Single-axis disturbance compensation system

The invention relates generally to an improved system for sensing and compensating for external disturbance forces acting on a satellite while in orbit. A proof mass member is housed within an enclosure and shielded from external, non-gravitational forces. The proof mass is electromagnetically levitated to move in a purely gravitational orbit, along an axis aligned with the satellite's velocity vector. The proof mass is subjected to a controlled magnetic biasing field and is caused to have a constant reaction to the resultant biasing force, by means of a thermal control system which maintains constant resistivity of the proof mass, during operation. The position of the proof mass with respect to its axis is detected optically and is utilized to control the firing of spacecraft thrusters. As a result, the satellite is caused to maintain a substantially constant position relative to the proof mass and thereby also is caused to follow a purely gravitational orbit.

BACKGROUND OF THE INVENTION 
It is generally well-known in the art of satellite control to provide an 
unsupported proof mass aboard a satellite, shielded from external 
non-gravitational forces so that it follows a purely gravitational orbit 
during operation of the satellite, and means responsive to motion of the 
satellite relative to the proof mass for controlling thrusters onboard the 
spacecraft which force the spacecraft to also follow the gravitational 
orbit, free from the effects of external surface forces such as solar 
radiation pressure and atmospheric drag. As a result, satellite position 
in orbit is predictable well in advance, thus significantly increasing the 
value of the satellite for navigational purposes, for example. 
One major problem which was encountered in the use of early disturbance 
compensation systems of this type involved the need for very accurately 
determining mass attraction forces between satellite-carried components 
and the proof mass, in order to prevent such mass attraction forces from 
influencing the proof mass and thereby providing faulty operation of the 
system. In an effort to overcome the problem of mass attraction forces, 
U.S. Pat. No. 3,785,595 proposed the use of movable compensation masses or 
chargeable magnets (reacting with a diamagnetic proof mass), on three 
orthogonal axes relative to the proof mass, which could be controlled by 
commands from the ground tracking station, once the spacecraft was in 
orbit, to set up forces which counterbalance the mass attraction forces. 
SUMMARY OF THE INVENTION 
The present invention provides a single-axis disturbance compensation 
system much simpler than the three-axes compensation system previously 
proposed and disclosed, for example, in U.S. Pat. No. 3,785,595. More 
particularly, the present invention utilizes a proof mass fabricated from 
electrically conductive material which is suspended electromagnetically, 
by eddy current forces, to counterbalance mass attraction forces due to 
the spacecraft, while moving with essentially no friction along an axis 
aligned with the velocity vector of the spacecraft. The proof mass is 
housed in an enclosure attached to the spacecraft, shielded from all 
non-gravitational exterior forces such as solar radiation pressure, 
atmospheric drag and electrostatic attraction forces, so that the proof 
mass follows a purely gravitational orbit. A controlled magnetic biasing 
field is generated within the enclosure adjacent to the proof mass and 
exerts a preselected and controlled eddy current biasing force level on 
the proof mass in order to balance any along axis constant component of 
proof mass disturbance force. Essentially, this biasing system comprises 
two oppositely wound coils each configured to provide a constant force 
over the proof mass range of motion. In order to assure that the proof 
mass always reacts in a constant fashion to the biasing magnetic field, a 
thermal control system is provided in order to assure constant resistivity 
of the proof mass materials. 
When the satellite is placed into orbit and is controlled to assume its 
orbital configuration, the proof mass is positioned at the center of mass 
of the spacecraft in alignment with the velocity vector. An optical 
detection system then monitors any movement of the proof mass both along 
and transverse to the velocity vector, as results from external forces on 
the satellite due to solar radiation pressure and atmospheric drag. The 
output of the optical detection system is utilized to operate thruster 
control apparatus which force the satellite to maintain a substantially 
constant relative position with respect to the proof mass and whereby the 
satellite is also caused to follow the purely gravitational orbit. 
In light of the above, one object of the present invention is to provide an 
improved system for controlling a satellite to maintain a purely 
gravitational orbit, not affected by external forces such as solar 
radiation pressure and atmospheric drag. 
Other objects, purposes and characteristic features of the present 
invention will be pointed out as the description of the invention 
progresses and will in part be obvious from the accompanying drawings.

