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Capabilities of lightsat constellation for operational altimetry
Pergamoo
00!&5765@4)0027~2
AcraAstronourica Vol. 3s. No. 1, pp.429-435, 1995 Copyright Q 1995ElscvicrScienceLtd printedin Great Britain. All rights reserved 0094-5765195$9.50+ 0.00
CAPABILITIES OF LIGHTSAT CONSTELLATION FOR OPERATIONAL ALTIMETRY? J. P. AGUTTES
and G. BELLAICHE
Centre National d’Etudes Spatiales, 18 av. Edouard Belin, 31055 Toulouse Cedex, France (Received 16 March 1994; received for publication I November 1994) Abstract-A radar altimeter mission may involve more than one satellite in space, especially when the objective is to collect data on meso-scale features. However, dedicated satellites can be very small thanks to the integrated technology of modem altimeters such as the French POSEIDON instrument. For these reasons, light satellites appear to be an affordable concept for the deployment of this type of constellation and, moreover, give more flexibility than “piggy-backing” altimeter packages on several distinct, heavier, multimission platforms. This paper gives an overview of typical system and mission trade-offs resulting from operational, technological and economic constraints in different domains: payload and satellite configuration, orbital constellation adjustment, ground tracking facilities etc.
I. INTRODUCTION I.I.
The TOPEX/POSEIDON
legacy
The forthcoming launch of the TOPEX/POSEIDON satellite (a French-U.S. co-operative venture) will make concrete reality of nearly 10 years of effort by the CNES to develop and master altimetry techniques and data utilization. The DORIS mission on-board SPOT 2 was an essential first step, providing full-scale validation of the precise orbitographic techniques needed to back up the altimeter measurements. Thus, right from its first flight, the TOPEX/POSEIDON mission will be able to make good use of the measurements from the POSEIDON altimeter, a new generation instrument using highly integrated technology. TOPEX/POSEIDON should confirm the soundness of CNES’s original approach, combining pragmatism of concept (a single frequency of the altimeter, one-way measurements for DORIS) with audacity of technology (solid-state altimeter, new generation of ultra-stable oscillators for DORIS) and leading to a minimum payload of small size that should encourage the setting up of more altimetry missions, either as passengers on host platforms or in small, dedicated satellites. 1.2. New opportunities for the future The topographic sea signal has spectrum, depending on the time and measurement, and altimetry can serve of applications. TOPEX/POSEIDON
a very large space scale of different types is a scientific
__tPaper IAF-92-87 presented at the 43rd Astronautical Congress, Washington D.C., U.S.A., 28 August5 September 1992.
mission optimized for the study of large scale features (more than 5000 km) and their seasonal and multiyear variations. But the scientific community also has interests which involve a smaller space scale observation, like the study of polar ice or the determination of a precise sea geoid. Up to now meso-scale features were mainly related to defence concerns but due to the involved energies and interactions with other scales, such phenomena are also attracting more and more scientists’ interests. This paper mainly draws its inspiration from studies carried out on behalf of the French DOD for military missions. The meso-scale aspect of these missions sets the first conditions of the system as it imposes the use of several detectors in-orbit at the same time. But such systems, as shown here, can continue to serve the scientific community with the addition of some marginal constraints. This synergy and the flexibility of implementation with light satellites should bring new opportunities for the future of scientific altimetry. 1.3. Light satellites The light satellite is a concept causing much excitement in the space world today. Techniques have reached a point that enables important payloads to be made in a small volume, but the real cause of all the enthusiasm is the emergence of small, low-cost launchers. The association of small satellites with small launchers means that even modest missions can take control of their access to space, i.e. above all. the choice of orbit, the date of the first launches and the strategy for replacing satellites in orbit at the end of their lifetime. All this is obviously more difficult when a launcher, or worse, a large space platform, has to be shared with several other missions. 429
430
J. P. Aguttes and G. Bellaiche
The meso-scale altimetry mission is, in this respect, a typical case showing the interest of the combination because of the number of satellites brought into play and the requirements as far as choice of orbit and system availability are concerned.
