SmallGEO (Small Geostationary Satellite Platform) Initiative / HAG1 Mission
SGEO (or SmallGEO) is an ESA and DLR funded project with the objective to develop a general-purpose small geostationary satellite platform followed by a subsequent mission which will enable European industry to play a significant role on the commercial telecommunications market for small platforms. The contract is an element of ESA's ARTES-11 (Advanced Research in Telecommunications Systems) program which calls for: 1) 2) 3)
• The development of a generic Small Geostationary Platform capable of supporting a payload mass of up to 300 kg, a payload power of up to 3 kW, and a lifetime of up to 15 years.
• The second part involves the development and launch of a Small Geostationary Satellite and associated mission to provide flight qualification as well as in-orbit demonstration for the platform.
The ultimate goal of SGEO is the establishment of a general purpose, small geostationary satellite product, which will be competitive in the commercial telecom market for small platforms and which may also be used for institutional applications.
In November 2008, ESA signed a contract with the respective industrial primes, namely: OHB- System AG of Bremen, Germany and Hispasat S. A. of Madrid, Spain. The core team of companies (consortium) which is jointly developing the platform and will subsequently commercialize it, includes OHB-Sweden AB [formerly SSC (Swedish Space Corporation)], Oerlikon Space of Switzerland (note: in 2009, Oerlikon Space was acquired by RUAG Space of Switzerland), OHB-System AG of Bremen and its Luxembourg-based subsidiary LuxSpace.
The activities/studies towards SGEO started already in 2005. The consortium was formed in 2007 and has been working on the preliminary development of SGEO. The main goal is to qualify the platform on orbit with a mission representative of the target (commercial) application. 4) 5)
As far as the SGEO mission is concerned, Hispasat, the leading operator for the Spanish- and Portuguese-speaking markets, has put together the necessary industrial team to implement all the mission elements. OHB-System AG is acting as the project prime contractor. Tesat GmbH of Backnang, Germany will be the Payload Prime and will be responsible for the integration of the Transparent Repeater (produced by Tesat) and the REDSAT payload.
The resulting satellite, Hispasat-AG1 (Advanced Generation satellite 1), or simply HAG1, will be placed into a geostationary orbit at an altitude of ~36000 km, where it will supply Spain, Portugal, the Canary Islands and South America with multimedia services. It will serve as an important and distinctive asset in Hispasat’s satellite fleet, delivering:
• A communications capacity of up to 24 transponders in Ku-band and 3 transponders in Ka-band
• A highly innovative REDSAT payload, a project led by Spanish industry. The REDSAT payload will include an active direct radiating array antenna providing four reconfigurable uplink beams in Ku-band and an advanced onboard processor offering four regenerative channels of 36 MHz. The goal is to supply Spain, Portugal, the Canary Islands, and South America with broadband multimedia services.
Development of the REDSAT payload will be led by Thales Alenia Space España, which will coordinate a group of mainly Spanish companies with Thales Alenia Space España TAS-E) being responsible for the OBC (Onboard Computer), and EADS CASA responsible for the reflector and DRA (Dual Receive Antennas).
On Sept. 30, 2011, ESA/TIA (Telecommunications & Integrated Applications) and OHB System AG signed an extension to the contract for development of the SmallGEO geostationary satellite platform. The added features in this extension contract will optimize the SmallGEO platform for a number of different commercial satellite services beyond the Hispasat AG1 mission. 6) 7)
The SGEO bus has been selected as the baseline platform for the EDRS (Data Relay Satellite) of ESA. The contract for the first commercial mission, using the Luxor bus (OHB refers to the commercial version of the SGEO bus as “Luxor”), was signed in November 2008 between ESA and Hispasat S. A. of Spain. In June 2009, OHB-System AG and Hispasat S. A. signed in turn a contract for the manufacture of Hispasat-AG1. Hispasat-AG1, due for launch in 2014, will be the first satellite to use Europe’s new SmallGEO platform, developed through a PPP (Public–Private Partnership) between ESA and Germany’s OHB.
The Hispasat-AG1 satellite includes the advanced communications payload REDSAT that will permit a more efficient use of its power and features. The REDSAT regenerative system is fully compliant with two European standards: DVB-RCS (Digital Video Broadcast-Return Channel Satellite) and the recently issued DVB-S2 (Digital Video Broadcasting-Second Generation). 8) 9)
The Hispasat-AG1 spacecraft will be launched into GTO. A bi-propellant system on-board will provide the injection into geosynchronous orbit. Final orbit transition and placement is however performed with the EP (Electric Propulsion) system. Electric propulsion is then used for all nominal station-keeping and momentum management for the entire lifetime of 15 years (Ref. 13).
Figure 1: Artist's rendition of the generic SGEO spacecraft in orbit (image credit: SGEO consortium)
SGEO offers a highly flexible and modular geostationary platform, able to accommodate a wide range of payloads and a fast recurring delivery time of 18 to 24 months. ITAR free subsystems & components based on European technologies are envisaged. The platform design provides the capability for direct injection into a geostationary orbit as well as injection into GTO (Geostationary Transfer Orbit), both configurations offer a high mass efficiency. In the GTO configuration, the platform will have a launch mass of ≤ 3200 kg.
• Electrical propulsion is being used for all stationkeeping functions.
• An AEM (Apogee Engine Module) for GTO transfer with a bipropellant
• All band communication payloads (P, L, S, C, X, Ku, and Ka-band)
• Payload capacity up to 400 kg & 3.5 kW.
• 3-axis stabilization
• Spacecraft 4.5 kW EOL average power
• Platform wet mass < 1700 kg (GEO), < 3200 kg (GTO)
• Mission lifetime of up to 15 years
• Spacecraft dimensions of 2.3 m x 1.9 m x 2.5 m (height)
The satellite structure has been designed to be compatible with two different launch scenarios: a direct launch into GEO orbit and a GTO injection with successive transfer phase to GEO. The architecture difference between these two scenarios lies in the implementation of the CPPS (Chemical Propulsion Subsystem)) for GTO transfer. The EPPS (Electrical Propulsion Subsystem) is mainly used for GEO operations; it is based on the combination of the EPTA and the CGTA, both fed with xenon propellant (Ref. 19).
