GMS-5 & SFU Mission and Launch Information (Vehicle H-II TF#3)

Material from NASDA Press kit "H-II Launch Vehicle No. 3" and NASDA 
Press Release March 18, 1995.

=========================================================================
This document is B44507 in NSSDC's Technical Reference File.
Added Note: GMS-5 is 95-011B, and SFU is 95-011A, launched 3/18/95.
=========================================================================

From:	AMES::"yamada@ysc.go.jp" "Yoshiro Yamada" 16-MAR-1995 03:08:16.39
To:	sardi@ncf.span.nasa.gov
Subj:	from H-II Press kit

   GMS-5/SFU launch is now Mar 18 (07:55-09:02 UTC), 1995.



The following  is an extract typed by Ms. Reiko Shindo, YSC.

----------------------------------------------------------------
From NASDA Press kit "H-II Launch Vehicle No. 3":


Table 1-2 Principal Specifications of the H-II Launch Vehicle Test 
Flight No.3
                            Overall vehicle
Name               H-II Launch Vehicle Test Flight No.3
Length      (m)    51
Diameter    (m)    4.0 (core vehicle)
Total weight(t)    278.9 (without payload)
Guidance system    Inertial guidance system
                   First   Solid rocket  Solid sub       second  Payload
                    stage   boosters(SRBs) boosters(SSBs) stage  fairing
Length      (m)    34.4       23.4          9.0           10.3    14.1
Diameter    (m)     4.0        1.8          1.1            4.0   5.1/4.1
Weight      (t)    98.2     140.8(2 units) 21.0(2 units)  16.8     2.1
                    (*1)
Propellant weight  86.3     118.3(2 units) 16.8(2 units)  14.0
Average thrust(t) Main      316.0(2 units) 69.0(2 units)  12.4
                  engine 86          (*2)                   (*3)
                        (*2)
Burn time   (s)  Main        94            66              497
                 engine 347
Propellant      LOX/LH2    Polybutadiene   polybutadiene   LOX/LH2
                           composite solid composite solid
                           propellant      propellant 
Propellant feed            Turbo pump                     Turbo pump
sustem
Specific         445(*3)    273(*3)       267(*3)          452(*3)
impulse(s) 

Attitude
control

pitch and     Main engine  Movable nozzle               Gimbal(during 
yaw axes      Gimbal                                    poweredflight
                                                        phase)RCS(during
                                                        coasting phase)
Roll          Auxillary                                 RCS:Reaction 
              engine                                    Control System

Radio         Guidance     Telemetry                (1)Guidance control
systems       control      package                     equipmet
              equipment    measuring                (2)Radar transponder
              Telemetry    epuipment                   (2 units)
              transmitter                           (3)Telemetry trans-
              (SHF.VHF)                                -mitter(UHF.VHF)
              (2 units)                                (2 units)
                                                    (4)Command Destruct
                                                       Receiver(2 units)    
    *1 Interstage section included
    *2 Sea level
    *3 In vacuum


1.3 Capability and launch sequence

 Table 1-4 shows the launch capability of the basic H-II rocket for
delivering a single satellite into geostationary orbit. Fig. 1-3 shows 
the sequence for delivering a geostationary satellite. Fig. 
1-5 and 1-6 show typical trajectories.
          
          Table 1-4 Launch Capability of the basic H-II rocket

   Orbit                    Example                    Payload
Geostationary transfer  Apogee altitude:36,000 km  Approx. 4 tons 
orbit                   Perigee altitude:250 km    (2.2 tons into
                        Inclination:28.5 deg       geostationary orbit)
Sun-synchronous orbit   Altitude1: 800 km           Approx. 4 tons
                        Inclination:99  deg
Medium Earth orbit      Altitude: 1000 km          Approx. 5 tons
                        Inclination:30  deg
Low Earth orbit         Altitude: 250 km           Approx. 10 tons
                        Inclination:30 deg
Earth escape trajectory  venus and Mars probe      Approx. 2 tons



          4. Geostationary Meteorological Satellite-5 (GMS-5)

4.1 Outlines

 The Geostationary Meteorological Satellite-5 (GMS-5) will provide 
meteorological satellite services, forming part of a network of 
meteorological satellite of Global Observing System (GOS) for the 
World Weather Watch (WWW) program planned by the World Meteorological 
Organization (WMO). As a member of WMO, Japan responded to the needs 
of WWW program by developing a geostationary meteorological satellite 
(GMS). In July,1977, GMS, known as "Himawari (a sunflower)", was
launched into geostationary orbit, approximately 36,000 km above the 
equator at 140 deg E longitude. Subsequently, GMS-2 was launched in 
August 1981, GMS-3 in August 1984 and GMS-4 in September 1989 to 
continue this meteorological satellite service.
 GMS-5, the successor of GMS/GMS-2/GMS-3/GMS-4, is to continue this 
meteorological satellite service and contribute not only to 
improvement of meteorological satellite service, but also to 
development of related technology. GMS-5 will be launched by the 
third  H-II test vehicle from the Tanegashima Space Center on February 
1, 1995.