the economic trade-off decided upon for the mission. An optimal strategy can be defined (see Section 5) fixing the lifetime of the satellites, the modes of replacement in orbit or on the ground and the time between launches. 2.4. System coverage (or service area)
2. TYPICAL MISSION REQUIREMENTS FOR A LIGHT SATELLlTE SYSTEM
2.1. Spatio -temporal sampIing Meso-scale phenomena are characterized by distance scales of 60-300 km and time scales of lo-30 days. With three satellites in orbit, it is possible to take samples on the ascending satellite tracks over 10 days/85 km compatible with the observation of the phenomena, given the oversampling contributed by the descending tracks. 2.2. Orbital cycle and track maintenance The conventional mission approach for obtaining the meso-scale signal consists of using repetitive orbits, keeping on track to within about a kilometre. In a first learning phase, the accumulation of passes allows the mean ocean profile to be reconstructed. This is considered to be stable and is subtracted at each useful pass so that the meso-scale signal can be isolated. There are two ways of obtaining the desired sampling, each considering three satellites following each other at regular intervals along the same nominal orbit: -ach satellite performs the required temporal sampling through the choice of an 11-day orbital cycle. The difference of 120 between the satellites multiplies (because 11 is not divisible by 3) the tracks and ensures spatial sampling; -the opposite: each satellite covers the required spatial mesh with a time cycle of 30 days. The 120 phase difference along the orbit means that the three satellites follow each other over the same track (because 30 is divisible by 3) at lo-day intervals. The 30-day solution is preferable since, if one satellite fails, the other two can be reconfigured at 180. This allows temporal degradation to be kept regular (15 days instead of an alternation of 10 and 20 days) and the same tracks, with their known mean profile, to be used. 2.3. Operational conditions Here again, severe constraints are imposed by the meso-scale mission. Continuity of service must be ensured, although temporary degradations of the sampling, and thus of the number of satellites, is tolerable. These degradations can be quantified at the level of the value of the service provided, allowing the availability objective for the constellation, and thus its maintenance costs, to be adjusted depending on
A meso-scale mission may be restricted to a few areas of operational interest. This is not the case for general circulation missions which are, in essence, of a global nature. Nevertheless, analyses show that, for optimal extraction of the meso-scale signal, measurements are needed from well beyond the useful area and there is thus a synergy between the two areas of interest. 2.5. TopographicaI performance Apart from the intrinsic performance level of the altimeter, the elements that affect the topographical performance are: -whether or not accurate orbit reconstitution is available. This, in fact, depends on whether DORIS is on-board, since this is the only system that has been validated in the required performance range (a few centimetres) to date. The importance of its presence depends on the type of mission; -the quality of the tropospheric propagation error correction. This requires the presence on-board of a three-frequency radiometer and the use of pressure fields provided by the meteorological service; -the quality of the ionospheric propagation error correction. Up to now, it has been necessary to use the best available models of the ionosphere if the heavy solution of a two-frequency altimeter was to be avoided. On all these points, the meso-scale mission adds no new constraints with respect to the large-scale needs (climatology, general circulation). DORIS orbito-graphy even proves to be less essential in this case. Nevertheless, this two-frequency ranging system has shown that it is able to provide ionospheric corrections applicable to the single frequency altimeter measurements. Moreover, it reduces the load on (and the cost of) the ground network of tracking stations by supplying the ranging necessary for the operations. We can conclude that, even within a framework strictly limited to the meso-scale, there is no justification for doing without DORIS but, on the other hand, its presence is a prerequisite if the system is to be extended to cover scientific and environmental needs. 3. SYSTEM CONSTRAINTS
3.1. Launch jhcilities The current reference launcher for size and mass consideration is PEGASUS, which has already been
Lightsat constellation
431
flown twice and for which the U.S.A. have already placed many firm orders. Its main features are its low cost (USSS M) but also its modest performance with respect to the mass it can satelhze and the volume inside the fairing. We recall that the originality of this launcher lies in the concept of being drop-launched from an aircraft. The TAURUS launcher (U.S.A.), which lifts off in the conventional way from the ground, offers performance levels and cost that are more than twice as high. Theoretically, this would allow the double launch of two satellites that would have been launched separately by PEGASUS. The existence of TAURUS is far from sure since it is still only a study on paper but this launcher is a fairly good illustration of a need which will almost certainly be filled in the future, possibly by a European initiative based on a derivative of ARIANE 5 (advanced project name DLA-P).
three additional, altimetry satellites because DORIS relieves the stations of the localization function and because a local time of 06.00/18.00 h is not used by the other “customers” of the network as yet. The control centre is specific. It simultaneously follows the orbit of the three satellites plus, possibly, a fourth on standby in orbit. It is also responsible for the station acquisition operations whenever a new satellite is launched, these operations being doubled if the launcher has the capacity for a double launch. Apart from the number of satellites controlled, the specificity of this centre lies in the accuracy of its orbital mechanics operations: accurate acquisition and maintenance of the ground track, keeping the phase difference between the three satellites, reconfiguring to two satellites when a failure occurs, reactivating and repositioning a satellite left on standby in orbit, etc.
3.2. Choice of orbit
Replacement in orbit. Nominally, three active satellites are necessary. Maintaining a continuous operational service supposes a procedure for replacing satellites at the end of their lives. One can decide to launch a new satellite when a failure occurs. This “on call” approach can only be envisaged if launch delays are not too long. A warning threshold must be defined, i.e. the minimum number of satellites in working order for a warning not to be issued. This will be three in principle but may be four if an in-orbit spare, easy to reactivate, is required at all times (such as in the case of really long launch delays). If double launches are used, the threshold will have to be set differently (to two and three, respectively). It is also possible to launch a new satellite (or two for a double launch) at regular intervals, a satellite being replaced whether it is in working order or not. This “scheduled” approach is less flexible but may result in lower launch costs. It is necessary to “adjust” the interval between two launches.
