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 |