4.2 Objectives

 GMS-5 has the following mission objectives.

(1) Observation by the Visible and Infrared Spin Scan Radiometer(VISSR)
 GMS-5 will observe the earth cloud cover location, the earth 
temperature such as ground, sea surface and cloud top temperature and 
atmospheric  water vapor distribution from geostationary orbit both 
in the visible and infrared spectra (earth image acquisition). 
Earth images are taken every 30 minutes by VISSR with 1.25 km visible
and 5 km infrared resolution. These images are transmitted an processed
on the ground to get the following data on:

  .severe weather such as typhoons and torrential rains.
  .cloud distribution and wind speeds.
  .sea-surface temperature.
  .atmospheric water vapor distribution.

(2) Weather Data Ralay

 Weather data sent by the Data Collection Platforms (DCP), which are 
installed mainly on ships, buoys and aircrafts, are relayed by the 
GMS-5 communication subsystem and received at the Command and Data 
Acquisition Station (CDAS) to be processed at the Data Processing 
Center (DPC). 

(3) Observation Data Relay

 Earth image data transmitted by the satellite are to be processed 
by the Stretched VISSR(S-VISSR) signals for lower transmission speed 
at CDAS. Then, the data are received at the Medium-scale Data 
Utilization Stations (MDAUS) via GMS-5. The earth data processed 
at DPC are received in the form of WERAX signals with 800 scan lines 
at the Small-scale Data Utilization Stations (SDUS) via GMS-5.

(4) Relay of Search and Rescue (SAR) Signal

 A relay test will be performed to transmit emergency signals (SAR 
signals*) from a distressed ship using GMS-5 as a relay satellite 
to search and rescue organizations.

*SAR signals, originated from the Emergency Locator Trensmitter (ELT) 
for aircraft and the Emergency Position Indicating Radio Beacon (EPIRB)
for ships are relayed by GMS-5 to the search and rescue organizations.



4.3 System Outline and Major characteristics

4.3.1 Outline

 GMS-5 is a spin-stabilized satellite with mechanical despun antennas. 
The configuration and characteristics of the satellite are based 
on those of GMS/GMS-2/GMS-3/GMS-4 and its major subsystems have adopted 
those of having successful flight experiences.

 The satellite is 444 cm tall (at launch), 354 cm tall (on orbit) and 
215 cm in diameter. It weighs 746 kg at the separation from the H-II 
vehicle and 345 kg on orbit. It consists of a despun earth-oriented 
antenna assembly and a spinning section rotating at 100 rpm. A despun 
antenna provides necessary gain for communications between the satellite 
on orbit and the ground stations. The despun section is controlled by 
earth and sun sensors of the spin section. A noncontacting RF rotary 
joint, mounted coxially within a Despin Bearing Assembly (DBA), feeds RF 
signals between despun antenna snd a spinning equipment shelf.

 The spinning section consists of the forward and the aft. The forward 
assembly carries electronics devices including VISSR, and the aft has 
an Apogee Kick Motor (AKM) and separation hardware. the aft will be 
jettisoned from the satellite body after AKM burnout.

 The satellite mission life and design life are 5 years. Redundancy of 
critical functions to accomplish the mission is provided. The solar 
panel power produces 291 W with the margin of about 22 W at the end of 
5 years (summer solstice).
 The configuration of the satellite is shown in Fig.4-1 and its 
functional blocks in Fig.4-2.


4.3.2 Major Characteristica

 The major characteristics of GMS-5 is illustrated in Table 4-3.