Local time. Unless it is counter-indicated
for the mission, heliosynchronism is preferred in low orbit as this offers the advantage of fixing the satellite’s energy and thermal configuration with respect to the Sun. A local time of 06.00/18.00 h is favourable for the energy budget as the solar panels are always illuminated except for a short eclipse during solstice seasons. Satellite studies have shown that in these conditions it is worthless to limit the payload duty cycle for matching a given ground coverage required by the meso-scale and that permanent operation (except during eclipse) of the payload and quasi global coverage, as required by scientific applications, can be obtained without oversizing. Altitude and orbital cycle. Within the above constraints on the orbital cycle, there are many candidate orbits, differentiated by their altitudes. With PEGASUS, the mass that can be placed in heliosynchronous orbit varies little in the 600-800 km altitude range, part of the difference in favour of low altitudes being compensated by the extra hydrazine needed to overcome atmospheric friction. Again because of friction, track-keeping to within + 1 km may involve too high a frequency of manoeuvres, making extraction of the meso-scale signal difficult, if a minimum orbital altitude is not respected. This trackkeeping quality is indispensable because of the uncertainty on the transverse slopes of the geoid. All this leads us to consider the nominal orbit to be 750 km with a 30.day cycle (14 + 13/30). 3.3. Operations and ground control facilities Satellites in low-Earth orbit require tracking stations at various points of the globe to monitor and control their orbit and operation. These considerable technical and human resources cannot be specific to a single mission. The CNES network of stations is used, among other things, for the SPOT and HELIOS missions, and it is only possible to take charge of
3.4. System deployment and maintenance
3.5. Redundancy and satellite lifetime The failure of the altimeter obviously leads to the satellite being abandoned and replaced (“on call” case). However, if the radiometer or DORIS fails, it may be decided to continue using the satellite and not to replace it. The operational attitude has to be optimized taking into account, on the one hand, the prejudice caused by the loss of these functions to the service provided and, on the other, the savings made by tolerating these losses. Including redundant parts may increase the satellite’s life but also increases the cost, so no immediate conclusions can be drawn as to which is the better approach from an economic point of view. 3.6. Optimizing the maintenance strategy The number of design parameters (lifetime, launch delays) and running parameters (warning threshold,
432
and G. Bellaiche
J. P. Aguttes
time between satellite deliveries and launches, etc.) makes it difficult to set up a deployment and maintenance strategy for the system. All these parameters have the influence of the annual running cost of the system, but optimization is only possible because they also have an impact on the value of the service provided by the system. The latter may be connected with how long the system stays in the various degraded states, through the reduction or prejudice coefficient (absence of one or two satellites of the DORIS function, etc.). The approach in this study is based on the implementation of modelling and Monte Carlo simulations (see Section 5).
4. THE ALTIMETRY
LIGHT SATELLITE
4.1. Options considered Under the study performed by CNES, AEROSPATIALE studied the satellite configurations capable of satisfying the various system and mission constraints mentioned above. Two cases were studied, depending on the intended lifetime (2 or 4 years) and thus on the electronic redundancy to be implemented. For each of the options, compatibility for a single launch on PEGASUS, a double launch on TAURUS and a triple launch with AR6 DLA-P is sought. The aim of this compatibility is to ensure that launches are possible for this mission for a long time to come.
\ Radiometer
Earth DORIS
and altimeter
sensor
antenna
Y
PEGASUS
TAURUS
LAUNCH
Fig.
I
LAUNCH
433
Lightsat constellation Figure 1 shows the satellite option with redundancies in flight configuration and its launch configurations PEGASUS and TAURUS. The main design difficulties stem from:
-the small volume under the launcher fairing, which makes it necessary to fold the common radiometer/altimeter antenna for launch; -the total acceptable launch mass, a particularly severe constraint for PEGASUS; -the target cost, which makes it necessary to seek commonality with future programmes in the design of the platform (i.e. a spacecraft without a payload).
The propulsion system provides precise orbit control and is also used to circularize the orbit after injection for the case of a PEGASUS launch. The system uses a 34 kg capacity hydrazine tank and 6 thrusters of 5 N. It is important to keep attitude control requirements to a minimum for light satellites. The chosen concept, using a single momentum wheel, ensures a control of +0.45 and a reconstitution of +0.2. The system uses three magnetotorquers and the wheel as actuators plus magnetometers and Earth sensors. This performance is favoured by the dawn-dusk orbit which reduces the pertubation constraints.
4.2. Payload
4.4. Compatibility
We recall that DORIS and POSEIDON altimeters already exist. DORIS, made by Dassault Electronique, is currently flying on SPOT 2 and will soon fly on SPOT 3, SPOT 4 and TOPEX/POSEIDON. Its performance relies on an ultra-stable oscillator (OUS), made by FE1 (U.S.A.) for SPOT 2 and by CEPE (France) for the rest. POSEIDON is made by Alcatel Espace and will have its first flight on TOPEX/POSEIDON. The three-frequency radiometer remains to be developed. The technology and know-how exist in France at Alcatel and Matra; it will draw much of its inspiration from the ERSl two-frequency radiometer.
The mass of the satellite is 204 kg without redundancy, and 270 kg with it. These figures, which include development margins, show full compliance with PEGASUS whose performance is estimated at around 330 kg for injection on the elliptic transfer orbit. TAURUS offers very comfortable margins, including those for double launches. The double configuration is not an easy one to obtain, however, considering the limited dimensions of the fairing. Let us not forget that neither the launcher nor its cost is a reality at present. It would be good if the additional cost engendered by the use of this higher performance launcher were accompanied by an increased fairing diameter, thus easing the arrangement of the antennas or facilitating multiple launch. In this way AR6 DLA-P has been shown to be able to launch the three satellites together.