Referring now to FIG. 1, a gravity-gradient stabilized satellite is 
illustrated in its orbital configuration; that is, four solar panels 11 
are deployed from the main satellite body 12 along with the 
gravity-gradient stabilizing boom 13. In this configuration, the 
satellite's communication antenna 14 is pointing towards Earth, for 
example. The satellite is yaw-stabilized, in a conventional manner, by a 
momentum wheel (not shown) which generates an angular momentum vector 
which tends to align itself with the orbit normal; the accuracy of the yaw 
stabilization being dependent upon the requirements of practice. In one 
practical application, for example, attitude stability was designed with a 
3.degree. maximum amplitude periodic librations for roll, pitch and yaw 
which average to zero about the respective bias angles of .+-.0.5.degree. 
in yaw and .+-.1.5.degree. in pitch relative to the velocity vector. 
Variations in the yaw bias angle due to thermal boom bending, solar panel 
deflection, and other intermittent sources were designed to cause less 
than 0.042.degree. deviations in the yaw bias angle over a preselected 
time period, e.g. seven days. 
In accordance with the present invention, the satellite carries a 
disturbance compensation system package, generally designated at 15, which 
is carried on the boom assembly 13 and is positioned in alignment with the 
satellite velocity vector when in orbit (designated as the .+-.X axis) and 
at the known center of mass of the satellite when in its illustrated 
orbital configuration. In the practical application noted above, the 
proposed sensor unit 15 was designed for alignment of its axis within 
.+-.0.2.degree. in yaw and within .+-.1.0.degree. in pitch of the velocity 
vector. 
FIG. 2 of the drawings illustrates, in somewhat simplified form, the 
structural details of the sensor unit 15 comprising the preferred 
embodiment of the present invention. As shown, the sensor unit 15 is 
essentially cylindrical in over-all configuration and comprises a pair of 
mating cylinders 20 and 21 fabricated from suitable non-conductive 
materials. In one practical application, the outer cylinder 21 is 
fabricated from grade G-10 epoxy fiberglass; whereas, the inner cylinder 
20 is fabricated from beryllium oxide which has high thermal and low 
electrical conductivity. 
A disassembled view of the cylinder members 20 and 21 is shown in FIG. 3. 
The inner cylinder 20 carries a bifilar Nichrome heater winding 22 which 
is wound over the length of the cylinder 20 and is connected, via its 
right-hand end in FIG. 3, to a suitable energizing source (not shown) to 
be described hereinafter. The cylinder 20 also carries a thermistor 23 
which is mounted, e.g., within a suitable slot in the cylinder wall, so as 
to monitor the temperature within the cylinder 20. The thermistor 23 is 
connected electrically by wire 24 to a temperature control arrangement, to 
be described, whereby the internal temperature of the sensor unit 15 is 
maintained constant. The cylinder 20 is formed with end and intermediate 
ridges 25 which engage against corresponding ridges on the internal 
surface of the slightly larger diameter cylinder 21, as can be best seen 
in FIG. 2. A slight space is thereby created between the cylinders 20 and 
21 to accommodate the windings 22 and the thermistor electrical 
connection. 
The cylinder 21 carries a pair of oppositely wound biasing coils 26, 27 
which are disposed in suitable, spaced notches extending around the 
periphery of the cylinder 21, in groups containing preselected integral 
numbers of windings of the biasing coils. The ridges on the interior wall 
of the cylinder 21 are each provided with four pairs of longitudinal 
slots, each adopted to receive therein a length of resistive type wire 28 
which extends along the inner surface of the cylinder 21 and terminates at 
a suitable connector element 29 whose outermost surface has a threaded 
hole, as shown at 30. In one practical application, the wires 28 are #18 
gauge manganin wire. 
Centrally disposed within the cylinder 20 is an optical system generally 
designated at 31 in FIG. 2. Specifically, the optical system 31 comprises 
a pair of parabolic mirror members 32 which are supported on opposite 
sides of the sensor unit 15 by means of mounting rings 33 whose edges 
engage one of the ridges 25 on the inner surface of cylinder 20. The 
optical system also comprises a light source represented at 34; a beam 
splitting prism 35; a pair of photocells 36 disposed on opposite sides of 
the prism 35; and, a third photocell 37. This third photocell 37 receives 
a beam of light from the source 34 by means of mirrors 38 which are 
mounted at the opposite ends of the cylindrical sensor structure, on the 
longitudinal .+-.X reference axis. 