The SGEO platform was born to support mainly telecommunication missions, but its modular structure is flexible to accommodate different types of payloads. The actual configuration can board payload mass up to 400 kg and payload power up to 3.5 kW at end of life. The design lifetime of the platform is 15 years, after this time the satellite will be transferred to a disposal orbit. The payload capacity for SGEO is extended by the innovative electric propulsion subsystem, as well as the possible launch opportunities. Indeed the satellite is compatible with the two launch scenarios of a standard geo-transfer orbit and a direct-to-geosynchronous orbit injection.
Figure 2 depicts the SGEO modules. The satellite design follows a modular approach: the core platform module, the propulsion module and the repeater module with the antenna assemblies are separately integrated and then mated during the last satellite integration phases.
The payload antennas are deployed from the East and West panels, the boresights are parallel to the positive z-axis (nadir). The y-axis points to South and the x-axis is oriented positive towards the orbital velocity vector. The apogee engine nozzle protrudes from the negative z-axis (zenith) side of the satellite. - The repeater module hosts the payload, manufactured by Tesat.
The central tube is the main structural element; the propulsion module is built around it. The CPPS propellant tanks are installed inside the central tube, while the helium and xenon tanks are symmetrically placed around it and as close as possible to the z-axis in order to minimize the excursion of the center of gravity.
The spacecraft structure, made of aluminum and composite materials in order to minimize the platform weight, develops from the central tube. The core platform module accommodates the electronic units which belong to the satellite subsystems and the reaction wheels. Heat pipes run across the north and south panels to remove the heat dissipated by the units. The batteries are accommodated on the anti-nadir panel.
Figure 2: Modular concept of SGEO (image credit: OHB System)
Figure 3: Conceptual view of SGEO subsystem electrical interfaces (image credit: SGEO consortium)
AOCS (Attitude and Orbit Control System):
The AOCS architecture is a three-axis stabilized system using a reaction wheel assembly for attitude control, star trackers for attitude control, and the EPPS (Electric ProPulsion Subsystem) for orbit maintenance. The AOCS design is characterized by a number of advances in technology beyond traditional telecom satellite designs. Perhaps the largest deviation from a traditional design is complete reliance on EPPS for orbit control. Angular momentum management of the reaction wheels relies solely upon EPPS in the nominal modes. The EPPS is not used in the safe modes and therefore a cold gas system is included on-board. The cold gas system is very efficient in terms of mass because it uses Xenon, the same fuel as used by the EPPS. The AOCS and EPPS are supplied by OHB-Sweden. 12) 13) 14) 15)
Another advance is the reliance upon APS-based star trackers. APS (Active Pixel Sensor) star trackers have a number of advantages over their CCD-based cousins in terms of robustness. As backup as well as for failure detection and for GTO, medium performance gyros are being used. The traditional fine sun sensor is simplified to a fault tolerant system of solar cells with 4π steradian coverage giving low, but more than adequate, accuracy.
The AOCS commands the following actuation devices based upon inputs from the following sensors: SPS (Sun Presence Sensors), GYR (gyros), and ST (Star Trackers).
- RWA (Reaction Wheel Assembly)
- Cold gas thrusters (XREA)
- EPPS (Electric ProPulsion Subsystem)
- Vernier thrusters (VREA)
- LAE (Liquid Apogee Engine)
- SADM (Solar Array Drive Mechanism)
Sun sensor: The chosen sun sensor is supplied by Bradford Engineering. It has redundant detection elements each with a near hemispherical field of view. The outputs approach a cosine function of the angle of incidence of the sunlight onto the detection elements. For this reason the redundant device is named CoSS (Cosine Sun Sensor). The design exploits the extensive heritage of dual chip devices used by TNO in its Sun Acquisition Sensor flying in a variety of space programs.
Figure 4: CoSS sun sensor (image credit: (Bradford Engineering)
Gyrometer: The baseline gyrometer is the Honeywell MIMU. Two units are used in cold redundancy. The gyro should support the Safe modes. Its reliability is not high enough to ensure continuous operation during the whole 15-year long mission, but its performance is good enough to enable bridging of ST outages, up to 10 minutes, even at start-up. The gyro is also used as a support to the ST measurements during more demanding mission phases (GTO transfer, platform commissioning, station acquisition, orbit relocation and graveyard orbit insertion).
Figure 5: The AOCS architecture (image credit: SSC, SGEO consortium)
Star Tracker: The ASTRO-APS of Jena-Optronik system of redundant star trackers has been selected by the project. The STAR1000 detectors (Cypress) of the ASTRO-APS device represent the most radiation hard CMOS APS (Active Pixel Sensor) on the market. The APS technology provides an individual readout scheme for each pixel, thereby avoiding “blooming” of charge over to neighboring pixels. This improves performance with bright objects in the field of view, e.g. the moon. The centroid calculations are done on the same integrated circuit resulting in tracking rates, e.g. 10 Hz, that are much faster than CCD-based star trackers. Built-in software algorithms automatically identify and compensate for anomalies such as white spots. A Peltier cooler is included. The star sensor is a single unit with camera head, electronics, and a shorter baffle than on CCD-based star trackers. 16)
Figure 6: Illustration of the ASTRO-APS star tracker (image credit: Jena-Optronik)
GNSS receiver experiment: SGEO will carry a GNSS receiver as an experiment onboard. The experiment will demonstrate in orbit the feasibility to use GPS signals in 36.000 km altitude for satellite position determination. Use of the Mosaic GPS receiver of Astrium. The device has a mass of 3.9 kg, a power consumption of 10 W, and a size of 272 mm x 284 mm x 92 mm. The expected accuracy in GEO is 150 m (3σ). 17)
Figure 7: Flight unit box of redundant Mosaic GNSS receiver (image credit: Astrium)
EPS (Electrical Power Subsystem):
The SGEO solar array is provided by Astrium GmbH and is mechanically composed of fully qualified and flight proven EuroStar-3000 & Alphabus equipment and components. The electrical components are qualified in the frame of E2000+, ARTES 3 EuroStar-3000 second generation (SG), Alphabus and Galileo. The generic SGEO solar array is based on panels of size of 2730 mm (length) x 2216 mm (width). All panels of the solar array are equipped with 3G triple junction solar cells with integrated bypass diode. 18)
The EPS is based on a 50 V regulated main bus powered by GaAs triple-junction solar cells disposed on six steerable solar panels. The 9 m solar arrays are deployed after orbit acquisition. The main bus is capable to distribute more than 6.5 kW power at BOL to the spacecraft, roughly divided in 4.1 kW to the payload and 2.4 kW to the platform in nominal mode. At EOL, the payload available power reduces to 3.5 kW including margins.