                 Table 4-3. Major Characteristics(1/2)


  Items                 Major characteristics

Dimensions      Diameter : 215 cm
                Height   : 444 cm (with AKM)
                           354 cm (without AKM))
Weight           746 kg launch
                 345 kg on orbit
Life             Mission life : 5 years
                 Design life : 5 years
Reliability      5 years  : 0.0699 
Orbit            Geostationary : 140 deg E longitude
                 Station keeping : +-0.5 deg (E-W)
                                 : +-1 deg (S-N)
Launch           Vehicle   : The third H-II test vehicle
VISSR and VDM                      Visible            Infrared
System           Spectral band    0.55-0.9          10.5-11.5 micro-m  
                                     micro-m        11.5-12.5 micro-m
                                                     6.5-7.0 micro-m
                 Resolution       1.25 km               5 km

Telemetry and                    Telemetry             Command
Command          Capacity      256 x 2 channels      pulse : 254
                                                     serial:   3
                 Modulation   USB  PSM/PSK/PM        PCM/FSK.AM/PM
                              FM/PM,FSK/PM 
                 S-band       PCM/PSK                PCM/FSK.AM/PM
                              FM/PM,FSK/PM
                 Bit rate     250 BPS                128 BPS




                 Table 4-3. Major Characteristics(2/2)


  Items                      Major Characteristics

Systems                    S-band          UHF           USB
                Uplink    2026-        402-           2100.164
                (MHz)       2034.974     406.05 
                
                Downlink  1681.6-      468.875-       2280.721
                (MHz)       1698.35      468.924
                
                Antenna     Parabola     Helical      Bicone
                Coverage    Global       Global       Omni 
                G/T         -22 dB/K     -23 dB/K     -50 dB/K

                EIRP 40dBm+-1.5dB(+-6.5 deg)*  46.0dBm+2dB 27.5dBm+-
                    34dBm+-1.5dB(+-6.5 deg)**  -1.5dB      1.5dB
                    56.5dBm+-1.5dB(+-6.5 deg)*** 
                    44dBm+-1.5dB(+-6.5 deg)****
                    54.5dBm+-1.5dB(+-6.5 deg)*****

                   *S-band telemetry    **ECPR/SAR
                 ***VISSR data        ****Trilateration
               *****WEFAX

Control        Spin-stabilized  (spin rate 100+-1rpm)
               Antenna pointing: +-0.5 deg(N-S Attitude Control)
                               : +-0.39 deg(E-W Precision Sun Reference)

Electric       Solar Array Power : 290 W (EOL Summer Solstice)
  Power        Batteries         : 4.8 A-hr x 2 sets

RCS            propellant   : Hydrazine
               3 tanks, 2 axial and 4 radial thrusters

AKM            Thiokol TE-M-616-5 (STAR-27)

Thermal        Passive
Control         (heaters used in RCS, AKM, Despin motor and VISSR)





                  5.THE SPACE FLYDE UNIT (SFU) PROGRAM


5.1 Outline

5.1.1 Purpose

 The SFU is a modular, resuable orbital platform that, unlike 
conventional space systems designed only for a specific mission, 
allows multiple users access to space for various observations and 
experiments.

5.1.2 Schedule

 The SFU will be one of two payloads launched by the Third H-II Test 
Vehicle from NASDA's Tanegashima Space Center sheduled for February 
1995. The other payload is the Geostationary Meteorological 
Satellite-5 (GMS-5). The SFU will be be placed in orbit at 
approximately 330 km attitude. It will then boost itself to higher 
orbit of 500 km where it will perform multiple observations and 
experiments in space for a period of several months.

5.1.3 Retrieval

 In cooperation with the National Aeronautics and Space Administration 
(NASA), the SFU will be retrieved during mission STS-72 by Space Shuttle 
Endeavour scheduled to launch in November 1995.

5.2 Responsibilities of the Participating Organizations

    Each organization has the following responsibilities for 
    the SFU program:

5.2.1 STA/NASDA 

 Responsible for
 .launching the SFU on the Third H-II Test Vehicle,
 .developing the Exposed Facility Flyer Unit(EFFU) which is a flight 
  demonstration model of the Japanese Experiment Module (JEM) for the 
  international space station, 
 .cooperating tracking and control sevices during launch and retrieval 
  of the SFU.

5.2.2 MDE/ISAS

 Responsible for 
 .managing and integrating all SFU systems,
 .interfacing with NASA,
 .coordinating all SFU systems,
 .developing bus equipments including navigation, guidance and 
  control subsystems,
 .developing experimental equipments including the Infrared Telescope 
  in Space (IRTS), which is a cooperative project with NASA,
 .providing tracking and control during the SFU mission at 
  the Sagamihara Operation Center.