Light emanating from the light source 34 is basically divided into three 
beam paths: one (shown in dash dot lines) which impinges on the right-hand 
end mirror 38 and is directed from right to left in FIG. 2 between the end 
mirrors 38 and eventually to the photocell 37, and two light beams (one 
shown in short dash lines and the other shown in solid lines) which are 
directed upwardly in FIG. 2 against the upper parabolic mirror 32, then 
downwardly against the lower parabolic mirror 32 and finally to the prism 
35 which directs one of these two beams against each of the oppositely 
disposed photocells 36. The latter two light beams reflected between the 
mirrors 32 are arranged to intercept the ends of a cylindrical proof mass 
member 39, so that the division of light between the photocells 36 
provides indication of the longitudinal position of the proof mass 39 
along the reference axis .+-.X; whereas, the light beam reflected between 
the end mirrors 38 is sized to correspond with the diameter of a central 
aperture through the proof mass 39 so that the light impinging upon 
photocell 37 provides indication of the position of the proof mass 39 
transverse to the reference .+-.X axis. 
Positioned on either side of and abutting the mirror support rings 33 is a 
pair of cylindrical spacer members 40 also fabricated from grade G-10 
epoxy fiberglass, for example, which have a plurality of holes fabricated 
therein for communicating the heat produced by energization of the heater 
coil 22 into the interior of the sensor unit 15 so as to maintain the 
temperature of the proof mass 39 constant. Abutting the extending ends of 
the spacer members 40 are a pair of cup-shaped epoxy fiberglass members 
42, each formed with a central aperture 42a. The interior surface 42b of 
the members 42, the interior surface 40a of the spacers 40 and the exposed 
surfaces of the optical system support structure are made electrically 
conductive, for example, by painting them with a black silicate paint 
doped with carbon, such as that developed by NASA and known as MSA-94B. 
This renders the internal sensor cavity optically absorbing in order to 
protect the photocells 36, 37 from stray radiation from the optical system 
lamp 34 and, at the same time, develops a common electrically 
equipotential surface within the sensor to remove localized electrostatic 
potential build-up in order to prevent forces of electrostatic attraction 
on the proof mass 39 by all components within the sensor interior cavity. 
This equipotential surface also bleeds off electrical charge created by 
radiation from the lamp 34 and by the incident flux of high-energy 
particles present in the space environment. However, the surface is not 
conductive enough to sustain eddy currents generated by the bias coils 26, 
27 of a magnitude large enough to distort the biasing coil magnetic field. 
The proof mass 39 itself is comprised of a cylinder of highly conductive 
material in which eddy currents can be created. In the practical 
application noted above, the proof mass 39 is formed of 0.410" 
I.D..times.0.670" O.D..times.1.240" long cylindrical shell of pure 
aluminum, with a 10,000 A"20,000 A gold surface plating to prevent 
formation of oxides which create localized electrostatic potential and to 
allow the proof mass 39 to reach electrostatic equilibrium within the 
inner sensor cavity. The proof mass 39 is also purposely designed with a 
very low temperature coefficient of resistivity; i.e., .ltoreq.0.004 
ohms/ohm/.degree.C., so that when operating in the controlled temperature 
environment existing within the sensor 15, the proof mass will react in a 
constant manner to the biasing (eddy current) forces. 
At each end of the sensor unit 15 is a spoked ring member 43 fabricated 
from a suitable plastic such as Lexan. The inner surface of each member 43 
is provided with a collar 44 to support the ends of cylinders 20 and 21, 
in assembled position. The outer surface of each member is formed with 
four radially extending conductive strips 46 (see FIGS. 2, 5, and 6), each 
of which is bifurcated to connect to the four pairs of resistive wires 28 
extending longitudinally along the inner surface of the biasing coil 
cylinder member 21, as previously described and as shown in FIG. 3. In 
particular, the outer extending ends of conductive strips 46 are connected 
to the wire pairs 28 by means of screws 47 which extend through the 
members 43 and into the holes 30 in the wire terminal members 29 shown in 
FIG. 3. 