SGEO propulsion systems:
The architecture of SGEO has been designed to account for a direct GEO injection or a GTO injection. The sole difference between the launch types lies in the implementation of a chemical propulsion system for GTO transfer. 19)
SGEO is a hybrid version where the spacecraft will rely on the Electric Propulsion Subsystem (EPPS) for all routine propulsive tasks once in GEO (Geostationary Orbit) after orbit raising with chemical propulsion. The orbital maneuvers include station acquisition; station keeping for a service life of 15 years; momentum management during GEO mission using reaction wheels; intermediate repositioning, if required; and transfer to graveyard orbit at end of life. 20)
For this purpose, the EPPS baselined on the SGEO platform comprises two independent (redundant) branches of four thruster units each, where the eight thrusters are hard-mounted onto the satellite walls, with no thrust pointing mechanism. This is in similar fashion, for example, to the Russian GALS, Express, Sesat, Kazsat, and Yamal family of telecommunications satellites.
• CPPS (Chemical ProPulsion System): A bipropellant MMH/NTO (Mono-Methyl Hydrazine / Nitrogen and Tetroxide) system for GTO-GEO insertion. CPPS consists of one 400 N LAE (Liquid Apogee Engine) and 4 main + 4 redundant reaction control (RCT) 10 N bipropellant thrusters for the stabilization of the satellite during transfer. The CPPS is only used for orbit transfer from GTO to GEO, nominally with four LAE burns of varying length. The Vernier thrusters are used for attitude control before, during, and after the LAE burn. In case of LAE failure, the subsystem is dimensioned for transfer with Vernier thrusters only. Upon successful transfer, the subsystem can be passivated with two opposing vents.
• Two different EPPS (Electric Propulsion Systems) for GEO orbital maneuvers
• A Xenon cold gas system for detumbling and safe mode operations, referred to as CGTA (Cold Gas Thruster Assembly). The EPPS includes in addition eight fixed Cold Gas thrusters, operating with a common Xenon supply.
EPPS (Electric ProPulsion Subsystem):
EPPS is being developed at OHB Sweden (EPPS is based on the Smart-1 mission heritage): The objective of the EPPS is to perform all on-orbit maneuvers (stationkeeping, intermediate repositioning, transfer to graveyard orbit) including momentum management during all phases except for the transfer from GTO to GEO. 23) 24) 25)
EPPS is a fully redundant system consisting of two branches with 4 EPPS thrusters each is used. The thruster layout as illustrated in Figure 10 is chosen such that the center of mass (CoM) movement during the lifetime is bounded by the eight thrust vectors, with an additional margin to ensure that torques can always be generated for wheel unloading.
The EPPS consists of the following subassemblies:
- XTA (Xenon Tank Assembly). XTA consists of two 60 liter composite-overwrapped titanium liner tanks. Each tank will house 110 kg of Xenon at 186 bar.
- PSA (Pressure Supply Assembly). The primary function of the PSA is to decrease the Xenon storage pressure to the operational pressure range required both by the EP thrusters and the cold gas thrusters. 26)
- CGTA (Cold Gas Thruster Assembly). The CGTA is used for initial de-tumbling after separation from the launch vehicle and for safe mode. The two groups of EPPS thrusters are powered by a separate electronic unit (PPU/TSU), which allows to operate any of the four thrusters within its group as commanded by the AOCS.
- EPTA (Electric Propulsion Thruster Assembly). The EPTA and the CGTA will be fed with Xenon stored in two tanks by the same propellant supply assembly. Its main function is to decrease the Xenon pressure from the main inlet tank pressure (150 bar BOL) down to a regulated pressure of 2.2 bar at the outlet. The Xe flow rate to each EPPS thruster is regulated by the XFC (Xe Flow Control) unit.
The EPTA includes two EP types. Four fixed HEMPT (High Efficiency Multistage Plasma Thruster) and the PSCU compose the nominal (EPTA1) branch, while the redundant branch (EPTA2) contains four SPT-100, the PPU, the ETSU and FUs. The XTA and the PSA are shared with the CGTA.
The high pressure of the xenon contained in the XTA is decreased to the values required by the thrusters via the PSA. The PSA provides pressure regulation and a central distribution system to multiple thruster branches; it consists of the fluidic hardware (PRP) produced by IberEspacio and the electronic control unit (SCE) provided by Crisa Astrium. The actual design reduces the propellant pressure from 186 bar (BOL) to 2.2±0.11 bar at the output. The PSA is designed around a bang-bang regulator mechanism to fulfill the different mass flow rate requirements of the propulsion assemblies. The architecture is similar to the one used on SMART-1.
Figure 9: Functional block diagram of the SGEO EPPS (image credit: OHB Sweden)
The thrusters are mounted in pairs as depicted in the Figure 10 with thrust directions symmetrically oriented around the nadir vector. In nominal SK (Station Keeping) operations, each thruster has thrust vector components in the directions orthogonal to the orbital plane and tangential to the satellite velocity vector in inertial space.
The HET (Hall Effect Thruster) assembly is procured from Snecma (France) and consists of four flight proven SPT-100 thruster modules including XFC (Xe Flow Controller) and two cathodes, one PPU (Power Processing Unit) and one ETSU (External Thruster Switching Unit). The PPU includes an internal TSU (Thruster Switching Unit) allowing the selection of two thrusters. In order to switch between four thrusters, the ETSU has been added. The HET features for SGEO are:
- Thrust: > 75 mN
- Average Isp: > 1500 s EOL
- Operational duration: < 3000 h
- Cycles: < 6500
- PPU input power: < 1600 W.
