The ARTEMIS satellite (Giubilei 1992; Dinwiddy 1992; Benedicto 1990; Pe rrotta 1990; Lisi 1991; Morando 1991; Dickinson 1990; Laurent 1991) will validate land mobile satellite services (LMSS), inter-orbit data relay and ion propulsion system technology. This multi-payload experimental geostationary orbit satellite is current ly under development for ESA by a team of European sub-contractors and launch is planned for 1995. A top view of the spacecraft is shown in Figure 5.2. This mission has undergone recent changes which were approved by the European C ouncil of Ministers in the fall of 1991. The initial role of ARTEMIS as an advanced technology demonstration mission has been de-emphasized with greater emphasis being given to operational functions. The present plan is for ARTEMIS to be located initial ly between 6 degrees and 20 degrees E from 1996 until 1998 where it will be used to carry out land mobile communications and inter orbit link experiments. Then, in 1998 when the first DRS is launched, it will be moved to 59 degrees E to provide, with DRS , an operational data relay service for the Polar Platform Mission and later missions. During this phase, the LMSS payload will no longer be used. ARTEMIS entered the C/D implementation phase in late 1992.
The major ARTEMIS technical missions are:
The feeder link for the inter-orbit link payloads (SKDR and SILEX) will be at Ka-band (20/30 GHz) and the feeder link for the land mobile payload will be at Ku-band (12/14 GHz).
LLM Characteristics. The LLM payload (see also Chapter 3) will provide for mobile communications between fixed earth stations and terrestrial mobile vehicles. To provide two-way communications, the paylo ad has two transponders, one for the forward and one for the return link. A block diagram of the payload is shown in Figure 5.3. The LLM is required to provide in-orbit spare capacity for the L-band European Mobile Communications S ystem, a payload to be flown aboard the ITALSAT F2 spacecraft (EMS/F2). The LLM 3 m reflector antenna will provide L-band coverage generating one dual-polarized "Euroglobal" beam and two single- polarized and one dual-polarized spot beams cover ing Europe. Over the last several years the design requirements for the LLM payload have been relaxed, particularly for the antenna design
Figure 5.2. ARTEMIS Top Floor View
Figure 5.3. LLM Payload Block Diagram
In 1990 a very aggressive design plan was reported which, in addition to the Euroglobal beam, required six fixed spot beams and a steerabl e beam from a 5 m reflector (Benedicto 1990; Perrotta 1990). Several approaches for implementing the antenna had been studied including an inflatable design. The mass of the current design is less than 100 kg compared with 150 kg for the 1990 design. T he characteristics of the current LLM payload design are given in Table 5.1
Characteristics of LLM Payload Design
The 1.5 GHz forward direction provides seven teen user links channelized as follows:
I t will be possible to operate up to six forward 1.5 GHz user-link channels simultaneously.
The 1.6 GHz return transponder user links are channelized as follows:
Fi gure 5.4. LLM Eurobeam Coverage
It will be possible to operate up to 15 return 1.6 GHz user link channels simultaneously.
For the forward links, the total EIRP is specified as 44 dBW for the Euroglobal beam and 47 dBW to be shared across the thr ee spot beams. For the Euroglobal beam, the ratio of gain to noise temperature (G/T) at the receiver at 1.6 GHz will be greater than -2 dB/K, while for the spot beams, the G/T will be greater than 0 dB/K.
The L-band EIRP requirements have been sized with the intention of providing a good balance between RF power and spectrum availability. The communications capacity of the payload is more than 400 voice channels of 4.8 kbits/sec data rate (rate 3/4 Viterbi encoding), quadrature phase shift keyed (QP SK) modulated with a voice channel spacing of 10 kHz.
Figure 5.5. LLM Spot Beam
The LLM payload exhibits a high degree of in-orbit reconfigurability to ease the operation of the spacecraft, and to optimize, at any time, the use of the on-board resources to maximize system capacity. Reconfiguration of the payload can be implemented at different levels:
The payload exhibits the EIRP and G/T performance shown in Table 5.2.
LLM EIRP and G/T Performance
T he available RF power resources can be shared in any ratio among the Eurobeam and any of the fixed spot beams at L-band, in order to match the nonuniform distribution of traffic across the coverage area.
