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MISSION DESIGN OF SCOPE –SMALL SATELLITES FORMATION FLYING MISSION
FOR MAGNETOSPHERIC TAIL OBSERVATION–
Yuichi Tsuda(1), Kiyoshi Maezawa(1), Hirotsugu Kojima(2), Masaki Fujimoto(3), Iku Shinohara(1),
Yoshifumi Saito(1), Ken Higuchi(1), Tomoaki Toda(1) , Takeshi Takashima(1)
(1)Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, 3-1-1, Yoshinodai, Sagamihara Kanagawa, Japan, tsuda@isas.jaxa.jp (2)Research Institute for Sustainalble Humanosphere, Kyoto University
(3)Tokyo Institute of Technology
ABSTRACT
Japan Aerospace Exploration Agency (JAXA) is currently planning the next generation magnetosphere observation mission called “SCOPE.”(cross-Scale Coupling in Plasma universE) SCOPE consists of five satellites, one 450kg mother satellite and four 90kg daughters, which are assumed to be launched by JAXA’s M-V rocket at once. Five satellites fly in basically a tetrahedron formation, minimum 5km apart from each other, at their perigee of 30 Earth-radii distance. To obtain the highly accurate spatial distribution of the magnetospheric phenomena, the quality of the clock synchronization and the relative orbit determination between satellites are essential. SCOPE satellites are to be equipped with the intersatellite communication and ranging system, which is now being developed, to realize the accuracy of 5usec as to clock synchronization and 50m as to the relative distance measurement for that purpose. This paper introduces the current system design and the technical challenges of SCOPE to realize the highly accurate magnetotail formation flying mission.
1. INTRODUCTION
Japan Aerospace Exploration Agency (JAXA) is currently planning the next generation magnetosphere observation mission called “SCOPE”(cross-Scale Coupling in Plasma universE, Fig.1).[1] Its predecessor “GEOTAIL”, launched in 1992, played a big role for understanding the ion behaviour in the Earth’s magnetotail and triggered interests toward further microscopic scale (electrons scale) phenomena. Following this successful work of GEOTAIL, SCOPE aims at observing the Earth’s magnetotail where the ions and electrons interact with each other, with 5 satellites flying in formation. To fully resolve the time-domain behaviour and spatial distribution of the magnetospheric phenomena, a simultaneous observation by spatially distributed electro-magnetic instruments is essential.
The launcher for SCOPE is assumed to be M-V rocket of JAXA, and their orbit is a highly elliptical orbit with
its apogee 30Re from the Earth center. SCOPE consists of one 450kg mother satellite and four 90kg daughter satellites, flying 5 to 5000km apart from each other. The inter-satellite link is used for telemetry/command operation as well as ranging to determine the relative orbit of 5 satellites in a small distance, which cannot be resolved by the ground-based orbit determination. The technical challenges of SCOPE mission are as follows;
(1) 4 daughters are small satellites of less than 90kg including propellant for formation control maneuver.
(2) 5 satellites fly very close to each other in formation. The typical formation is a 5km tetrahedron at apogee.
(3) The spacecrafts’ onboard clocks are synchronized within 5usec by the inter-satellite link.
(4) The relative distance is measured between all the satellites pair by onboard ranging equipments. (5) The relay operation of 4 daughter satellites via
mother satellite to enable the concentrated management and operation.
In this paper, the current status of SCOPE mission R&D is introduced, focusing mainly on the engineering and system design as well as the above technical challenges.
2. ORBIT DESIGN 2.1 Tailbox Definition
The primary place of interest for observation is a tailbox
(Fig.2), where the most dynamic interaction between ionic and electronic effects is expected to be observed. For the engineering purpose, the tailbox is defined as a simple 25Re 10Re 4 Re× × rectangular prism, whose position and orientation is defined in the neutral sheet coordinates.
The neutral sheet coordinates
Σ
ns is a non-inertial, right-handed Cartesian coordinate system, with its x-axis anti-sun directional from the Earth’s centre, y-x-axis perpendicular to the x-axis along the equatorial plane, and z-axis perpendicular to x- and y-axis (Fig.3). The point P in Fig.3 is the point of the highest interest in the tailbox, which is defined in such that, proceed 10Re from the Earth’s centre along the intersection of the equatorial plane and x-z plane, toward anti-sun (+x) direction (now on the position Q in Fig.3), then proceed toward +x direction so that the distance from the Earth’s center become 30Re. From this definition, P is below ecliptic plane in winter, above ecliptic plane in summer as is indicated in Fig.2. In the spring and fall, P moves across the ecliptic plane.2.2 Orbit Insertion
M-V rocket is assumed as the launch vehicle. To insert as much payload as possible, the direct ascent method is applied to reach the 30Re apogee and Re+3000 to
7000km perigee. Additionally, the orbit perturbation is positively utilized for plane change (Fig.4), which is required to insert the spacecraft into the taibox from the M-V launch site of Uchinoura space center. The resulting payload to be inserted in the taibox is 1620kg. The following is the baseline orbit insertion sequence. (1) M-V 3rd stage establishes Re+150km perigee,
14000km apogee.