The opposite inner ends of the conductive strips 46 are interconnected 
electrically by screws 48 which extend through the associated member 43 
and are threaded into a suitable electrically conductive collector member 
mounted on the internal surface of the end member 43, to thereby connect 
the conductive strips 46 in series with a central conductor member 50 
which extends longitudinally through the end mirrors 38 and the proof mass 
39 between the end members 43, on the .+-.X reference axis. At the 
left-hand end of the sensor unit (see FIGS. 4 and 6), the conductor 50 is 
connected to a centrally-disposed disc 51 of conductive material, formed 
with an integral terminal tab 51a. A slotted annular conductor member 52 
surrounds and is insulated from the conductor disc 51 and is electrically 
connected to the conductive strips 46 on the opposite surface of the Lexan 
end member 43 by means of the aforementioned screws 48. 
As shown in FIG. 6, the conductor disc 51 has a centrally disposed base or 
stem portion 51b which is threaded at one end and which extends through 
the Lexan end member 43 to receive a nut 54 which locks the member 51 in 
position. The conductor disc 51 is provided with an aperture which extends 
through the stem portion 51b thereof to accommodate the central conductor 
50, whose ends are threaded to receive a nut 55 which is inserted into the 
stem portion 51b of the conductor disc to firmly attach the ends of the 
central conductor 50 in assembled position in the sensor unit 15. 
The radial conductive strips 46 and the equal relative high resistance 
paths provided by the resistive wires 28 between the ends of the central 
conductor 50 assure even symmetrical distribution of return currents 
lengthwise of the sensor unit 15 and thereby prevent undesired biasing of 
the proof mass 39 which would occur if asymmetrical return currents 
existed. 
Secured to the inner surface of the conductor disc member 51 is a mirror 
assembly for supporting the mirrors disc 38 in an adjustable position on 
the reference .+-.X axis. As shown in FIG. 6, for example, the mirror 
plate 38 is formed with a raised base portion adapted to be bonded, e.g., 
by epoxy, into a circular recess formed in a canted end of a mounting 
pedestal member 56. The member 56 has a base flange 57, around which is 
disposed an anchoring collar 58 for connecting the mirror assembly to the 
conductor disc 51 by means of suitable screws, for example. An annular 
space exists between the base flange 57 and the anchoring collar 58 to 
accommodate a spring washer 60 which resiliently biases the base of the 
mirror mount 56 against the conductor disc 51. Three mirror adjusting set 
screws, one of which is shown in FIG. 6 at 61, are disposed at regular 
120.degree. intervals around the conductor 51 to adjust the position of 
the mirror 38 against the biasing force of the spring washer 60, in order 
to assure that the light beam reflected between mirrors 38 is properly 
aligned with the reference .+-.X axis, the central aperture in the proof 
mass 39, the photocell 37, and the light source 34. 
At the right-hand end of the sensor unit 15, in FIG. 2, is a caging 
mechanism for holding the proof mass 39 in fixed position, during 
launching of the satellite and prior to deployment of the boom assembly 
13, with the sensor unit 15 attached, into the orbital configuration. 
FIGS. 7 and 8 of the drawings illustrate the details of the proof mass 
caging mechanism. In particular, the caging mechanism comprises a pair of 
substantially L-shaped caging fingers 70 which are adapted to engage the 
upper and lower surfaces of the proof mass 39, by means of a central ridge 
which extends around the proof mass 39 and mates with notches on the inner 
surface of each caging finger 70. Each caging finger 70 has an extending 
leg portion which extends through a holder 71 and terminates in an 
enlarged cap 72 installed at its end. The holder 71 is secured, by means 
of mounting ears 71a, to the single conductor disc 52a which is disposed 
within a central circular recess 43a in the Lexan end member 43. The 
conductor disc 52a is connected electrically to the four conductive strips 
46 (not shown) that extend along the spokes of member 43 on the outermost 
right-hand surface thereof and connect the right-hand ends of the 
resistive wires 28 to the disc 52a (and central conductor 50). A coil 
spring 73 is disposed between the enlarged cap 72 of the finger 70 and the 
upper surface of the holder member 71, as illustrated in FIG. 8. The 
forward or left-hand surface of the holder members 71 is provided with a 
partial slot 74, best shown in FIGS. 2 and 7, which is slightly wider than 
the extending leg portion of the finger 70 so that the fingers 70 may move 
radially upward from, i.e., uncage, the proof mass 39 as will be discussed 
shortly. 