Figure 10: Layout of the EPPS (image credit: SGEO consortium)
HEMPT (High Efficiency Multistage Plasma Thruster):
THALES Electronic Systems (TES) of Ulm, Germany is developing and qualifying a novel type of ion propulsion system based on HEMPT technology in the course of DLR’s HEMPTIS project (HEMPT In-Orbit Verification on SmallGEO). The EPTA1 is based on the HEMPT assembly and is intended for being integrated on OHB’s SmallGEO satellite platform to perform attitude and orbit control maneuvers. 27) 28) 29) 30)
The HEMPT assembly consists of four HTM (HEMP Thruster Modules), including 4 XFC (Xenon Flow Controller) and neutralizers, and a single PSCU (Power Supply and Control Unit) developed by Astrium (Germany). The PSCU supplies the HEMPT modules with electric power and controls their operation. Each of the HEMPT modules integrates a HEMPT 3050 ion thruster, an HCN 5000 (Hollow Cathode Neutralizer) and a FCU (Flow Control Unit) which doses the Xenon propellant into the thruster.The FCU is produced by Bradford MOOG. The particular positioning of the four HEMPT modules on SmallGEO allows for all necessary position correction maneuvers in GEO and for momentum wheel off-loading, respectively.
The HEMPT 3050 is a new thruster design and is currently under development at TES. The thruster has successfully passed the PDR and EM units are undergoing endurance tests. The development program of the HEMPT foresees the full qualification before its integration on the SGEO platform and a full dual lifetime test before the launch of the satellite. 31)
The HEMPT performance parameters for SGEO are:
- Nominal Thrust: ≥ 44 mN
- Average Isp: > 2500 s EOL
- Operational duration: < 4800 h
- Cycles: < 6500
- PSCU Input Power: < 1600 W.
Figure 11: Schematic view of the HEMPT Assembly for SmallGEO (image credit: TES)
Legend to Figure 11: The 4 HEMPT modules and the PSCU are shown as 3D drawings, the electrical and propellant harness are indicated schematically as orange/red and blue lines, respectively.
Figure 12: A 3D view of the HTM (image credit: TES)
The HEMPT is based on a particular magnetic confinement of the Xenon propellant discharge which at the same time allows for efficient propellant ionization and ion acceleration. Besides its performance the HEMPT exhibits the unique feature of negligible thruster erosion and therefore shows excellent long-life capabilities. In addition, the HEMPT design concept and operational characteristics enable ion propulsion system architecture with minimum complexity and thus high reliability and cost efficiency.
Flight hardware status: Manufacturing of HTM FM1.4 has been accomplished. The PSCU manufacturing has also been completed (Ref. 30).
CPPS (Chemical Propulsion System):
CPPS is provided by AVIO as system prime and is composed of a well proven EADS S400 – 12 LAE (Liquid Apogee Engine) and 4 main and 4 redundant well proven EADS S 10-18 RCT (Reaction Control Thrusters). The main performance parameters of the LAE and RCTs are:
- LAE thrust: 420 N, Isp: 318 s
- RCT thrust 10N, Isp > 287 s.
The propellant is stored in two 700 liter propellant tanks allocated inside the central CFRP tube. The CPPS carries up to roughly 1300 kg and is capable to provide the nominal delta V of 1500 m/s for GTO to GEO insertion. The GTO sequence depends on the chosen launcher but is expected to be performed with 3-4 burns. At the end of the orbit insertion into GEO and before the station acquisition, the pressurant section is passivated and only the electric propulsion systems and the Xe cold gas thrusters are then used for all orbital maneuvers including end of life disposal.
The SGEO satellite can recover the transfer GTO – GEO phase in case of the LAE failure. In this contingency event the CPPS will utilize all 8 RCT thrusters and the electric propulsion system. Around 90% of the required total impulse will be provided by the RCTs and the remaining 10% will be delivered with the electric propulsion system.
Figure 13: Overview of the CPPS accommodation (image credit: SGEO consortium)
CGTA (Cold Gas Thruster Assembly):
The CGTA, a subsystem of the EPPS, and developed and manufactured by TAS-I (Thales Alenia Space-Italia), consists of the following elements: 32)
- CGCS (Cold Gas Conditioning System)
- 4 main + 4 redundant Thrusters Units (TV). The TV functions are to provide: a) the thrust according to a calibrated relationship vs. inlet pressure with the required accuracy and fast time response, and b) thruster closure of the nozzle throat with very low leakage.
- ADE (Actuator Drive Electronics)
- Piping and harness.
Figure 14: Overview of the CGTA architecture scheme (image credit: SGEO consortium)
The function of the CGTA is to provide reaction control torques to the spacecraft for the detumbling after separation from the launch vehicle and momentum management of sun pointing in Safe mode. For these tasks the total required Xenon budget over the life is in the order of 5 kg. As Xenon is already available on the platform from the EPPS, Xenon cold gas was preferred compared to an additional Nitrogen cold gas system.
The size of the thrusters guarantee a minimum thrust of 50 mN and an Isp higher than 25 s over the life. The thrust is generated by an on-off command of a pressurized piezo thruster (nozzle) valves which are driven by a dedicated redundant electronic (ADE) The ADE provides TM/TC interface, power interface with the S/C, monitor the pressure sensors inside the CGCS, and actuates all the valves.
Figure 15: Layout of the EPPS and CPPS on the SGEO platform (image credit: SGEO consortium)
Table 1: Summary of satellite key features for the HAG1 mission (Ref. 3)
Figure 16: Illustration of the SGEO spacecraft in stowed configuration (image credit: OHB, SSC)
Launch: A launch of the first SGEO platform is scheduled for late 2015 on board an Ariane 5 ECA vehicle from Kourou. The mission will fly a payload for the Spanish satellite operator Hispasat S.A.,known as HAG1 (Hispasat Advanced Generation 1). Hispasat will integrate the HAG1 spacecraft into its existing fleet of geostationary communications satellites (Ref. 2).
Orbit: Geostationary orbit, altitude ~ 35786 km , longitude = 0º.