(2) Perigee up maneuver, establishing Re+600km perigee, 14000km apogee half a period after the launch.
(3) Wait until orbital plane aligns the final orbital plane (for about 2week to 1 month, Fig.4)
(4) Apogee up maneuver, establishing Re+600km perigee, 30Re apogee.
(5) Perigee up manuerver and small orbital plane collection (which maximize the taibox staying period), establishing the final orbit of Re+3000 to 7000km perigee, 30Re apogee.
2.3 Tailbox Encounter Design
The launch window is open throughout a year for this orbital sequence. The launch day determines when SCOPE encounters the tailbox. Because the taibox is placed always in the anti-Sun direction, while the apogee direction is fixed in the inertial space, SCOPE can get into the taibox roughly 1 month a year.
The eclipse analysis indicates that spring and fall taibox encounters are difficult, because of the long ecliptic duration by the Earth (Fig.5). As is explained in the section 2.2, spring and fall taibox are located on the ecliptic plane, the area largely covered by the Earth’s shadow. The current equipment plan cannot allow the battery operation of 5 SCOPE satellites for more than 180 minutes.
On the contrary, the perturbation analysis indicates that the summer and fall taibox encounters are not preferable, because the orbital plane is perturbed much in these orbits due mainly to the lunar gravity (Fig.6). The sensitivity of the inclination on the taibox staying time becomes maximum in the mid-summer and mid-winter, because the semi-major axis aligns the inclination angle direction. In these orbits, the staying time in taibox Fig.2: Tailbox Image
Fig.3: Tailbox Definition (winter tailbox)
decreases drastically year by year. This is not affordable for 4 year baseline SCOPE mission plan.
The trade off between these two conditions suggests four seasonal possibilities for taibox encounter, which are February, April, July and October. SCOPE scientists don’t impose a strict seasonal condition for observation, so that this level of restriction as to system design and flexibility as to observation timing seems feasible for SCOPE mission.
2.4 Orbital Shape Trade-off
The current baseline mission adopts a highly elliptic orbit of 30Re apogee, Re+3000 to 7000km perigee, whose eccentricity ranges from 0.87 to 0.90. The period of this orbit is 3.66 to 3.77days. The orbit encounters the taibox for about 1 month a year, when the apogee aligns anti-Sun direction.
Another possible design is a 30Re circular orbit. Because this orbit can be achieved with just a little bit
increase of delta-V if we use a lunar gravity assist, it is also possible with M-V launch vehicle. Different from the highly elliptic orbit this orbit is expected to encounter the taibox evenly throughout a year. But because the period of the orbit is as long as 9.4days, the total staying time within the taibox is almost the same as the highly elliptic orbit. Moreover, the circular orbit always crosses the tailbox at the same altitude, while the highly elliptic orbit can sweep a whole area of the taibox.
From these reasons, the nominal orbit now is 30Re elliptic orbit, inserted by the direct ascent by M-V launch vehicle.
3. INTERSATELLITE RANGING AND CLOCK SYNCHRONIZATION
3.1 Requirement for Formation Determination To resolve the time and spatial distribution of ionic and electronic phenomena to the significant level, scientists require the spacecrafts’ clocks should be synchronized and the relative positions between the spacecrafts should be obtained throughout observations. The accuracy requirements are shown in Tab.1. The requirements for scientific observations are as follows; (1) The clocks must be synchronized to the accuracy
indicated in Tab.1 throughout the observation. (2) The relative ranges (1-dimensional) should be
obtained on board the spacecrafts with the accuracy indicated in Tab.1, so that on-board data selection and correlation process can be executed.
(3) The 3-dimensional relative positions should be obtained with the accuracy indicated in Tab.1. As a real time determination is not required for this information, the positions are to be calculated on ground based on the telemetry data.
(4) As to the absolute positions with reference to the Earth, only the normal level of accuracy of orbit determination is required.
