From H-II Press kit

From: Yoshiro Yamada <>
Date: Thu, 16 Mar 1995 02:55:15 -0500
   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
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)
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
Specific         445(*3)    273(*3)       267(*3)          452(*3)


pitch and     Main engine  Movable nozzle               Gimbal(during
yaw axes      Gimbal                                    poweredflight
                                                        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
                 S-band       PCM/PSK                PCM/FSK.AM/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

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

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:


 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.


 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

Microgravity   : less than 0.0001 g

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

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
 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
               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

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 :
5-2-1 Yokodai Isogo-ku                    |
Yokohama 235, JAPAN                       |
Received on Thu Mar 16 1995 - 04:41:10 UTC

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