Payload/experiment complement: (Telecommunication payload, REDSAT)
The first SGEO mission based on the new satellite platform will be launched with a payload for the Spanish satellite operator HISPASAT S. A. as HAG1 (Hispasat Advanced Generation 1). HAG1 will provide multimedia services for Spain, Portugal, the Canary Islands and South America. The telecom operator HISPASAT will be integrating HAG1 in its existing fleet of geostationary communications satellites. The name "Advanced Generation 1" shows that new technologies will be tested with this mission for their use in space. Hispasat AG1, as the first SmallGEO mission, marks the first time in more than 25 years that Germany has assumed system management for telecommunications satellites. 33)
The telecommunication payload for HAG1 is supplied by Tesat Spacecom of Backnang, Germany. 34)
Optical interfaces: Telecommand and telemetry (TM/TC) data of satellite payload equipment is nowadays mostly accomplished by using discrete electrical signals, i.e. high level pulse signals for on/off switching and commanding, and digital or analog signals for telemetry acquisition. This involves a high number of individual electrical lines which have to be routed to each payload equipment leading to a complex and heavy TM/TC harness. By using a single optical fiber instead of several electrical lines, the complexity and mass of the TM/TC harness can be reduced considerably. Moreover, optical data transmission is insensitive to electro-magnetic interference and provides inherent galvanic isolation. These characteristics make optical fibers an ideal candidate for substituting electrical TM/TC interfaces in future satellite payloads.
The major benefits of optical interfaces with respect to electrical interfaces are:
• Galvanic isolation
• No electromagnetic interference (EMI)/ electromagnetic compatibility (EMC)
• Reduced mass
• Ease of integration
• No damage of payload equipment through erroneous harness or connector mating
• Potential for higher transmission data rates (>> 100 Mbit/s)
• Potential for longer cable distances (> 10 m).
Tesat-Spacecom has developed MPM (Microwave Power Module) with an optical interface and an OPLIU (Optical Payload Interface Unit), satisfying the need of controlling new complex payload equipment with the benefits of optical interfaces.
The optical network is comprised of a master node and several slave nodes. The master node controls the transfers on the optical network and forwards telecommands or telemetry requests received from the SMU (Satellite Management Unit) to the addressed slave. The slave node itself sends only telemetry data to the master node, when the telemetry was requested by the master. The master node transfers the gathered telemetry data to the SMU via the Spacecraft interface e.g. MIL-STD-1553 bus. Figure 17 illustrates the network architecture.
Figure 17: Schematic view of the network architecture (image credit: Tesat Spacecom)
Network protocol: The data link layer used for the optical network is based on the CAN protocol, ensuring reliable data transmission between the OPLIU and the payload equipment and cutting the overall costs for equipment integration and test. The CAN protocol provides sufficient transmission data rate (1 Mbit/s) and address space, supports variable packet lengths and is robust. The protocol stack is illustrated in Figure 18.
Figure 18: The network protocol stack (image credit: Tesat Spacecom)
Each data transfer on the network is either be acknowledged by the CAN-Acknowledge or by the 1553 Status Word. If a message transfer error occurs on the optical network, the sending node will retransmit the message again. The maximum retry attempts are limited. Additionally, the master node implements a timeout function for salves which are not responding.
Optical cables and connectors: Commercial fiber assemblies are being used to cut the overall costs for the optical harness.
• Radiation hard fiber assembly - tested up to TID of 80 Mrad
• Optical fiber: Multimode step index, 50/125 µm diameter; Wide temperature range -65ºC to +300ºC
• Space qualified cable: Low mass: 6.4 gram/meter; Wide temperature range: -40ºC to +160ºC
• Space qualified connectors AVIM: Compact, lightweight, 5.9 gram; Excellent performance (typical insertion loss 0.2 dB, return loss type < 45dB); Temperature range: -40ºC to +85ºC.
Figure 19: Optical cable assembly (image credit: Tesat Spacecom)
Optical transceiver: Tesat is currently qualifying an optical transceiver for space application, which originally has been developed for commercial data communication purposes. This transceiver operates at wavelengths of 1310 nm and 1530 nm and relies on a Fabry-Perot laser diode, two PIN photo diodes and micro-optics. It allows full-duplex data transfer on a single fiber. A block diagram is illustrated in Figure 20. The optical transceiver has the following features:
• Operating Temperature: -40ºC to +85ºC
• Integrated WDM filters for dual wavelength Tx/Rx operation at 1310 nm / 1530 nm
• Transmission rate: DC to 5 Mbit/s (FP laser diode transmitter and PIN diode receiver suitable for data rates up to 1.25 Gbit/s)
• Hermetically sealed optical Tx and Rx path.
Figure 20: Block diagram of the optical transceiver (image credit: Tesat Spacecom)
Two different types of transceivers are used to establish a point-to-point connection (Figure 21). The master node uses transceiver which emits light at a wavelength of 1530 nm and receives at a wavelength of 1310 nm. In slave nodes transceiver emit light at 1310 nm and receive at 1530 nm.
The major advantage of this full-duplex operation mode is that only one fiber is needed for a point-to-point connection with the opportunity of bidirectional data transfer. So the weight of the optical harness will be reduced compared to a optical connection which needs two fibers for bidirectional data transmission.
The optical transceivers will be set to an optical output power of about 1.8 dBm ex fiber over the temperature range of -40ºC to +85ºC. The power control is realized by a passive stabilization network adjusting the laser diode current regarding the operation temperature of the transceiver.
Figure 21: Bi-directional optical link (image credit: Tesat Spacecom)
OPLIU architecture: Since the currently available SMUs (Satellite Management Units) and PLIUs (Payload Interface Units) do not provide optical interfaces, Tesat Spacecom has developed a so called OPLIU (Optical Payload Interface Unit). The OPLIU has a modular architecture and can control up to 40 MPMs. The MPMs are connected in a star topology to the OPLIU.
Figure 22: Overview of the OPLIU equipment (image credit: Tesat Spacecom)
The OPLIU is comprised of a nominal and redundant control unit, a nominal and redundant DC/DC converter (power supply unit), a backplane and up to five independent OIF (Optical Interface) boards. The mechanical outline of the OPLIU is shown in Figure .
The backplane is mounted in plane with the mounting plate. All other units are mounted perpendicular to the mounting plate. All external electrical and optical interfaces are accessible on the top side of the OPLIU assembly. A redundant internal serial bus is used for interconnection of the OIF-boards and the control unit. The internal bus is based on the CAN protocol with transceivers according to the RS-485 standard.
OPLIU accommodation parameters:
• 50 V main bus interface
• Mass of ~ 6 kg
• Power consumption 10 W (typically)
• Dimensions L x W x H = 190 mm x 280 mm x 190 mm
• High reliability.