The fingers 70 are retained in the caging or locking position relative to 
the proof mass 39 by means of plunger assemblies 75 each comprising a 
housing 76, which is fixed to the end member 43, and a spring-loaded 
plunger element 77 which extends through a hole in the conductor disc 52a, 
through the rear surface of the caging holder 71 and through the extending 
leg of the caging finger 70. Each plunger mechanism 75 is provided with a 
locking pin 78 which maintains the plunger 77 in its depressed position 
(as shown in FIG. 8) prior to mounting the sensor unit 15 onboard the 
satellite. Once the sensor is mounted on-board the satellite, in its 
prelaunch configuration, the plungers 77 are maintained in their depressed 
position by means of a suitable plate member or the like, shown dotted at 
79 in FIG. 8, which is anchored to the satellite body in a position such 
that it contacts the ends of the plungers 77 until the boom assembly 13 is 
deployed to an orbital condition (see FIG. 1). At this time, the 
spring-biased plunger elements 77 are ejected and the springs 73 urge 
fingers 70 out of contact with the proof mass 39 and thereby release the 
mass 39. 
With current applied to the central conductor 50, an eddy current 
suspension force is exerted radially on the proof mass 39 proportional to 
the radial offset of the proof mass and to the square of the applied 
current. This force acts to electromagnetically levitate the proof mass 39 
relative to the conductor 50. As will be described hereinafter, when 
considering operation of the sensor unit 15, the proof mass 39 is also 
acted upon by a biasing force directed along the reference .+-.X axis, 
also resulting from repulsion of eddy currents induced in the proof mass 
39. 
When the satellite is first placed in its orbital configuration, the proof 
mass 39 will slowly migrate along the .+-.X axis as the satellite is 
subjected to acceleration/deceleration as a result of solar radiation 
pressures and atmospheric drag. During this migration, the proof mass 39 
will pass through the optical detection region of the sensor 15 and will 
be captured in its desired operating position, by a control system now to 
be described. 
FIG. 9 of the drawings illustrates, in block diagram form, the control 
circuitry associated with the sensor unit 15. In particular, a pick-off 
electronics unit 80 of conventional circuit design is utilized to energize 
the sensor lamp 34 from a 20.8 KHz clock supply carried onboard the 
spacecraft and also to monitor the output of the photocells 36 and 37, in 
order to detect along axis and cross or transverse axis positions of the 
proof mass 39. An on-board command data receiver shown in dotted form at 
81 is connected, via any convenient communications link, to a ground 
control station. The command data received at 81 is input to a data 
command decoder 82 which detects or decodes the incoming commands and 
outputs them to the remainder of the control system. The command decoder 
82 is also connected to a clock frequency source 83 and outputs three 
basic clock frequencies; i.e., a 20.8 KHz clock which is applied via the 
pick-off electronics 80 as the AC energizing source to the lamp 34, a 2.08 
KHz clock which is applied as input to a suspension current generator 84, 
and a 1.04 KHz clock which is supplied to the illustrated bias coil 
current generator 85. 
The suspension current generator 84 is connected by wires 84a, b (see FIG. 
4) between the central conductor disc 51 and the surrounding conductor 
ring 52, to thereby provide an energizing source to the series circuit 
comprised of the central conductor 50, the conductive strips 46 at each 
end of the sensor, and the connected resistive wire pairs 28 which extend 
along the inner surface of the cylinder 21, as shown in FIG. 3. The 
resultant A.C. current flow through the axial conductor 50 produces an 
eddy current repulsion force on the proof mass cylinder 39 which levitates 
or centers the proof mass about the conductor 50 for frictionless motion 
along the reference .+-.X axis. This suspension force, commandable to 
different levels, maintains the proof mass 39 suspended in directions 
normal to the reference axis (and the central conductor 50) and 
compensates for all external forces normal to this axis, e.g., solar 
radiation pressure and mass attraction forces resulting from nearby 
spacecraft-carried components. 
Forces on the proof mass 39, directed along the reference axis, are 
controlled by a biasing system which provides a vernier compensation for 
spurious, constant level forces acting on the proof mass 39, along the 
reference axis, for example as a result of mass attraction forces produced 
by an unbalanced location of spacecraft-carried components. Initialization 
of the bias force acting on the proof mass 39 is derived from tracking 
data indicating orbit anomalies being experienced by the satellite during 
its normal operation. 