4.3. Platform Electrical power is supplied by fixed solar panels (315 W at the end of life), the size of which can be very small thanks to the use of GaAs cells. A NiCad battery of 4 Ah allows to cope with the eclipse. Communications with the ground are carried out through an omnidirectional S-band antenna, both for control links and for instrument data telemetry. These are stored on-board during periods when no reception stations are visible. The dumping data rate is 160 kbps.
5. TYPICAL
with launchers
OPERATIONAL PERFORMANCE LIGHT SATELLITE SYSTEM
OF A
5.1. Introduction
The results which follow correspond to the operational optimization of the meso-scale mission which, as mentioned before (see Section 2.3) drives the
Service
value
prejudice
For the degradation loss of radiometer
[Reference
cay
coefficients
of each remaining -> 0.8; of DORIS
,,
S/C -> 0.97
,,
High reliability satellite (R = 0.6 at 4 years) single launch wtth PEGASUS on call S/C procurement rate: I S/C each 2 years I
2
3
4
5 Years
6
7
8
9
10 II
12 13 14 15 16 17 18 19 20
from the S/C production
start Fig. 2
AA 3517-s
434
J. P. Aguttes and G. Bellaiche Reference case Same but at satellite option with low reliability (R = 0.5 2 years). procurement rate: I S/C /year
; 100 r
450 95 -
400
90 -
350
85 -
250
300
200 80 -
150 100
75 -
\
1 2
3
4
‘$
7
5 6
8 9 1011
12 13 14 15 16 17 181920
Years from the S/C production
I
2
’ 3
’ 4
” 5 6
” ’ ” ’ ” ’ ” 7 8 9 1011121314151617181920
Years from the S/C production
start
’ T-b
start
Fig. 3
requirements. It is clear, however, that large scale and scientific applications are satisfied by only one satellite and can even cope with data outages provided that long-term continuity is achieved. In this way these applications are perfectly covered by the minimum meso-scale system described here. Above all, these results are intended to illustrate the operational flexibility of the light satellite concept and quantify the key role of certain parameters in the “economic” optimization of the mission. operational
5.2. Running cost and service provided Figure 2 shows the expected availability over 20 years for three different operating states (three, two or one satellites). The first 3 years correspond to the bringing into production of the satellites, the fourth year is used for MOS (mean ocean surface) learning and the service only really starts in the fifth year. This figure corresponds to the reference case issued by the optimization study: the satellite corresponds to the long lifetime option (R = 0.6 after 4 years), launches are single and take place on call, there is no in-orbit spare, the procurement rate of the satellite is 1 S/C every 24 months. n
0
The prejudice coefficients attached to each degraded state allow the value of the service provided, normalized with respect to the ideal case of continuous operation of the three satellites without any degradation, to be derived from the availabilities. Prejudice coefficients are indicated in Fig. 2, those attached to the degradation of the satellite are applied to those concerning the constellation state proportionally to the number of satellited affected among the remaining ones. The satellite degradations are not severe and justify that the satellite is kept in operation in spite of DORIS or radiometer failure. Figures 3 and 4 plot the service value and the annual running costs for the system. These costs include the procurement of satellites (but not the development costs), any pre-launch storage and the launch itself. Ground operating costs (estimated at around 60 MFF/year) have to be added to finally get the overall annual running cost. The procurement rate for the satellites is the major factor in the costs and marks the profile of the curve. This rate also fixes the availability and thus the service. Too low a rate may lead to long “repair” delays for the constellation because there are no satellites ready to be launched.
Reference case Same but dual launch with TAURUS. use of m-orbit spare mode, procurement rate: 2 S/C each 45 months 500
100
450 95
400 350 300
85
250 200
80
150 100
75 JO 1 2
3
4
I
I
I
I I I
I
I
I I I j
I
I
I
5 6
7
8 9 1011121314151617181920
50
01““““““““‘T 1 2
Years from the S/C production start
3
4
5 6
7
8 9 1011121314151617181920
Years from the S/C production Fig. 4
start
Lightsat constellation 5.3. Main results It appears from Fig. 3 (reference case) that the annual running cost of such a system serving, as a matter of fact, the full range of altimetry missions is around 250 MFF/year. Figure 3 shows also that the low reliability satellite is less efficient from the overall economical point of view. The low reliability satellite has no equipment redundancy while for high reliability there are redundancy of the altimeter and the key equipment of the platform. Figure 4 shows that the double launch option with TAURUS is competitive. This approach supposes the use of an in-orbit satellite spare so as to insure that the number of operating satellites is nominally three or two during launch phases and that worse constellation degradation is almost precluded.
6. CONCLUSIONS: THE LIGHTSAT AS A TOOL INTERNATIONAL CO-OPERATION
FOR
The distinction between the different altimetry missions depends mainly on the number of satellites used and the operational requirements concerning the continuity and availability of the service. In contrast, there is no reason to make a distinction between
435
scientific satellites and meso-scale defence satellites, even the orbit can be the same. The light satellite concept is fully adequate for these altimetry missions and allow a great flexibility of the choice of the rocket. The specificity of the launch offers the designer and operator a set of degrees of freedom for optimizing mission economy on condition that the value of the service provided can be quantified. Here, we have illustrated the case where the heaviest mission (mesoscale) is set up from scratch, but advantage can also be taken of the flexibility of the light satellite to upgrade a more or less discontinuous scientific mission (small or large platform) to a mission that is more operational or has a denser sampling pattern. Finally, this flexibility is a unique advantage for the setting up of an international programme, as cooperation can take place with the minimum of constraints between countries and the maximum synergy for the whole. Each participant can work as he pleases according to his own resources and objectives, developing his own satellites, launching them at his own rate, operating them himself. A minimum agreement needs to be set up for the sharing and exchange of data and for the choice of orbits, the most demanding partner in terms of operationality accepting to fill a few “gaps”.