• Discrete TM/TC (DC/DC converter on/off commands, primary current, bi-level status, temperature)
• 40 optical interfaces
• 80 high level pulse commands to switch the payload equipment on or off, respectively.
• 40 bi-level status signals, which represent the power status (on/off) of the connected payload equipment.
DC/DC converter: The OPLIU internal DC/DC converter is dual redundant. Each converter is split into two parts. One part powers the control unit, whereas the other part powers the optical interface boards. Each converter is overvoltage and overcurrent protected.
Control unit: The main function of the control unit is to decode and execute the telecommands received via the MIL-STD-1553B bus from the spacecraft. The command decoding and execution is implemented in hardware with a radiation tolerant FPGA. Four different command types/transfers are supported:
• On/off command: Used to switch payload equipment or OPLIU internal power supplies on/off.
• Register load command: Provides write access to the OPLIU internal or to the payload equipment internal memory.
• 8 bit serial digital data acquisition: Provides 8 bit telemetry data from the OPLIU internal memory or payload equipment memory.
• 128 bit serial digital data acquisition: Provides 128 bit telemetry data from the OPLIU internal memory or payload equipment memory.
In addition, the control unit supports functions to supervise and control the OPLIU internal equipment e.g. the supervision of the DC/DC converter voltages.
Figure 23: Photo of the control unit breadboard model (image credit: Tesat Spacecom)
OIF (Optical Interface) boards: Each OIF board is designed to control up to eight Tesat MPMs (Microwave Power Modules). All electrical parts except the optical transceivers are redundant on the OIF board. The power for the nominal and the redundant parts of the OIF board can be switched on by the control unit. The routing functions between the internal bus and the optical interfaces are implemented in a radiation tolerant FPGA. The power consumption of one OIF board is about 1 W.
OIF board interfaces:
• 8 optical interfaces
• 16 high level pulse commands
• 8 bi-level status signals.
Figure 24: Photo of the OIF board breadboard model (image credit: Tesat Spacecom)
REDSAT (Regenerative Processor System):
The REDSAT payload is based on the initiative started in the framework of the Spanish Space National Program. This payload includes a receive active antenna and all the subsidiary RF units needed to provide the service in the areas defined as coverages across the land masses of the orbital slot. The system provides 4 reconfigurable uplink beams in Ku-band and up to 4 regenerative channels of 36 MHz to provide broadband services in Ku-band or Ka-band.
REDSAT is an advanced communication OBP (On-Board Processor) developed by TAS-E (Thales Alenia Space España) which also includes EADS CASA Espacio. The REDSAT payload includes a regenerative on-board processor based on the DVB S2/DVB RCS standard, and an advanced Ku- band active antenna. 35) 36)
Note: DVB-RCS (Digital Video Broadcasting - Return Channel via Satellite). It is a specification for an interactive on-demand multimedia satellite communication system initially formulated in 1999 by the DVB consortium.
The REDSAT payload features two major innovations: a broadband multimedia system and an advanced antenna system.
• The broadband multimedia DVB system is an end-to-end advanced communication system based on an OBP (OnBoard Regenerative Processor) and providing DVB-RCS/DVB-S2 capabilities including an IP router in the space. Thales Alenia Space España will define, develop and produce the whole system.
• The advanced antenna system will allow the AG1 mission to meet its performance specifications for both the transparent and regenerative repeaters in the Ku-band frequency. EADS CASA Espacio will develop and produce the reflector and the DRA-ELSA (Direct Radiating Array-Electronically Steerable Antenna).
The Redsat OBP also features two Thales L-band processor units, one Mier Communicaciones Ku-/L- band down and up converter unit and a box of Thales filters with its flight harness. It is able to simultaneously process four 36 MHz communication channels. The OBP is in charge of demultiplexing, demodulation and decoding the received DVB-RCS uplink signals, and re-multiplexing the received data in order to build four DVB-S2 compliant downlink TDM (Transport Data Mechanisms). 37)
Figure 25: Redsat OBP flight units : LBP-Tx, LBP-Rx, DUC, filters (from left to right), image credit: ESA
DRA-ELSA (Direct Radiating Array-Electronically Steerable Antenna)
DRA-ELSA is an active array specifically designed for new satellite telecommunication missions and future needs of the broadcasting market (DVB-S2). The dedicated receive antenna, to be operated in the Hispasat AG1 mission, will mainly allow on-demand high flexibility and electrical steerable capabilities as differential features when compared to standard products and services currently offered. 38) 39) 40)
Table 2: Summary of DRA-ELSA requirements
The DRA-ELSA allows the reception of Ku-band carriers over four simultaneous spot beams working in horizontal polarization. The antenna is able to synthesize each one of these spot beams in any direction over the Earth from the defined orbital position with the required sidelobes level by means of controlling the excitations (amplitude and phase) of the array radiating elements. Primarily designed to operate with steerable pencil or spot beams, each beam can be independently pointed at any Earth location readily whenever requested by command.
Additionally, DRA-ELSA can support beam hopping capabilities, that is, time sharing service based on switching automatically the beams between a set of pre-defined locations and configurations. This feature would introduce adaptability to the presented array antenna.
The driving parameter of the DRA-ELSA RF architecture has been to provide simultaneously:
• The required G/T over the four simultaneous spot beams withstanding the antenna temperature and the DRA-ELSA noise temperature and directivity for each beam, as well as the S/C post-DRA noise figure.
• The output power within the required range considering the PFD of the in-band signals, the antenna directivity and the gain and losses of the architecture needed to fulfil the G/T requirement.
• The required linearity of the architecture.
DRA-ELSA architecture: The antenna provides 4 independent and fully reconfigurable beams, working in the uplink Ku sub-band from 14.25 to 14.50 GHz in linear polarization.
The array, which follows a modular approach, comprises 100 radiating elements (subarrays) arranged in a square configuration with a total aperture of 650 mm x6 50mm. Figure 26 shows a 3D model of the antenna. The radiating square array is under the shaded green area (thermal blanket isolation). The total dimension of the antenna is 900 mm x 700 mm and 400 mm of height with a total mass close to 60 kg. The antenna is firmly supported by four legs that are attached to the top floor of the spacecraft.
