tracking and trajectory information during launch and Preliminary Orbit Determination. The standard orbit is H = 817 km, polar sunsynchroneous circular, i = 98.69”, period 101.35 min., local time lo:30 descending, repetition 24 days. For future missions are also under study CSS-orbits at 720 and 570 km.

1. INTRODUCTION IRS-P3 Spacecraft has been developed to support the PSLV launcher development and to enhance and improve the IRS mission capabilities toward operational&y and application. The spacecraft is a low cost mission and uses MIL-STD-883B parts with upgrading as necessary, also for the payload instruments. To reduce the cost and development time has been used of-the-shelf items of other ISRO programs and instrument developments. Apart from technology testing and demonstration the mission objectives are: l provide the opportunity for RS application in the areas of land, atmosphere and oceanographic investigations l to validate new RS methods and develop affiliated application potential l to provide opportunity for experiments in X-ray astronomy. IRS-P3 was successful launched D3 rocket on March 2 1, 1996.



with PSLV-


The IRS-P3 bus is derived from flight proven IRS-l A,B,C and IRS-P2. Main technical parameters are given in Tab. 1. Some new parts are the processor based attitude and orbital control system (AOCS) from INSAT-II for large angle manoeuvres, the processor based telecommand system and FPGA based


The Polar Satellite Launch Vehicle (PSLV) is four stage rocket with height of 42,2 m and lift 711

Small Satellites for Earth Observation


TM system. ISRO batteries (24 Ah).







Diahotsr *S2m VatMa Dimatt)t


Fig. 1 PSLV Configuration The satellite bus structure consists of four vertical panels and two horizontal decks supported on a central load bearing cylinder of 930 mm 0 and 1188 mm height. At the inner sides of the panels and decks are mounted the bus subsystems. The outer side of earth viewing panel carries the payload data antennae, the TTC antenna and attitude sensors. The scientific payload is on the outer side of the upper deck, which is oriented in flight direction in orbit. The thermal control system of IRS is configured to provide suitable temperature environments for various subsystems during all phases of satellite orbit life. The design philosophy of maximum use of passive elements and minimum use of semi-active elements for end-of-life surface radiation

characteristics of the satellite is followed. This is achieved by extensive use of thermal control coatings. thermal control tapes, optical solar reflectors (OSRs), multilayer insulation blankets (MLI), conductive grease etc., and temperature control surface heaters to compenr-te for the shortfall in beginning-oflife power dissipation. While the subsystem packages are maintained between 0°C and 40°C the scientific instruments and battery are maintained within narrow temperature ranges. The power system consists of solar array, battery, core and distributed power electronics to provide uninterrupted power of approximately 370 W to the spacecraft. The configuration is based on dual bus of 42 volts. The solar array consists of 2 sections feeding to the two buses. There are two batteries directly connected to the respective raw bus. The solar array power is controlled by shunt switches. The battery charge control is achieved by dissipationless pulsewidth modulated taper charge regulators. The solar array is designed to provide a power of 810 W at end-of-life trough 9.6 square meter suntracking rigid solar panels. Two Ni-Cd batteries comprising of 28 cells each, having a capacity of 24 Ah supply continuous eclipse power and support payload operation in orbit. Instantaneous peak loads are also supported by batteries. The telecommand system operates in Sband with PCMIFSKIFMIPM modulation and provides time-tag command execution facility. It supports auto commanding with suitable override commands for back up. It houses both programmable and fixed duration timers for payload operation. The telemetry system adopts PCM/PSK modulation in S-band (2203 and has dwell mode facility. MHz) Additionally it has provision to store approximately one orbit data that can be played back during next visible pass. The tracking is provided by range and doppler measurements using S-band TTC transponder. The payload data is transmitted in S-band (2280 MHz) with BPSK modulation at a data rate of 5.2 MBPS. The Attitude and Orbit Control System (AOCS) does the function of attitude orientation and orbit control of the spacecraft to the required accuracys and does also the 3axis autoacquisition and control from the moment of injection into the orbit and also

Small Satellites for Earth Observation total lift off mass Solar arrays Attitude and OCS Pointing accuracy Transponder 1Frequency Date rate onboard losses Antenna gain for maximum range RF Power