00!&5765@4)0027~2
AcraAstronourica Vol. 3s. No. 1, pp.429-435, 1995 Copyright Q 1995ElscvicrScienceLtd printedin Great Britain. All rights reserved 0094-5765195$9.50+ 0.00
CAPABILITIES OF LIGHTSAT CONSTELLATION FOR OPERATIONAL ALTIMETRY? J. P. AGUTTES
and G. BELLAICHE
Centre National d’Etudes Spatiales, 18 av. Edouard Belin, 31055 Toulouse Cedex, France (Received 16 March 1994; received for publication I November 1994) Abstract-A radar altimeter mission may involve more than one satellite in space, especially when the objective is to collect data on meso-scale features. However, dedicated satellites can be very small thanks to the integrated technology of modem altimeters such as the French POSEIDON instrument. For these reasons, light satellites appear to be an affordable concept for the deployment of this type of constellation and, moreover, give more flexibility than “piggy-backing” altimeter packages on several distinct, heavier, multimission platforms. This paper gives an overview of typical system and mission trade-offs resulting from operational, technological and economic constraints in different domains: payload and satellite configuration, orbital constellation adjustment, ground tracking facilities etc.
I. INTRODUCTION I.I.
The TOPEX/POSEIDON
legacy
The forthcoming launch of the TOPEX/POSEIDON satellite (a French-U.S. co-operative venture) will make concrete reality of nearly 10 years of effort by the CNES to develop and master altimetry techniques and data utilization. The DORIS mission on-board SPOT 2 was an essential first step, providing full-scale validation of the precise orbitographic techniques needed to back up the altimeter measurements. Thus, right from its first flight, the TOPEX/POSEIDON mission will be able to make good use of the measurements from the POSEIDON altimeter, a new generation instrument using highly integrated technology. TOPEX/POSEIDON should confirm the soundness of CNES’s original approach, combining pragmatism of concept (a single frequency of the altimeter, one-way measurements for DORIS) with audacity of technology (solid-state altimeter, new generation of ultra-stable oscillators for DORIS) and leading to a minimum payload of small size that should encourage the setting up of more altimetry missions, either as passengers on host platforms or in small, dedicated satellites. 1.2. New opportunities for the future The topographic sea signal has spectrum, depending on the time and measurement, and altimetry can serve of applications. TOPEX/POSEIDON
a very large space scale of different types is a scientific
__tPaper IAF-92-87 presented at the 43rd Astronautical Congress, Washington D.C., U.S.A., 28 August5 September 1992.
mission optimized for the study of large scale features (more than 5000 km) and their seasonal and multiyear variations. But the scientific community also has interests which involve a smaller space scale observation, like the study of polar ice or the determination of a precise sea geoid. Up to now meso-scale features were mainly related to defence concerns but due to the involved energies and interactions with other scales, such phenomena are also attracting more and more scientists’ interests. This paper mainly draws its inspiration from studies carried out on behalf of the French DOD for military missions. The meso-scale aspect of these missions sets the first conditions of the system as it imposes the use of several detectors in-orbit at the same time. But such systems, as shown here, can continue to serve the scientific community with the addition of some marginal constraints. This synergy and the flexibility of implementation with light satellites should bring new opportunities for the future of scientific altimetry. 1.3. Light satellites The light satellite is a concept causing much excitement in the space world today. Techniques have reached a point that enables important payloads to be made in a small volume, but the real cause of all the enthusiasm is the emergence of small, low-cost launchers. The association of small satellites with small launchers means that even modest missions can take control of their access to space, i.e. above all. the choice of orbit, the date of the first launches and the strategy for replacing satellites in orbit at the end of their lifetime. All this is obviously more difficult when a launcher, or worse, a large space platform, has to be shared with several other missions. 429
430
J. P. Aguttes and G. Bellaiche
The meso-scale altimetry mission is, in this respect, a typical case showing the interest of the combination because of the number of satellites brought into play and the requirements as far as choice of orbit and system availability are concerned.