Tab.1: Accuracy Requirements for Intersatellite Time Synchronization and Ranging
0 100 200 300 400 0 50 100 150 2015-06 γ=0deg 2015-06 γ=+10deg 2015-06 γ=-10deg Accum.
Day in Tailbox [day]
Rev
Fig.6: Taibox Staying Time for 2015.06 Encounter Orbits with Different Final Orbital Plane Collection Angles (1Rev=3.66days)
Because the conventional ground based orbit determination achieves the accuracy of about 1km for SCOPE orbit, these requirements lead to the need for on-board relative ranging and time synchronization system via intersatellite communication link.
3.2 Ranging and Clock Synchronization System The range between mother and daughter satellite is measured by PN code signal-reproduction ranging, where the daughters loop-back the signal from mother satellite, and the mother compares the PN code phases of the forwarded signal with the returned signal to obtain the distance information. At the same time, the daughter reads the PN code from mother and synchronizes the internal clock, so that the clocks of mother and daughter tick at the same speed, as long as the daughter receives the signal from mother.
Let us define the clock of mother (sat0) as
t
0, clocks of daughters (sat1, sat2,…) ast
1,t
2,"
. After the mother-daughter link is established, the clock is synchronized as follows; 1 0 01/
t
= +
t
L
c
(1) 2 0 02/
t
= +
t
L
c
"
where
L
ij is the distance between sati
and satj
,c
is the speed of light.L
0j is known by mother from the round-trip-time of the signal. The returned signal arrives at mother at;10 0
2
01/
t
= +
t
L
c
(2)so that the distance
L
01 is obtained by comparing the phases of forwarded and returned signal;(
)
1
01 2 10 0
L
=
c t
−
t
(3) There are also the links between daughters to measure
the daughter-to-daughter distance. The signal sent by a
daughter as a return link to the mother is to be received by the other daughters as well. The time when sat2 receives the signal from sat1 is expressed as follows;
12 1 12 0 01 12
/
/
/
t
t
L
c
t
L
c L
c
= +
= +
+
(4) From (1) and (4),(
)
12 2 01 12 02/
t
− =
t
L
+
L
−
L
c
(5) BecauseL
01 andL
02 are known from (3), the daughter-to-daughter distanceL
12 can be obtained. In this way, distances between all the pairs of satellites are to be measured.The equation (1)-(5) are the simplified analysis. As the positions of satellites always change due to the orbital motion, and only a pair of S-band link is assumed to be used, a filtering process to estimate momentarily states is required in the actual system.
In the SCOPE mission, the signal from mother (forward link) is broadcasted so that every daughter can synchronizes its clock simultaneously. On the other hand, the return link uses TDMA technique to share one frequency (Fig.7). The return link is switched every 2.5msec, and the transmission slot is handed over one by one to each daughter. Because of this fast switching, the mother satellite can virtually receive the continuous PN signal from all the daughter satellites at once. In this way, the PN return signal and the intersatellite telemetry data of 4 daughter satellites can be sent to the mother satellite on one frequency.
Each spacecraft switches two antennae (top/bottom and side) to cover the whole direction. Each 2.5msec TDMA slot is divided into two parts, and the same signal is transmitted twice from mother using the different antenna. The daughter selects the optimum antenna based on the AGC measurement at the slot allocated for the other daughters, as is indicated in Fig.7.
Currently this onboard relative ranging and clock synchronization system is at the BBM development phase. The data bitrate for mother-to-ground link is Fig.7: TDMA Sequence (Daughter 1 has a right to broadcast the ranging
planned to be 4Mbps (X-band), daughter-to-ground is 400kbps band), and daughter-to-mother is 40kbps (S-band, same frequency as the daughter downlink).
3.3 Formation Determination
The three dimensional shape and orientation, that is to say a “formation”, is obtained by a filtering process on ground. To know the orientation of the formation from the relative ranges, time domain history of the range data is required. The determination accuracy depends on the formation design itself, and how much amount of ranging data we can expect to obtain in every orbital period.
It can be read from Tab.1 that the accuracy requirements for formation orientation must be determined within 1/100 [rad]. According to our analysis so far, the relative range measurement for a few hours per 3.7days orbital period is sufficient to obtain a 1/100[rad] accuracy orbit.
3.4 Mission Phase and Formation Design
The resolution of observation must become the maximum in the tailbox region. For that purpose, SCOPE adopts nominally a tetrahedron formation at apogee, with its size varying depending on the mission phase.