5.2.3 MITI/NEDO.USEF

 For the SFU, USEF is the actual hardware developer under the 
 authority of MITI and NEDO.
 Responsible for
  .developing three different electrical furnaces for material 
   science experiments in the microgravity enviroment of space,
  .developing power, structure & thermal control, and data 
   processing subsystems,
  .developing facilities for mission operations and control.


5.3 Major Features

 Mission orbit : Altitude 330 km after launch
                          500 km during mission and in phaserepeating 
                          orbit 300-500 km at retrieval

                         Inclination 28.5 deg.

 Dimension     : Main Body 4.46 m diameter x 2.8m height (launch or 
                                 retrieval configuration)
                           24.4 m length (configuration with fully 
                            extended solar array paddles) x 2.4 m(w)

weight         : Launch    4.0 tons
                 Retrieval 3.2 tons

Attitude       : Sun pointing/3-axis stabilized
 Control

Microgravity   : less than 0.0001 g
 Condition     


5.4 Structure

 The major structure of the SFU cosisits of an octagonal-shaped 
aluminum alloys truss with detachable box-like payload units (PLUs) 
in which experiments are installed. The core system including attitude 
control, communication and power subsystem stowed in bus units, and 
on-board experimental equipments stowed in the PLUs. The interfaces 
for power, data and communication subsystems between the PLUs and the core 
subsystems are standardized to enable the use of different experimental 
equipments. The SFU solar array paddles, orbital change thruster, 
antennas and attitude control sensors are assembled directly on the 
main structure.
 Fig.5-1 shows the launch and in orbit configurations of SFU. In orbit 
solar array paddles are fully deployed. Fig. 5-2 shows the SFU core 
system.

5.4.1 Navigation, Guidance & Control

 The SFU controls its attitude in a three-axis stabilization. During 
nominal operation, the SFU maintains a sun-pointing mode using solar 
and earth sensors. Its attitude is controlled by reaction wheels, 
magnetic torquer and a Reaction Control Subsystem (RCS) with 3N 
thrusters.
 By using Global Positioning Satellite System (GPS), the SFU can 
automatically recognize its orbital position and velocity without 
any support of ground tracking system. It can also place itself in 
a planned orbit by firing the orbital change thruster after a command 
from the ground station.

5.4.2 Orbit Change Thruster

 The SFU has a unique Orbit Change Thruster (OCT) system. The OCT 
consists of eight 23N thrusters and hydrazine propellant tanks of 
650 kgcapacity. The OCT will be used by the SFU to ascend to its 
missionorbit after launch and descend to the shuttle rendezvous 
orbit forretrieval.

5.4.3 Communication and Data Management

 This system consists of an S-band communication subsystem and a data 
management subsystem with 16-bit micoprocessor. The SFU can transmit 
core system data to the Space Shuttle at a rate of 1 kbps. The SFU 
can transmit telemetry data to the ground station at rates of 1 kbps,
16 kbps and 128 kbps, sending the core system and experiment data 
in real time as well as playback provided by data recorders.
The recorders store 256 time-line commands and 128 timer-commands. 
In addition, four monochromatic still cameras are mounted to the SFU 
to monitor the solar array paddles and other components.


                   Table 1, System Performance of SFU

Dimension      4.46 m dia. x 3.02 m L Modified Octagonal Modularized 
               Type Structure

Weight         Total Mass.     : 4 ton at launch
               Mission Mass.   : 1 ton on orbit

Power          Solar Array Paddle
                  Power Generation  : More than 2.7 KW (BOL)
               Mission Electric Power  : 850W (Averege on Orbit)
               Total Power Supply : More than 1.4 KW (Mean Value)
               Bus Voltage
                (Floating for Eclipse) : 32.5V-51.5V DC (at load)
               Battery   : NiCd 19AH x 4 ea.
               Bettery Depth of : Less than 30% (Mission steady phase)                    Discharge     : Less than 60% (BOL/Retrieval phase)

Navigation     Navigation : Autonomus Navigation System with GPS 
Guidance                    and IMU(Pointing Accuracy+-50m/
and                         Velocity+-0.1m/sec.)
Control        Guidance   :Guidance with On-board Computer S/W
Orbit          Attitude
Control        Control :Three-Axis Stabilize Pointing Accuracy+-1 deg.                 Accuracy : Position+-100m, Velocity+-0.1m/sec.