920 k 8 9.6m ,817 W,24AhNiCd Battery 3-axis stabilisation 0.2” Nadir Mode 0.0 lo Stellar Mode 2.28 GHz 5.2 Mbps 2 dB 3.5 dB 5W

Tab. 1 Satellite Characteristics puts the spacecraft in a safe orientation in the case of contingency like earth loss. Besides the conventional earth pointing mode for remote sensing applications, AOCS also configures the spacecraft with star sensor in loop to orient X-ray payload astronomy observations. This mode is called as the stellar pointing mode (the mode of spacecraft operation with roll axis pointing towards pre-defined X-ray stars). AOCS consists of attitude sensors (earth sensors, sun sensors, inertial sensors, star sensor), Attitude and Orbit Control Electronics (AOCE) and actuators like hydrazine thrusters, reaction wheels and magnetic torquers. Earth sensors give error information about pitch and roll axes of the spacecraft and the inertial sensors (DTG) give yaw error information. Sun sensors provide attitude information with respect to sun for acquisition and sun orientation. There are two magnetic torquers for momentum dumping of wheels. The Propulsion system is a monopropellant hydrazine system and is configured with 100% redundancy having 4 propellant tanks (two bladder type and two surface tension type) cross connected with each other through latch valves and feeding two functionally redundant thruster blocks with 8 thrusters (1 .O Newton) in each block. One 11 .O Newton thruster is mounted along negative roll axis of the spacecraft that is used for correcting initial injection errors and inclination correction. The spacecraft carries 84 kg of fuel.



IRS-P3 spacecraft was launched successfully by PSLV-D3 vehicle on March 21, 1996. The spacecraft was injected into 810 x 844 km orbit with 98.7897 deg. inclination and the dispersion errors of the vehicle are very well within limits (A V of around 12 m/set.). During initial phase of the mission (6 weeks), the different payload and bus performance was validated/verified. Following this the spacecraft has been put into a nominal 8 I7 km, circular sun-synchroneous orbit and is phased 180 deg. with respect to IRS-1C (which also carries WiFS payload) for faster coverage. The satellite bus performance is normal. The solar array is generating around 980 watts in earth pointing mode and 1060 watts in stellar orientation. The performance of technology demonstration elements like processor based command system, FPGA based telemetry system, inhouse developed Ni-Cd battery, stellar orientation mode of AOCS with star sensor inloop has been nominal. The AOCS in earth orientation provides pointing accuracy of + 0.2 deg. in all axes and the same in stellar orientation is better that f 0.01 deg. (with the use of star sensor in control system loop). The spacecraft orientation in stellar pointing from earth pointing was demonstrated successfully by orienting to X-ray source sgn-xi. The spacecraft has the capacity to orient to any direction in space along with maximum power optimisation with star sensor in loop. All these large angle manoeuvres were first carried out with thrusters initially and later through reaction wheels. The performance of X-ray astronomy payload is satisfactory and so far the spacecraft has been oriented towards 5 Xray sources (both bright and weak ones). The orbit performance of the RS instruments WiFS and MOS is good, both sensors are under validation. The spacecraft TTC support for payload operation is provided by ISRO networks at Bangalore, Lucknow, Mauritius, Bearslake station near Moscow besides Weilheim station of DLR. The remote sensing payloads data is received and processed at Shadnagar (Hyderabad) India and Neustrelitz, Germany. In 1997 the number of ground stations will be extended, e.g. Maspalomas ESA) and a mobile station for test sites at other regions,


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5. THE PAYLOAD OF IRS-P3 5.1 Wide Field Sensor (WiFS) The WiFS is a solid state push broom camera operating in three band in visible and shortwave infrared (SWIR) regions (B3, B4 and B5). It is a wide angle camera providing a ground resolution of 188 meters (from 817 km orbit). Each band has two detectors centred at + 13.6 deg to achieve the wide swath of 8 10 km on ground, which enables a repetitive cycle of 5 days (see Tab. 2.). The band B3 and B4 use CCD array 143A with 2048 elements with size of 13x13 micron. The B5 2100 element THX 3 1901 SWIR detector operates in 1.55 - 1.70 micron spectral range at temperature of -5 or -10” C respectively. The optics system consists of 8 lenses with spectral band pass and neutral density filters for each spectral band. The WiFS is designed to have saturation radiance setting for different gains(4) and spectral bands. The dynamic range in each gain is 7 bits. The absolute radiometric accuracy is better than 10% with relative radiometric accuracy among the bands within 2%. The data rate for B3 and B4 is 2.6 MBPS and that for B5 is 1.73 MBPS with dummy pixels.