the economic trade-off decided upon for the mission. An optimal strategy can be defined (see Section 5) fixing the lifetime of the satellites, the modes of replacement in orbit or on the ground and the time between launches. 2.4. System coverage (or service area)
2. TYPICAL MISSION REQUIREMENTS FOR A LIGHT SATELLlTE SYSTEM
2.1. Spatio -temporal sampIing Meso-scale phenomena are characterized by distance scales of 60-300 km and time scales of lo-30 days. With three satellites in orbit, it is possible to take samples on the ascending satellite tracks over 10 days/85 km compatible with the observation of the phenomena, given the oversampling contributed by the descending tracks. 2.2. Orbital cycle and track maintenance The conventional mission approach for obtaining the meso-scale signal consists of using repetitive orbits, keeping on track to within about a kilometre. In a first learning phase, the accumulation of passes allows the mean ocean profile to be reconstructed. This is considered to be stable and is subtracted at each useful pass so that the meso-scale signal can be isolated. There are two ways of obtaining the desired sampling, each considering three satellites following each other at regular intervals along the same nominal orbit: -ach satellite performs the required temporal sampling through the choice of an 11-day orbital cycle. The difference of 120 between the satellites multiplies (because 11 is not divisible by 3) the tracks and ensures spatial sampling; -the opposite: each satellite covers the required spatial mesh with a time cycle of 30 days. The 120 phase difference along the orbit means that the three satellites follow each other over the same track (because 30 is divisible by 3) at lo-day intervals. The 30-day solution is preferable since, if one satellite fails, the other two can be reconfigured at 180. This allows temporal degradation to be kept regular (15 days instead of an alternation of 10 and 20 days) and the same tracks, with their known mean profile, to be used. 2.3. Operational conditions Here again, severe constraints are imposed by the meso-scale mission. Continuity of service must be ensured, although temporary degradations of the sampling, and thus of the number of satellites, is tolerable. These degradations can be quantified at the level of the value of the service provided, allowing the availability objective for the constellation, and thus its maintenance costs, to be adjusted depending on
A meso-scale mission may be restricted to a few areas of operational interest. This is not the case for general circulation missions which are, in essence, of a global nature. Nevertheless, analyses show that, for optimal extraction of the meso-scale signal, measurements are needed from well beyond the useful area and there is thus a synergy between the two areas of interest. 2.5. TopographicaI performance Apart from the intrinsic performance level of the altimeter, the elements that affect the topographical performance are: -whether or not accurate orbit reconstitution is available. This, in fact, depends on whether DORIS is on-board, since this is the only system that has been validated in the required performance range (a few centimetres) to date. The importance of its presence depends on the type of mission; -the quality of the tropospheric propagation error correction. This requires the presence on-board of a three-frequency radiometer and the use of pressure fields provided by the meteorological service; -the quality of the ionospheric propagation error correction. Up to now, it has been necessary to use the best available models of the ionosphere if the heavy solution of a two-frequency altimeter was to be avoided. On all these points, the meso-scale mission adds no new constraints with respect to the large-scale needs (climatology, general circulation). DORIS orbito-graphy even proves to be less essential in this case. Nevertheless, this two-frequency ranging system has shown that it is able to provide ionospheric corrections applicable to the single frequency altimeter measurements. Moreover, it reduces the load on (and the cost of) the ground network of tracking stations by supplying the ranging necessary for the operations. We can conclude that, even within a framework strictly limited to the meso-scale, there is no justification for doing without DORIS but, on the other hand, its presence is a prerequisite if the system is to be extended to cover scientific and environmental needs. 3. SYSTEM CONSTRAINTS
3.1. Launch jhcilities The current reference launcher for size and mass consideration is PEGASUS, which has already been
Lightsat constellation
431
flown twice and for which the U.S.A. have already placed many firm orders. Its main features are its low cost (USSS M) but also its modest performance with respect to the mass it can satelhze and the volume inside the fairing. We recall that the originality of this launcher lies in the concept of being drop-launched from an aircraft. The TAURUS launcher (U.S.A.), which lifts off in the conventional way from the ground, offers performance levels and cost that are more than twice as high. Theoretically, this would allow the double launch of two satellites that would have been launched separately by PEGASUS. The existence of TAURUS is far from sure since it is still only a study on paper but this launcher is a fairly good illustration of a need which will almost certainly be filled in the future, possibly by a European initiative based on a derivative of ARIANE 5 (advanced project name DLA-P).
three additional, altimetry satellites because DORIS relieves the stations of the localization function and because a local time of 06.00/18.00 h is not used by the other “customers” of the network as yet. The control centre is specific. It simultaneously follows the orbit of the three satellites plus, possibly, a fourth on standby in orbit. It is also responsible for the station acquisition operations whenever a new satellite is launched, these operations being doubled if the launcher has the capacity for a double launch. Apart from the number of satellites controlled, the specificity of this centre lies in the accuracy of its orbital mechanics operations: accurate acquisition and maintenance of the ground track, keeping the phase difference between the three satellites, reconfiguring to two satellites when a failure occurs, reactivating and repositioning a satellite left on standby in orbit, etc.
3.2. Choice of orbit
Replacement in orbit. Nominally, three active satellites are necessary. Maintaining a continuous operational service supposes a procedure for replacing satellites at the end of their lives. One can decide to launch a new satellite when a failure occurs. This “on call” approach can only be envisaged if launch delays are not too long. A warning threshold must be defined, i.e. the minimum number of satellites in working order for a warning not to be issued. This will be three in principle but may be four if an in-orbit spare, easy to reactivate, is required at all times (such as in the case of really long launch delays). If double launches are used, the threshold will have to be set differently (to two and three, respectively). It is also possible to launch a new satellite (or two for a double launch) at regular intervals, a satellite being replaced whether it is in working order or not. This “scheduled” approach is less flexible but may result in lower launch costs. It is necessary to “adjust” the interval between two launches.