In termediate Frequency Processor (IFP). One of the most important features of the IFP is its high efficiency in handling the L-band spectrum. The use of offset frequency conversions for adjacent frequency slots eliminates the loss of spectrum due to f ilter guard bands, at the same time easing the out-of-band rejection requirements for the surface acoustic wave (SAW) filters and reducing the coherent multipath component through the repeater. The IFP contains a bank of SAW filters operating in the rang e 140 to 169 MHz. The filters provide more than 37 dB of out-of-band rejection at 250 kHz away from the pass-band. The filter bandwidths are 1 MHz and 4 MHz. All filters are manufactured on quartz substrate to achieve good temperature stability. Each of the filter outputs, upconverted to L-band, can be routed to any of the 11 input ports of the beam forming network (BFN) by the use of miniaturized microwave monolithic integrated circuit multithrow switches which provide over 40 dB isolation.
In or der to maximize the efficiency in handling the L-band spectrum, a novel technique of filter guard-band reduction has been implemented. Each 1 MHz slot at intermediate frequency is up-converted by a different local oscillator signal. Adjacent local oscil lator signals are offset by an amount equal to the filter transition band (250 MHz) and in this way 100 per cent spectrum efficiency is obtained at L-band. The frequency overlapping at L-band allows reuse of the same frequencies in three beams. Moreover , each of the 1 MHz frequency slots is controlled in frequency band by a different frequency synthesizer so that up to six slots can be overlapped in frequency to provide six-fold frequency reuse, i.e., the use of the same band in all spot beams, for a mu ltibeam code division multiple access (CDMA) system. The implementation of such an IFP is largely dependent upon the use of miniaturized frequency synthesizers. Seven such synthesizers are needed for the LLM IFP, the outputs of which are reused by the f orward and return transponders. Each synthesizer has 250 kHz resolution and 25 MHz bandwidth in the L-band frequency range (1,500 MHz), with outstanding phase noise and spurious performance, the mass of each unit is 150 grams, and the power consumption 2 .5 W.
L-Band Transmit and Antenna Subsystems. In order to meet the EIRP requirements, the Eurobeam edge-of-coverage directivity is required to be around 27.5 dBi and the spot beam directivity around 28.5 dBi. The L-band spot beams provide 15 to 20 dB of isolation between the west (A) and east (C) beams. This feature will be used to perform experiments of frequency reuse by spatial discrimination in frequency division multiple access (FDMA) and CDMA systems. In addition, RHCP is provided in the central spot beam (B) for experimental purposes.
A novel semi-active antenna configuration has been selected for the LLM payload, the "Multimatrix-fed reflector" (ESA patent pending). In the baseline LLM configuration a 2.9 x 3.3 m proj ected aperture reflector is used with a feed array of six elements at its focus, fed by a multimatrix transmit section. All solid-state amplifiers operate at the same input drive level, independent of the actual beam-to-beam traffic distribution. This f eature maximizes the DC to RF conversion efficiency. Power division and phasing for each beam at the amplifier inputs is provided by a low level BFN. A fixed BFN is used for the fixed beams.
SKDR P ayload Characteristics. The SKDR payload will provide one bi-directional inter-orbit link (IOL) at S- or Ka-band, a feeder link section for the optical terminal (SILEX), plus service channels for tracking, telemetry and command (TT&C) signals, pi lot tone and transmission of data from some experiments on the satellite. The ARTEMIS data relay service will provide the accesses and channels listed in Tables 5.3 and 5.4.
ARTEMIS Link Accesses
ARTEMIS Return Link
It should be noted that the last row of Table 5.4 (Mode 4) refers to an opera ting mode envisaged only when ARTEMIS is positioned in a DRS position (59 degrees E). The mass and power of the payload are given in Table 5.5.
The SKDR payload is equipped with a European feeder link antenna (EFLA) and an inter -orbit link antenna supporting both S- and Ka-band links. A block diagram of the SKDR payload is shown in Figure 5.6. The EFLA provides two beams, one for each specific coverage over Europe. The antenna simultaneously performs the transmission and reception of a single linear polarization. The antenna configuration consists of a shaped offset reflector (about 1 m diameter) fed by two switchable feeds, each optimized for one of the two orbital positions. The payload can receive u p to three uplink signals in the 30 GHz band:
Ta ble 5.5
ARTEMIS Mass and Power
After low noise amplification (low noise amplifier (LNA) noise figure (NF) 4.5 dB), the signal is divided by four and part of the signal goes to the 20/30 GHz TT&C subsystem for p rocessing. The communication signal is converted to a common IF (5.5 GHz) by means of tunable frequency converters, which are driven by a reference frequency (100 MHz from the frequency generation unit) and are able to synthesize a local oscillator (LO) frequency that down converts any frequency in the 5.5 GHz band.