The misregistration of SWIR band B5 with visible band B3 and B4 is f 0.25 pixels.

Spatial resolution Swath

B 3:0.62 - 0.68 pm B 4:0.77 - 0.86 pm B 5:1.55 - 1.75 pm 188m 810km/4096pxl

Radiometric resolution Ground repetivity

7 bit 5 days

Spectral bands

Power so w Weight 25 Kg I Tab.2 Wide Field Sensor WiFS Characteristics

5 3 The Irmginp SpectroW


The MOS instrument [l] was designed for remote sensing of the Ocean-Atmosphere system in the VIS-NIR spectrum. It consists of spectrometer blocks: the two separate atmospheric spectrometer MOS-A provides 4 narrow channels in the OzA-absorption band at

- 760 nm to allow measurements that can be used to estimate the aerosoloptical thickness and stratospheric aerosols. It measures simultaneously with the bio-spectrometer MOS-B that has 13 channels of 10 nm width in the range from 408 to 1010 nm. MOS-IRS provides a 14th channel from the MOS-C camera in the SWIR for improved ocean surface term and roughness estimation (technical parameters see Tab. 3). Using the MOS-A measurements and the NIR-channels of MOS-B it is possible to remove the atmospheric influence from the MOS-B data for computation of the water leaving radiance (reflectance) on the surface level. The advantage of the OzA-method is to provide additional measurements on aerosol content and profile. The MOS-B channels are chosen in accordance with the spectral characteristics of ocean and coastal zones and appropriate to construct quantitative retrieval algorithms of different water constituents. They also give the opportunity of vegetation signature determination and estimation of Hz0 -vapour content in the atmosphere from the NIRmeasurements( e.g. po’C-band). High attention was drawn on radiometric accuracy, resolution and absolute calibration of the instrument. The MOS device is equipped with a Sun calibration unit.



5.3 MOS Cahbratlon Procedures The main goal was a good radiometric resolution. By careful design of the multiarray focal plane controlling and the analogue signal electronic (up to ADU) and temperature stabilizing at + 5 + 0.1 “C the resulting S/N allowed to work with 16 bit digitalization. The dynamic range of 64,000 counts is sufficient for the high values for the sun-calibration at the upper level and the dark ocean at the lowest level. Apart from spectral and radiometrical measurements before launch this parameters are determined in orbit by comparison with extraterrestrical solar radiation and by internal lamps. For calibration to the Sun the solar irradiance

is measured


a white



when the satellite

is crossing



( 02A-band spectral halfwidth Inm] FOY along track x ldegl




1.4 0.344

1 815:

945 (H20-


1 1

WpX) 10 0.094

1 1

I IW 0.14

Operational energy range FoV Pin hole size Distance to detector Detector Detector cell size, length Window Filling gas


DATA Tab. 3 Technical Parameters of the Modular Optical Imaging Spectrometer MOS-IRS


The X-ray astronomy payload is designed to study periodic and aperiodic intensity and spectral variations in galactic and extragalactic X-ray sources like pulsars, X-ray binaries, Seyfert galaxies, quasars etc. This is achieved by pointed mode observations of this sources with an array of three coaligned, collimated pointed mode proportional counters (PPCs). It in mutual operates anti-coincidence for significant reduction of background by cosmic ray and Compton interaction of gamma rays. 2 - 20 KeV 2”x2” 3 3 18 1.1 x I.1 cm2 25 km Mylar, 500 A Al coating Filling gas Ar + CH,, 800 torr Tab. 4 Parameters of the Pointing Proportional Counter (PPC) Energy range FoV No. of PPC No of layers per PPC No. of anode cells per layer Size of cell Entrance window

A further goal is the study of light curves and spectral evolution of transient and flaring X-ray sources and long term intensity monitoring of known binary X-ray stars and other bright X-ray sources. This is achieved by means of an X-ray-Sky Monitor (XSM) based on the principle of a pin hole placed above a position sensitive proportional counter in anticoincidence mode.