The in-orbit vernier control of the along-track bias forces on the proof 
mass 39, e.g., in order to compensate for mass attraction forces along the 
.+-.X axis, is provided by bias current amplitude data output from the 
decoder 82, on line 87, to the bias coil current generator 85. A bias coil 
select signal on line 88 determines whether biasing coil 26 or the 
oppositely wound biasing coil 27 should be energized, depending upon 
whether the spacecraft is tracked as accelerating or decelerating from its 
nominal orbit velocity. The bias coil current generator 85 responds to the 
signals on lines 87, 88 and the 1.04 KHz clock from decoder 82 to generate 
a 1.04 KHz sine wave energizing current, to coil 26 or 27, which is 
amplitude modulated by the amplitude data input on line 87. In the 
practical embodiment of the present invention mentioned above, the biasing 
coil current is commandable to a maximum value of 31.4 ma. (producing 
1.22.times.10.sup.-8 g force on the proof mass) in 0.01534 ma. increments, 
with a maximum incremental force of 1.2.times.10.sup.-11 g, thus providing 
precise in-orbit tailoring of the proof mass axial forces. 
As previously described, the biasing coils 26, 27 are designed, i.e., 
arranged in groups on cylinder 21, to produce a linear variation of the 
quare of the magnetic field B along the reference .+-.X axis, in the 
vicinity of the proof mass 39. The resulting bias force, produced by 
repulsion of eddy currents induced in the aluminum proof mass 39 is 
proportional to the integral over the proof mass length (1); i.e., 
##EQU1## 
Therefore, by designing each biasing coil 26, 27 to provide linear 
variation of the quantity B(x).sup.2, a constant axial biasing force is 
produced on the proof mass 39 independent of its axial displacement along 
the reference .+-.X axis. In the practical embodiment of the present 
invention discussed above, 26 ga. copper wire is used to form each biasing 
coil 26, 27, with integral numbers of 14, 3, 8, 8, 10, 6, and 26 turns in 
each of seven groups located -10.160, -6.353, -3.647, -0.478, +2.685, 
+5.819, and +9.009 cm. respectively in an axial position measured in 
opposite lengthwise directions from the center of the proof mass when in 
its normal position (centered as in FIG. 2). 
As previously noted, the photocells 36 are utilized to detect the 
along-track (axial) position of the proof mass 39. More particularly, the 
pick-off electronics unit 80 contains conventional circuitry which 
monitors the difference between the outputs from the photocells 36 and 
produces an error signal on line 90 which is applied to and controls the 
operation of a control loop electronics unit 91. The unit 91 is also of 
well-known design, its details being described on pages 16 and 17 of the 
APL Technical Digest, Volume 12, No. 2 (April-June 1973) published by the 
Applied Physics Laboratory of The Johns Hopkins University. Basically, the 
control unit 91 responds to the along-track position signal on line 90 and 
controls the spacecraft thrusters to adjust the position of the satellite 
relative to the proof mass 39 until the proof mass is returned to its 
initial precise position and thereby compensate for the solar radiation 
pressure and/or atmospheric drag on the satellite which initially caused 
the proof mass to move along the reference axis. Thereby, the spacecraft 
is controlled to follow the purely gravitational orbit of the proof mass 
39. As shown in FIG. 9, the along-track position of the proof mass 39 is 
also supplied, by control loop electronics unit 91, to the satellite 
telemetry system for telemetering back to the ground control station. 
Moreover, if a significant along-track mass attraction force exists on the 
proof mass 39, the resulting optical system output error signal on line 91 
will cause unwanted acceleration/deceleration of the satellite, unless the 
mass attraction force is compensated for, i.e., biased out. As discussed 
above, this compensation is accomplished by selective energization of the 
biasing coils 26, 27 (depending upon whether acceleration or deceleration 
is observed in the tracking data) with a current whose amplitude produces 
the proper amount of along-track biasing force on the proof mass 39 to 
cancel out the effect of the mass attraction force. 