Communicaton   Communication Frequency:(UP Link 2084, 4MHz/Down 
and Data                                Link 2263.6MHz)
Management     Function : Telemetry, Command and Ranging
               Telemetry data Rate:1Kbps(Fixed H/K),16Kbps(High 
                                   Speed Packet),128Kbps(Fixed H/K 
                                   and H/L Packet)
               Command Bit Rate : 1Kbps
               Command Memory Capacity : 256 Timeline Command,
                                         128 timer Command
               Data Recorder : 80Mbit(High Speed Packet Data)
                               40Mbit (Low Speed Packet Data)
               Main Comupter/Experiment Compuler:
                             12ch Intelligent Command, Packet 
                             Data Telemetry
               Monitor TV Camera: Monochromatic Still Image

Reaction       Thruster Type : Monopropellant Hydrazine Thruster
Control        No. of Thruster: 3N x 12/23N x 4(for RCS),23N 
and Orbit                       x 8(for OCT)
Change         Propellant Tank Capacity: 100 kg(for RCS),
                                         650kg(for OCT)


Structure      Main Structure: Aluminum Alloy Truss Structure
               STS Interface: 4 Longeron Trunnions and 1 Kell 
                              Trunnion
               H-II Interface: Aft Ring of Main Structure with 
                               4 Separation Bolts
               Stiffness : Launch Vehicle Lateral Axis > 10 Hz
                           Launch Vehicle Thrust Axis > 30 Hz
                           STS Support Condition > 6.4Hz
               But Unit : 2Units, Machine Aluminum Alloy 
                          Panel/Heneycmb Sandwich panels Structure
               Maximum Payload Unit : 200kg

Thermal        Exhaust Heat: 160-360W/Each Unit Box
Control        Temperature Control Range: 0 deg.~40deg.C with 
                                          Thermal Louver,Heat Pipe 
                                          and Passive Thermal Control
Acceleration   Less than 0.0001 g
Environment    


5.4.4 Electric Power

 The electric power system consists of two solar array paddles 
with 28,000 sillicon cells (2 x 4 cm), and four nickel cadminum 
batteries. The paddles are stowed for launch and deployed for 
operation. The power output is more than 2.7 kw at the beginning 
of life. The power system provides the 850w stable power to payload 
experiments and the necessary power to the core system for the 
overall operation of the SFU.

5.5 Operation and Retrieval

 The SFU will be launched in February 1995 by the Third H-II Vehicle.
The SFU will be placed at an altitude of 330 km with an orbital 
inclination of 28.5 deg. The operation of the SFU is shown 
in Fig. 5-3. After being injected into the orbit, the SFU will 
take a sun-pointing mode and deploy its solar array paddles.
After intitial checks, the SFU  will boost up to its operational 
orbit of 500 km using its Orbit Change Thruster (OCT). Because of
aerodynamic drag, the SFU will have to reboost regularly to maintain 
its proper orbit.  During reboosts, the SFU takes and earth-
pointing mode. Following a test period of approximately 7 days, 
on-board experiments will be conducted through time-line commands 
for about 4.5 months.  After completion of their missions, the SFU 
will descend to a phase repeating orbit (orbit repeated same pattern 
relative to ground position) of about 485 km altitude, where it will 
stand by for several months. The SFU will then rendezvous with 
the Space Shuttle at the altitude which will be arranged before 
the Space Shuttle launch.
 When the Space Shuttle is within approximately 20 km from the SFU, 
a communication link is established between the two spacecraft.
The Sagamihara Operation Center (SOC) will issue commands to the Space 
Shuttle which will in turn command the SFU to retract the solar array 
paddles and to go into a safety mode (the securing of all potential 
hazards to personnal, equipments or operations). If the solar array 
paddles do not retract properly, the SOC will issue a command to 
jettison the paddles. Then, the Space Shuttle will make a final 
approach to the SFU, capture the SFU with its robotic arm, and place 
it in the Shuttle cargo bay.  The SFU will switch off the SFU batteries,
and complete its mission before returning to earth.

5.6 Tracking and Control System

  The major ground station for the SFU mission is ISAS's Kagoshima 
Space Center (KSC).  The Okinawa Tracking and Data Acquisition Station 
(OTDS) is the backup ground station during the H-II launch and Space 
Shuttle retrieval phases. OTDS may slso be used as a backup station 
for KSC during the mission phase.  NASA's Deep Space Network stations 
(Goldstone, Canberra,Madrid) and Chile's Stantiago will be in service 
during the launch and retrieval phased (see Fig.5-4).
 ISAS's Sagamihara Operation Center(SOC) will transmit commands and 
monitor telemetry. When the SFU and the Space Shuttle will begin 
direct communication, commands will be given by SOC and the Space 
Shuttle crew, and telemetry data will be  monitored by the Space 
Shuttle crew, NASA's Mission Control Center, and SOC.