Local time. Unless it is counter-indicated
for the mission, heliosynchronism is preferred in low orbit as this offers the advantage of fixing the satellite’s energy and thermal configuration with respect to the Sun. A local time of 06.00/18.00 h is favourable for the energy budget as the solar panels are always illuminated except for a short eclipse during solstice seasons. Satellite studies have shown that in these conditions it is worthless to limit the payload duty cycle for matching a given ground coverage required by the meso-scale and that permanent operation (except during eclipse) of the payload and quasi global coverage, as required by scientific applications, can be obtained without oversizing. Altitude and orbital cycle. Within the above constraints on the orbital cycle, there are many candidate orbits, differentiated by their altitudes. With PEGASUS, the mass that can be placed in heliosynchronous orbit varies little in the 600-800 km altitude range, part of the difference in favour of low altitudes being compensated by the extra hydrazine needed to overcome atmospheric friction. Again because of friction, track-keeping to within + 1 km may involve too high a frequency of manoeuvres, making extraction of the meso-scale signal difficult, if a minimum orbital altitude is not respected. This trackkeeping quality is indispensable because of the uncertainty on the transverse slopes of the geoid. All this leads us to consider the nominal orbit to be 750 km with a 30.day cycle (14 + 13/30). 3.3. Operations and ground control facilities Satellites in low-Earth orbit require tracking stations at various points of the globe to monitor and control their orbit and operation. These considerable technical and human resources cannot be specific to a single mission. The CNES network of stations is used, among other things, for the SPOT and HELIOS missions, and it is only possible to take charge of
3.4. System deployment and maintenance
3.5. Redundancy and satellite lifetime The failure of the altimeter obviously leads to the satellite being abandoned and replaced (“on call” case). However, if the radiometer or DORIS fails, it may be decided to continue using the satellite and not to replace it. The operational attitude has to be optimized taking into account, on the one hand, the prejudice caused by the loss of these functions to the service provided and, on the other, the savings made by tolerating these losses. Including redundant parts may increase the satellite’s life but also increases the cost, so no immediate conclusions can be drawn as to which is the better approach from an economic point of view. 3.6. Optimizing the maintenance strategy The number of design parameters (lifetime, launch delays) and running parameters (warning threshold,
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time between satellite deliveries and launches, etc.) makes it difficult to set up a deployment and maintenance strategy for the system. All these parameters have the influence of the annual running cost of the system, but optimization is only possible because they also have an impact on the value of the service provided by the system. The latter may be connected with how long the system stays in the various degraded states, through the reduction or prejudice coefficient (absence of one or two satellites of the DORIS function, etc.). The approach in this study is based on the implementation of modelling and Monte Carlo simulations (see Section 5).
4. THE ALTIMETRY
LIGHT SATELLITE
4.1. Options considered Under the study performed by CNES, AEROSPATIALE studied the satellite configurations capable of satisfying the various system and mission constraints mentioned above. Two cases were studied, depending on the intended lifetime (2 or 4 years) and thus on the electronic redundancy to be implemented. For each of the options, compatibility for a single launch on PEGASUS, a double launch on TAURUS and a triple launch with AR6 DLA-P is sought. The aim of this compatibility is to ensure that launches are possible for this mission for a long time to come.
\ Radiometer
Earth DORIS
and altimeter
sensor
antenna
Y
PEGASUS
TAURUS
LAUNCH
Fig.
I
LAUNCH
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Lightsat constellation Figure 1 shows the satellite option with redundancies in flight configuration and its launch configurations PEGASUS and TAURUS. The main design difficulties stem from:
-the small volume under the launcher fairing, which makes it necessary to fold the common radiometer/altimeter antenna for launch; -the total acceptable launch mass, a particularly severe constraint for PEGASUS; -the target cost, which makes it necessary to seek commonality with future programmes in the design of the platform (i.e. a spacecraft without a payload).
The propulsion system provides precise orbit control and is also used to circularize the orbit after injection for the case of a PEGASUS launch. The system uses a 34 kg capacity hydrazine tank and 6 thrusters of 5 N. It is important to keep attitude control requirements to a minimum for light satellites. The chosen concept, using a single momentum wheel, ensures a control of +0.45 and a reconstitution of +0.2. The system uses three magnetotorquers and the wheel as actuators plus magnetometers and Earth sensors. This performance is favoured by the dawn-dusk orbit which reduces the pertubation constraints.
4.2. Payload
4.4. Compatibility
We recall that DORIS and POSEIDON altimeters already exist. DORIS, made by Dassault Electronique, is currently flying on SPOT 2 and will soon fly on SPOT 3, SPOT 4 and TOPEX/POSEIDON. Its performance relies on an ultra-stable oscillator (OUS), made by FE1 (U.S.A.) for SPOT 2 and by CEPE (France) for the rest. POSEIDON is made by Alcatel Espace and will have its first flight on TOPEX/POSEIDON. The three-frequency radiometer remains to be developed. The technology and know-how exist in France at Alcatel and Matra; it will draw much of its inspiration from the ERSl two-frequency radiometer.
The mass of the satellite is 204 kg without redundancy, and 270 kg with it. These figures, which include development margins, show full compliance with PEGASUS whose performance is estimated at around 330 kg for injection on the elliptic transfer orbit. TAURUS offers very comfortable margins, including those for double launches. The double configuration is not an easy one to obtain, however, considering the limited dimensions of the fairing. Let us not forget that neither the launcher nor its cost is a reality at present. It would be good if the additional cost engendered by the use of this higher performance launcher were accompanied by an increased fairing diameter, thus easing the arrangement of the antennas or facilitating multiple launch. In this way AR6 DLA-P has been shown to be able to launch the three satellites together.