The IF signal can be routed to three different sections:
The feeder link section of the return repeater consists of four identical chains. Each chain consists of a channel amplifier, a tunable frequency converter, and a TWTA working at saturation (30 W). No ring redundancy at the output is planned; each converter works on a fixed output frequency.
Figure 5.6. ARTEMIS SKDR Payload Block Diagram
The inter orbit link (IOL) antenna is the most critical ARTEMIS/DRSS antenna. It is a dual-band antenna (S- and Ka-band) providing bi-directional RF links between the satellite and a user spacecraft terminal. The design is based on an offset parabolic reflector, with a projected aperture diameter of 2.85 m, f/D=0.5 and a fixed feed assembly placed in the focus. The reflector rotates around the focus using a yoke of suitable design, implementing a new concept which offers several advantages: fixed feeding networks, lo w scanning losses and relatively small perturbation to the spacecraft attitude. The antenna can be steered over a half-cone angle of about 10°. The IOL antenna includes a pointing subsystem: at S-band the user tracking is performed open-loop, while at K a-band the user's acquisition and tracking is performed using an RF sensor integrated in the Ka-band feed.
A single channel system has been selected, in which a single receiving chain is used for both communication and tracking purposes. The key idea of the single channel system is the use of the same feed for both the communication and the tracking function, implementing a "conical scan" so that a single channel carries both communication and RF sensing information. A possible implementat ion of the feed consists of a cluster of four square horns, operating in circular polarization. In order to set the squint angle, each beam is obtained by synthesis, combining the four original beams with a set of weights. The weighting is performed by a BFN based on variable power dividers realized with ferrite technology. An alternative solution, at present in breadboarding phase, is to use a multimode feed with a monopulse converter device.
The communication signal, with the superimposed pointin g information, is derived at IF by selecting one of the possible Ka-band IOL return channels. The tracking channel is sampled and the error information is extracted using a tracking receiver which measures the power coming from each sensor. An antenna p ointing controller combines the measured powers, generates the error information and drives the control electronics (APMDE) of the actuators (APM), which move the reflector.
The SILEX optical communications package will be test ed with LEO satellites including SPOT and EURECA (Laurent 1991) and is described in detail in the next section.
ARTEMIS will fly advanced platform technologies in addition to the communications p ayloads. These technologies include the ion propulsion package (Silvi 1991; Protto 1991) which will apply ion propulsion for 10 years providing North/South station-keeping to demonstrate its operational capability for the follow-on DRSs and other future missions.
ARTEMIS Spacecraft Bus and Mission. A single ARTEMIS satellite is currently being designed by a team of contractors from several European nations led by Alenia Spazio. In addition to the L-band payload for the promot ion of the technology for future generations of mobile systems, ARTEMIS is intended to demonstrate new technologies and services to be included aboard three DRSs planned for launch in 1996 and in 2000. No detailed information on the satellite design, mas s properties or power requirements is available. However, the satellite mass into geosynchronous transfer orbit (GTO) will be in excess of 2,250 kg. It will have an operational lifetime of ten years.
As mentioned above, an Ariane 5/01 is scheduled t o place ARTEMIS in a GEO orbit at a position between 6 degrees E and 20 degrees E in 1996. From this initial position, the satellite will be capable of moving to 59 degrees E. It will maintain its station within a box centered at its initial position th at is 0.14 degrees in the E-W direction by 0.14 degrees in the N-S direction.
ARTEMIS is a precursor program of ESA's DRS. Seventeen aerospace companies from ten European countries are involved in DRS. The ARTEMIS satellite is expected to cost about 400 million "then year" dollars. Alenia Spazio was selected as the prime contractor in March 1989, program approval from ESA was obtained in July 1990, and development was scheduled to begin by the end of 1991. ARTEMIS will be deployed in 199 5.