Initialization of the position of the proof mass 39 is controlled by a 
command transmitted from the ground control station to the command data 
receiver 81, e.g., based upon the output of the optical system 31 as 
telemetered to the ground station. Thus, if the proof mass 39 is not at 
the center of mass of the orbiting satellite, modulation will be observed 
in the output of the transverse photocell 37 as the satellite orbits. In 
response, the ground station transmits an appropriate position bias data 
command which is outputted on line 86 by decoder 82, and summed with the 
output of photocells 36 in the pick-off electronics 80 to produce a 
position bias for the proof mass 39. As a result, an along-track position 
error appears on line 90 which actuates the thruster control, via control 
unit 91; the axial position of the proof mass 39 is thereby adjusted, 
placing it precisely at the center of mass of the satellite when in an 
orbital configuration (as verified, for example, by observing reduced 
modulation in the telemetered optical system output). 
As noted earlier, the proof mass 39 is electromagnetically suspended or 
levitated relative to the reference .+-.X axis by current supplied to 
central conductor 50 by the suspension current generator 84. More 
particularly, the suspension current generator 84 receives a 2.08 KHz 
clock signal from the data command decoder 82 and generates four possible 
amplitude levels of square wave energization current on lines 84a, b 
connected to the suspension circuit; i.e., the 2.08 KHz clock frequency is 
amplitude modulated depending upon the value of the suspension strength 
select signal applied to the generator 84 from the data command decoder 82 
via control line 93. This suspension current strength select signal on 
line 93 is ground-commanded, e.g., as two bits of a digital command word 
which permits four possible levels of suspension force, e.g., in even 
increments from 165 to 645 milliamps to produce from 4.times.10.sup.-8 g 
to 6.3.times.10.sup.-7 g force levels in the practical application 
mentioned above. The particular suspension force level selected is that 
which will enable the proof mass 39 to be effectively damped in a 
transverse or cross-axis direction, as will now be discussed. 
The transverse or cross-axis position of the proof mass 39 is monitored at 
pick-off electronics unit 80, in response to the output of the photocell 
37 (see FIG. 2). The resulting cross-axis position signal is applied to 
the suspension damping electronics unit 95 which functions to take the 
time derivative of the cross-axis position signal and thereby determine 
the rate of transverse motion of the proof mass 39. The transverse rate 
signal output from the unit 95 is applied to the suspension current 
generator 84 to effectively amplitude modulate the suspension current 
value, about its selected level, and thereby eliminate transverse 
oscillations of the proof mass 39 relative to the reference .+-.X axis. In 
the above-noted practical application of the invention, the suspension 
current is modulated about its selected value by a constant of 7.3 
ma/(mm/sec) times the transverse rate in mm/sec, up to a maximum of 1.1 
amps. 
Also shown in FIG. 9 is the closed-loop thermal control system for 
maintaining the temperature within the sensor 15 at a preselected level, 
in order to maintain constant resistivity of the proof mass materials and 
therefore also its eddy current forces. In the practical embodiment of the 
present invention referred to previously, the internal sensor temperature 
is stabilized to 70.degree. F..+-.0.5.degree. F. 
This temperature control includes the thermistor 23 mounted on the cylinder 
20 (see FIG. 3) to monitor the internal sensor temperature and temperature 
control unit 97 which is controlled by the thermistor 23 to energize the 
heater winding 22 (wound on the cylinder 20) as necessary to maintain 
constant temperature within the sensor unit, paritcularly in the vicinity 
of the proof mass 39. In order to help assure a constant temperature 
within the sensor unit 15, a suitable thermal blanket 94 is provided to 
cover the entire outer surface of the sensor unit. 
As also shown in FIG. 2, the ends of the sensor unit 15 are covered with 
suitable flat circular end plates 101 and 102 which are attached, e.g., by 
bolting them to the members 43 at each end of the sensor unit. The end 
plates 101 and 102 are provided, if need be, with electrical terminal 
members 101a and 102a respectively for permitting connection of the 
internal windings, etc. to the external control/energizing circuitry. The 
end plate 101 also includes a pair of circular apertures to accommodate 
the uncaging plunger assemblies 75 which eject when the sensor unit 15 is 
deployed by the stabilizing boom assembly, as previously described. The 
end plates 101, 102 are also covered with a suitable thermal blanket 
material in order to help assure constant temperature conditions within 
the sensor unit. 
Other modifications, adaptations and alterations to the present invention 
are of course possible in light of the above teachings. Therefore, within 
the scope of the appended claims, it should be understood that the 
invention may be practiced otherwise than as specifically described 
hereinabove.