4.3. Platform Electrical power is supplied by fixed solar panels (315 W at the end of life), the size of which can be very small thanks to the use of GaAs cells. A NiCad battery of 4 Ah allows to cope with the eclipse. Communications with the ground are carried out through an omnidirectional S-band antenna, both for control links and for instrument data telemetry. These are stored on-board during periods when no reception stations are visible. The dumping data rate is 160 kbps.
5. TYPICAL
with launchers
OPERATIONAL PERFORMANCE LIGHT SATELLITE SYSTEM
OF A
5.1. Introduction
The results which follow correspond to the operational optimization of the meso-scale mission which, as mentioned before (see Section 2.3) drives the
Service
value
prejudice
For the degradation loss of radiometer
[Reference
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coefficients
of each remaining -> 0.8; of DORIS
,,
S/C -> 0.97
,,
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2
3
4
5 Years
6
7
8
9
10 II
12 13 14 15 16 17 18 19 20
from the S/C production
start Fig. 2
AA 3517-s
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J. P. Aguttes and G. Bellaiche Reference case Same but at satellite option with low reliability (R = 0.5 2 years). procurement rate: I S/C /year
; 100 r
450 95 -
400
90 -
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Years from the S/C production
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’ T-b
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Fig. 3
requirements. It is clear, however, that large scale and scientific applications are satisfied by only one satellite and can even cope with data outages provided that long-term continuity is achieved. In this way these applications are perfectly covered by the minimum meso-scale system described here. Above all, these results are intended to illustrate the operational flexibility of the light satellite concept and quantify the key role of certain parameters in the “economic” optimization of the mission. operational
5.2. Running cost and service provided Figure 2 shows the expected availability over 20 years for three different operating states (three, two or one satellites). The first 3 years correspond to the bringing into production of the satellites, the fourth year is used for MOS (mean ocean surface) learning and the service only really starts in the fifth year. This figure corresponds to the reference case issued by the optimization study: the satellite corresponds to the long lifetime option (R = 0.6 after 4 years), launches are single and take place on call, there is no in-orbit spare, the procurement rate of the satellite is 1 S/C every 24 months. n
0
The prejudice coefficients attached to each degraded state allow the value of the service provided, normalized with respect to the ideal case of continuous operation of the three satellites without any degradation, to be derived from the availabilities. Prejudice coefficients are indicated in Fig. 2, those attached to the degradation of the satellite are applied to those concerning the constellation state proportionally to the number of satellited affected among the remaining ones. The satellite degradations are not severe and justify that the satellite is kept in operation in spite of DORIS or radiometer failure. Figures 3 and 4 plot the service value and the annual running costs for the system. These costs include the procurement of satellites (but not the development costs), any pre-launch storage and the launch itself. Ground operating costs (estimated at around 60 MFF/year) have to be added to finally get the overall annual running cost. The procurement rate for the satellites is the major factor in the costs and marks the profile of the curve. This rate also fixes the availability and thus the service. Too low a rate may lead to long “repair” delays for the constellation because there are no satellites ready to be launched.
Reference case Same but dual launch with TAURUS. use of m-orbit spare mode, procurement rate: 2 S/C each 45 months 500
100
450 95
400 350 300
85
250 200
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Years from the S/C production start
3
4
5 6
7
8 9 1011121314151617181920
Years from the S/C production Fig. 4
start
Lightsat constellation 5.3. Main results It appears from Fig. 3 (reference case) that the annual running cost of such a system serving, as a matter of fact, the full range of altimetry missions is around 250 MFF/year. Figure 3 shows also that the low reliability satellite is less efficient from the overall economical point of view. The low reliability satellite has no equipment redundancy while for high reliability there are redundancy of the altimeter and the key equipment of the platform. Figure 4 shows that the double launch option with TAURUS is competitive. This approach supposes the use of an in-orbit satellite spare so as to insure that the number of operating satellites is nominally three or two during launch phases and that worse constellation degradation is almost precluded.
6. CONCLUSIONS: THE LIGHTSAT AS A TOOL INTERNATIONAL CO-OPERATION
FOR
The distinction between the different altimetry missions depends mainly on the number of satellites used and the operational requirements concerning the continuity and availability of the service. In contrast, there is no reason to make a distinction between
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scientific satellites and meso-scale defence satellites, even the orbit can be the same. The light satellite concept is fully adequate for these altimetry missions and allow a great flexibility of the choice of the rocket. The specificity of the launch offers the designer and operator a set of degrees of freedom for optimizing mission economy on condition that the value of the service provided can be quantified. Here, we have illustrated the case where the heaviest mission (mesoscale) is set up from scratch, but advantage can also be taken of the flexibility of the light satellite to upgrade a more or less discontinuous scientific mission (small or large platform) to a mission that is more operational or has a denser sampling pattern. Finally, this flexibility is a unique advantage for the setting up of an international programme, as cooperation can take place with the minimum of constraints between countries and the maximum synergy for the whole. Each participant can work as he pleases according to his own resources and objectives, developing his own satellites, launching them at his own rate, operating them himself. A minimum agreement needs to be set up for the sharing and exchange of data and for the choice of orbits, the most demanding partner in terms of operationality accepting to fill a few “gaps”.