Legend to Table 3: The values designated with (*) can be optimized for any mission and payload interface specification. Values are an example of mission scenario (not necessarily those of Hispasat AG1).
The antenna is composed by the next main functional subassemblies:
• Subarray subassembly: the subarrays are grouped in 2 x 2 radiating elements to provide a higher level of integration form the manufacturing point of view.
• Filter subassembly: Below each group of 4 subarrays, there is a group of 4 filters to provide the necessary out-of-band rejection.
• MCCM subassembly: Next to the filters there is a hybrid circuit which amplifies-filters-amplifies the signal before splitting to process the signals coming from the 4 subarrays block with the proper amplitude and phase for each one of the spot beams to be synthesized. Besides the control of the illumination coefficients, this MCCM (Multi-Chip Control Module) subassembly provides the first level of signal combinations and a first part of the needed gain of the active chain to fulfil both the G/T with the post-antenna noise figure and the output power requirements for each beam.
• Combiners subassembly: As said above, the combination is done in two levels. Instead of the combination of 100:1 in one level, the first level of combination 4:1 is implemented into the MCCM subassembly, and the second level of combination 25:1 is implemented by means of 6 combiners of 5:1 per each one of the spots beams.
• MCCM Signal Conditioner subassembly: Next to the combiners network there are 2 modified MCCM modules, called MCCM Signal Conditioner, which provide the last gain stage required. Each one of the modules handles the signals of 2 spot beams. This subassembly includes a variable attenuator after the amplification of the signal in order to allow properly programming and adjustment to optimum of the power depending on the different operational scenarios.
• Test coupler subassembly: A coupler at the output of each beam signal path is included to test during integration.
The signals coming from each subarray are filtered out and then processed in the MCCMs (Multi-Chip Control Modules), which constitute the core of the antenna. The modules include amplification, variable attenuation and phase shifter to individually control each beam. Combination is done by means of four specific passive combiners. In order to optimize the performance versus the payload I/F, a signal conditioner module has been included at the output of each port allowing output power level adjustment in a range.
The antenna is capable of synthesizing steerable spot beams with a given beamwidth in any direction over the Earth optimizing the following electrical parameters:
• G/T is optimized for every beam pointing, withstanding different beam directivities and minimizing equipment noise temperature. Moreover the antenna can be configured to reduce and even make negligible the noise temperature contribution of the payload it is integrated in.
• Output power can be set within a specified range. It is adjusted for every beam pointing to achieve levels and a dynamic range compatible with the payload sub-units the antenna liases with. Furthermore the output power values and dynamic range can be configured to different payload interface requirements and also for different power levels of uplink signals.
• Linearity, defined as C/I (Carrier to Interference) value. The antenna amplifies the signals received from Earth through synthesized beams with great linearity purity. The intermodulation noise is reduced to very low levels even in the worst operation scenarios. The antenna is capable of operating when up to 12 different channels in the Ku uplink band operate simultaneously at Hispasat AG1 mission maximum power levels and from the same Earth location. C/I levels higher than 50 dBc are guaranteed even in the extreme case.
• Pattern contour. Beamwidth and Side Lobe Level (SLL) for the case of spot beams. The patterns are optimized to maximize directivity values inside the required beamwidth keeping the cross-polar level and SLL below a specified value. In order to achieve that, a suitable power pattern synthesis algorithm is applied. Levels of cross-polar discrimination higher than 28 dB and SLL below 25 dB are achieved for a beamwidth diameter of 1.5º. The DRA performance for shaped beams depends on the contour areas required.
In addition to the RF functionality, the MCCM is in charge of the electronic control of the active antenna. MCCMs are comprised of digital/analogue ASIC electronics that are monitored and telecommanded by the antenna central control unit (ICU) that interfaces with the satellite. A power unit (PSU) is the responsible to provide the secondary voltages demanded by the DRA sub-units from the primary S/C supply.
Beam pointing or shaping is achieved by means of controlling the excitations (amplitude and phase) of the array radiating elements. These excitations are either internally stored or uploaded from ground into the antenna. The control system of the antenna can be commanded from ground station to update the excitation configurations and so, change the operating scenario in any moment. Main control capabilities of the antenna are the following:
• Providing the TM (Telemetry) necessary to know the status of the antenna equipment (temperatures, consumptions, etc) to be able to react in case of any failure.
• Processing the TC (Telecommands) from the on-ground station to load any of the array configurations stored in the data base of the antenna or to store and load any other new configuration if needed. In-orbit control makes the antenna capable of being well suited to a possible change of the mission (orbital position, repointing of the spot beams, definition of new coverages of interest, etc.).
• Controlling the switching on/off of the active subassemblies or nominal/redundant units of the antenna in order to preserve its correct operation. It provides the antenna with some capacity of reaction in the event of failures (subarray, amplifier, etc) readapting the illumination coefficients of the rest of the elements.
Thanks to a dedicated thermal design control based on MLI (Multi-Layered Insulation), heat pipes, heaters and loop heat pipes, the system equalizes and regulates the operating temperature points and excursions of the electronics for an optimized RF performance. This aspect is deemed critical to guarantee a low and stable noise temperature that will contribute to the mission G/T performance parameter.
RF antenna architecture: Incoming signals from each subarray are filtered prior to being processed in the MCCMs (Multi-Chip Control Module) that includes LNAs (Low Noise Amplifiers), 1:4 divider, and variable attenuation and phase shifter to individually control each beam. After the processing, all signals are collected by means of a BFN (Beam Forming Network) comprised of two levels of 5:1 combiners to enhance the modularity of the antenna. Finally two modules, called MCCM-SC (Signal Conditioners), are placed at the output of the antenna. With these modules the antenna can adjust their electrical interface (mainly output power and noise figure) with the satellite payload (Ref. 39).
Subarray: The antenna aperture is made up of 100 subarrays arranged in a square lattice of 10 x 10. The subarray design is based on the recognized heritage of EADS CASA Espacio with printed patch antennas and arrays. Each subarray is made up of 4x4 circular disk patches fed up in linear polarization by the SSSL (Suspended Substrate StripLine) technology). This printed technology yields a very light and robust mechanical piece maintaining also very efficient electrical performance.