2 - 8 KeV 9o” x 90 o 1 cm2 16cm 32 proport. counters 1 x I cm2 x 32 cm 25 pm Mylar, Al coated Ar + CH,

Tab. 5 Parameters of the X-ray Sky Monitor


5.4 X-Ray


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The strategy of MOS development was a optimized hardware for registration of the influence of the atmosphere and target parameters and their variation (in time and space) on the measurable spectral signature. To maximize the retrievable information about a target, especially the determination of quantitative figures we found by simulations that for ocean remote sensing is necessary: - to use a large number of spectral channels in VWNIR and SWIR range for sampling the signature - to have a higher spectral resolution (Ah o 10 nm) as usually used in remote sensingmissions - to expand the spectral range to this extent, were you can get information about the target - to realize in the instrument a good signal to noise ratio and - to make a good and stable calibration for radiometric and spectral data. For the development of smart sensor algorithms is used a scheme of parameter retrieval from spectral high resolution measurements as shown in the diagram of Fig. 2. The used ,,hyperspectral linear estimator“ is derived and optimized for different remote sensing situations using lookup tables calculated by principle component analysis [2,3,4]. The new algorithm is applied directly to satellite radiances without atmospheric correction, makes use of as much as possible of multispectral information, is very fast in processing and stable in calculation convergence problems). (no Therefore it is tailored for more operational production of level-2 products. It is totally different to commonly used color ratio


Small Satellites for Earth Observation

algorithms which needs a separate atmospheric correction and water leaving spectral radiances reflectances) for retrieving water (or constituents via their bio-optical effect and color ratios. This new MOS Principal Component Inversion algorithm (MOS-PCI) must be tested, adapted and validated, at various regions of the oceans and coastal zones for a variety of different parameter combinations.

Fig. 2 Principle estimator




To provide this tests and the ,,regional tuning“ a long term validation program of the algorithm is now carried out by ground truth measurements at different test sites in European and Indian costal zones.

intercalibration with sensors from PRIRODA [l] and SeaWiFS Oceanographic applications of MOS data for retrieval of different water constituents in the open ocean and coastal zones Land applications of WiFS data for vegetation state and stress, land use and desertification, flood mapping Atmospheric applications of MOS-A data for validation of methods for correcting the atmospheric influence, estimation of aerosol and cloud parameters. Ecological monitoring for cloud/snow/ice discrimination and cloud type determination by the SWIR channels of WiFS and MOS-C Combined MOS-WiFS Data utilization Example of MOS and WiFS images and processed level 2 data images can be found at DLR-homepage


et. al., MOUPRIRODA - An Imaging VWNIR Spectrometer for Ocean Remote Sensing, SPIE Proceedings, Vol. 1937, pp 201-206, 1993 Krawczyk et. al., Investigation of Interpretation




Besides this special Ocean-algorithm problems the IRS-P3 mission covers several general aspects: -



Technological test of the PSLV launch vehicle, radioastronomy experiments using the X-ray payload Scientific remote sensing applications and algorithm development based on joint interpretation of MOS and WiFS data Pre-operational test of data processing and algorithm concepts.

The scientific part of the mission addresses significant where five major issues contributions are expected: - Calibration and validation of MOS and WiFS and its intercalibration, in-orbit

Possibilities of Spectral High Dimensional of Principal Measurements by Means Component Analysis - A Concept for Physical interpretation of Those Measurements, SPIE

Proceedings, Vol. 1938, pp 401-411, 1993 Neumann

et. al., A Complex Approach to Quantitative Interpretation of Spectral High Resolution Imagery,, Paper presented at ERIM’s Third Thematic Conference on remote sensing of Marine and Coastal Environment, Seattle, Washington, 18-20 September 1995, pp. II-641 - II-652 Krawczyk et. al., interpretation Marine


Potential of Multispectral Imagery,

Paper presented at ERIM’s Third Thematic Conference on Remote Sensing for Marine and Coastal Environments, Seattle, Washington, 1820 September 1995, pp. II-57 - II-68