-------------------
Yoshiro Yamada                            | tel    : +81-45-832-1166
Astronomy Section                         | fax    : +81-45-832-1161
      Yokohama Science Center             | e-mail : yamada@ysc.go.jp
5-2-1 Yokodai Isogo-ku                    | 
Yokohama 235, JAPAN                       |

=========================================================================
=========================================================================

From:	AMES::"yamada@ysc.go.jp" "Yoshiro Yamada" 19-MAR-1995 02:03:11.96
To:	sardi@ncf.span.nasa.gov
Subj:	H-II launch info


From NASDA PRESS RELEASE:

LAUNCH OF SFU & GMS-5 / H-II LAUNCH VEHICLE TEST FLIGHT NO. 3
-------------------------------------------------------------

06:00 PM(JST)
March 18, 1995
NASDA H-II TF#3
LAUNCH TEAM

The National Space Development Agency of Japan (NASDA) launched
the third H-II Launch Vehicle test flight (H-II TF#3) from
Tanegashima Space Center with the Space Flyer Unit (SFU) and
Geostationary Meteorological Satellite-5 (GMS-5) at 05:01:00 JST
(=UTC+9h) PM on March 18, 1995 at the launch azimuth of 97 degrees.

The flight of the first and the second stages was normal, and the
guidance and control system functioned as planned. The SFU was
separated approximately 13 minutes 16 seconds after the launch,
and GMS-5, at 27 minutes 47 seconds after the launch. All flight
events of H-II TF#3 were completed.


INJECTION INTO ORBIT OF THE SPACE FLYER UNIT (SFU)
--------------------------------------------------

March 18, 1995
NASDA HQ, TOKYO

The data from SFU after the first revolution was received at 06:37 PM
(JST) on March 18 at the Sagamihara Operation Center, ISAS.

The data confirmed that SFU was injected into the proper orbit.

SFU will deploy the solar array paddle and its status will be
confirmed around 08:20 PM on March 20 at the Sagamihara Operation
Center, ISAS.


FLIGHT CONDITION OF THE GEOSTATIONARY METEOROLOGICAL SATELLITE (GMS-5)
----------------------------------------------------------------------

11:00 PM
MARCH 18, 1995
NASDA GMS-5/H-II TF#3
TRACKING & CONTROL TEAM

NASDA launched GMS-5 "HIMAWARI-5" at 05:01 PM (JST) on March 18, 1995
from Tanegashima Space Center. The data from the satellite after the first
revolution was received at 06:25 PM on March 18 at Madrid Tracking and
Data Acquisition Station.

The Data calculation confirmed that "HIMAWARI-5" was inserted into
the first transfer orbit as planned.

The first transfer orbit is as follows:

Apogee Altitude: 36669 km
Perigee Altitude: 329 km
Orbit Inclination: 25.1 degree
Period: 10 h 50 m

The command for the apogee kick motor firing will be transmitted at
09:57 on March 19 from the Tracking and Control Center at Tsukuba Space
Center.


SOLAR ARRAY PADDLE DEPLOYMENT OF THE SPACE FLYER UNIT (SFU)
-----------------------------------------------------------

March 18, 1995
NASDA HQ, TOKYO

The Sagamihara Operation Center, ISAS confirmed that the SFU normally
completed the solar array paddle deployment at 07:25 PM (JST) on
March 18, 1995. The SFU is functioning normally after the solar
array paddle deployment. 


INJECTION ORBIT DETERMINATION OF THE SPACE FLYER UNIT (SFU)
-----------------------------------------------------------

March 18, 1995
NASDA HQ, TOKYO

Data calculation confirmed that SFU was placed in the designated
orbit.

The orbit determination is as follows:

Apogee Altitude: 336 km
Perigee Altitude: 322 km
Orbit Inclination: 28.45 degree
Period: 1 h 31 m

SFU functions normally.

SFU will maneuver to its mission orbit by firing its Orbit Change
Thruster (OCT) three times.
SFU will complete orbit transfer to its mission orbit around
March 21.

After orbit transfer completion, the SFU initial function testing
will be started.


-------------------
Yoshiro Yamada                            | tel    : +81-45-832-1166
Astronomy Section                         | fax    : +81-45-832-1161
      Yokohama Science Center             | e-mail : yamada@ysc.go.jp
5-2-1 Yokodai Isogo-ku                    | 
Yokohama 235, JAPAN                       |