Figure 27: DRA-ELSA subarray model (image credit: EADS CASA Espacio)
Figure 27a shows a subarray unit. The model has been mechanically built in blocks of 2 x 2 subarrays, for simplicity during integration. The patch has been designed to optimize the cross-polarization levels, achieving good efficiency and bandwidth performance, but mostly having a robust design in the Ku-band, able to withstand any manufacturing dispersions.
Figure 28 shows a normalized radiation pattern of the subarray at 14.50 GHz. The measurement has been performed with the configuration shown in Figure 27b surrounded by other subarrays to account for coupling effects that affect the final radiation pattern.
Cross Polar (XPD) inside the coverage is better than 32 dB, which is an excellent figure for the state-of-the-art and a very good reference for the antenna synthesis and final beam isolation. It can be observed that, near to boresight, XPD values exceed 40 dB. Measured subarray losses within the band are lower than 1.15 dB, including cables and connectors. The radiation efficiency is very important due to its direct influence in the antenna G/T performance.
All measured units showed an excellent agreement with predictions. The maturity of the design and manufacturing processes allow for an efficient development flow, first-shot and repetitive models.
Figure 28: Measured subarray radiation pattern (image credit: EADS CASA Espacio)
Filter: A waveguide filter is placed right after the subarray in order to filter interference signals out of the antenna receiving band. Figure 29 shows a filter model. The filter is sized in such a way that interfering signals that are present in any telecommunication mission can be highly rejected. The filters have shown low insertion loss and rejection levels higher than 50 dB in Ku-band and 70 dB in Ka- and X-bands.
Figure 29: DRA-ELSA input waveguide filter (image credit: EADS CASA Espacio)
MCCM (Multi-chip Control Module): Based on EADS-CASA Espacio's heritage, a multi-chip hybrid module was developed for the DRA-ELSA. This module consists of a multi-layer ceramic package that houses thin film passive circuits, MMIC low-noise amplifiers (LNAs), phase shifter and attenuator chips, ASICs and other discrete components for control and telemetry.
The main scope of this unit is to control the illumination coefficients (attenuation or phase) of the RF signal received from each subarray for the beam synthesis. Moreover, the signal must be amplified for a correct reception and noise figure balance. - Several MMIC dies were developed for the module: low noise amplifiers, digital phase shifter and digital attenuators. Figure 30 shows the discrete set of phase and attenuation states (the so-called “constellation”) that the MCCM can implement for each subarray to synthesize any required radiation pattern.
The control electronics of the MCCM is implemented with an integrated architecture based mix functionality on a specifically developed analog ASIC plus digital ASICs. The ASICs and their associated circuitry are in charge of commanding the Phase shifters and the Digital Attenuators, biasing the amplifiers and sensing current and/or temperature inside the hybrid module among other functionalities. The whole package is hermetically sealed with a lid, closed into an inert nitrogen environment.
Figure 30: Phase/Attenuation state constellation (image credit: EADS CASA Espacio)
BFN (Beam Forming Network): The signals from all the MCCMs are collected into 4 RF outputs (one per beam) by means of two modular levels of combiners (5:1): a primary level of 20 5:1 combiners and then a secondary level of 4 5:1 combiners. Each combiner is of a Wilkinson type, which provides good input matching and a substantial level of isolation between the input signals.
The design is based on a 1:8 Wilkinson splitter, built with 1:2 splitters in a low-loss stripline technology. In this solution, 3 outputs are loaded. This solution has a related impact on insertion losses, absorbed by the DRA-ELSA configuration.
MCCM-SC (MCCM Signal Conditioner): MCCM-SC is implemented over a slightly modified version of the above described MCCM package. For each spot beam, there are four amplification MMIC stages and four digital attenuators plus the ASIC to command them. The signal conditioning makes the antenna output power suitable for a wide range of payload requirements:
- Adjustment of output power to the payload requirements in order to optimize the noise figure of the antenna plus payload, and therefore G/T performance.
- Protection and redundancy against failures.
- Robustness against component dispersion and degradation due to aging.
DRA RF performance:
Pencil/spot beams: Figure 31 shows the G/T patterns for two spot beams pointing at Madrid and Rio de Janeiro at 14.25 GHz. The G/T performance offered by DRA-ELSA spot beams are significantly better than values typically offered by reflector antennas.
This is basically a consequence of the capability of the active antenna to adapt to the traffic demands in a specific area, optimizing the system utilization during the mission. It is important to note that one antenna beam could provide service for several areas using beam hopping services.
For this two pointing examples, side lobe levels (SLL) are lower than -25 dB, while cross-polar isolation (XPI) is around 32 dB (Madrid) and 42 dB (Rio).
Figure 31: G/T over Madrid at 14.25 GHz (left) and G/T over Rio de Janeiro at 14.25 GHz (right), image credit: EADS CASA Espacio
Figure 32 shows the G/T values for a spot beam as a function of its scanning angle off boresight. The array has a small penalization in directivity when it is pointed, due to grating lobe effects at the edge of the Earth.
Figure 32: G/T variation with scanning losses (image credit: EADS CASA Espacio)
Shaped beams: The DRA-ELSA antenna provides the capability and flexibility to reshape beams over any location on the visible Earth. An assessment of the G/T performance of the DRA-ELSA is presented for several scenarios. Figure 33 shows an example of a shaped pattern over small areas in America.
Figure 33: Localized shaped beams over America at 14.25 GHz (image credit: EADS CASA Espacio)
An extended and larger coverage over the American continent is presented in Figure 34. The shaped area is extended to all higher coverage. In this case, the obtained G/T performance is comparable to G/T values provided by a standard large reflector antenna.
An added advantage is that this coverage can be reshaped to reinforce G/T performance over coverage areas with more demand or where a market had not been foreseen, a sports event takes place or where a new potential customer arises in an atypical coverage.
Figure 34: Wide America pattern at 14.25 GHz (image credit: EADS CASA Espacio)
DRA-ELSA (Direct Radiating Array-Electronically Steerable Antenna) will be operated by HISPASAT in the Hispasat AG1 satellite mission. The antenna is meant to be the first reference of a very promising, entirely new generation of flexible and potentially adaptable broadcasting services for satellite telecommunication. It has been conceived to be an innovative and commercially attractive product capable of providing on-demand and flexible service definition Ref. 39).
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.