The mission phase is mainly divided into two phases (Fig.8). In the “Phase-1”, all the daughters fly within 100km from mother. In this phase, the intersatellite ranging and clock synchronization are fully utilized to satisfy the accuracy requirements of Tab.1. Additionally, the daughters are designed to be commanded from the ground station via mother satellite. This enables a concentrated operation of all the SCOPE satellites, using the mother as a gateway, to simplify the operation and mission management.
“Phase-2” aims at observing larger scale phenomena, allowing the daughters to fly up to 5000km from mother. As it is too far from mother, and the accuracy
requirement is looser in this range, each satellite is operated independently from the ground, like an ordinary satellite operation.
Over the nominal 4 year mission plan, the first two or three years are for Phase-1, while the rest is for Phase-2. In addition to the nominal tetrahedron formation at apogee, some other key regions will be observed at the maximum observability by optimizing the formation for that region.
4. STRUCTURE
4.1 Structural Configuration
SCOPE is launched by M-V launch vehicle at once. In the launch configuration, daughters are built up vertically on the mother satellite (Fig.9). After inserted into the initial orbit of Re+150km
×
14000km, the bipropellant thruster equipped on the mother takes them all together to the final orbit of Re+3000km to 7000km×
30Re apogee. After that, each daughter is undocked one by one, and moves to the objective relative position to form the formation by its own RCS equipment.The mother and 3 daughters are spin-stabilized satellites with their spin axes perpendicular to the orbital plane. The rest one daughter is also spin-stabilized, but its spin axis is oriented toward the Sun. This special daughter satellite is called “near-daughter”, as it is to fly within 100km even in the Phase-2 operation. The combination of the mother and the near-daughter enables us to obtain 3-dimensional information as to the electric field, measured by the wire antennae.
The mother satellite’s dry mass is 450kg, wet 1300kg. To accommodate the satellite’s body within the M-V envelope diameter of 2.2m, a long body propellant tank, and four long body oxidizer tanks are to be equipped, which are illustrated in the right figure of Fig.9.
The daughters are 90kg, 1300mm diameter, 300mm height cylinder. For weight saving, the outer wall
Fig.9: SCOPE Satellites (left: external view, right: internal view at launch configuration) Fig.8: Mission Phase of SCOPE
functions as a thrust tube as well as the place to attach the solar cells.
The satellites are connected by non-explosive separation devices in the launch configuration, which are being developed newly so that the shock at the separation should not affect the scientific instruments placed very close to the separation devices in the thin small daughter satellites.
4.2 Science Instruments
The mother satellite has four 50m wire antennae, two 3-5m radial masts, and two 3-5m spin-axis antennae, all of which are parts of science instruments (Fig.10). The daughter satellites have the same configurations as to masts and antennae (Fig.11), except for the near daughter.
The near-daughter has shorter (35-40m) wire antennae to reduce the angular momentum, as this satellite is Sun-oriented and its spin axis must be re-Sun-oriented 1deg/day. The spin-axis antenna, that all the SCOPE satellites are equipped with, must have high specific rigidity for the attitude stability of the spinning satellite. We are now developing new light-weight inflatable antenna, having sufficient rigidity for the SCOPE satellites’ spin rate of 20[rpm].
The science payloads are listed in Tab.2. 13 kinds of measurements covering wide energy range of electron, ion and magnetic fields are possible with the mother, while the daughters support 5 kinds of measurement, so that the combining and correlating of these data enables us to fully understand the 3-dimensional transient behaviour of the electro-magnetic phenomena in the tailbox.
5. CONCLUSION
This paper describes the current status of SCOPE mission research and development. SCOPE is now at the pre-phase A study, aiming at the launch in mid 2010s. SCOPE consists of 5 small satellites flying in formation. Their relative orbital motion must be measured accurately to determine the spatial distribution of the ionic and electronic phenomena in the tailbox region. An on-board ranging and time synchronization equipment is now in BBM development phase, which
enables the relative orbit determination by means of radio wave measurement. New structual, operation and orbit-control concepts are being adopted so that the ISAS’s first full-scale formation flying mission should become possible.
References
1. Fujimoto, M., Saito,Y., Tsuda,Y., Shinohara,I., Kojima,H., “The scientific targets of the SCOPE mission”, COSPAR04-A-01348,D2.3/E3.3/PSW2-0016-04, COSPAR 2004
2. Toda,T., Saito,Y., Tsuda,Y., “Communication System Design for Magnetospheric Formation Flying Mission: SCOPE”, IAC-05-B3.6.06, 55th International Astronautical Congress, 2005
Tab.2: Science Instruments on SCOPE Fig.10: Mother Satellite in
