U.S. patent number 3,730,457 [Application Number 05/056,034] was granted by the patent office on 1973-05-01 for nutation damper.
This patent grant is currently assigned to Hughes Aircraft Company. Invention is credited to Donald D. Williams, deceased.
United States Patent |
3,730,457 |
Williams, deceased |
May 1, 1973 |
NUTATION DAMPER
Abstract
A viscous nutation damper for a spin stabilized satellite said
damper comprising a sealed arcuate tube disposed near the periphery
of the spinning satellite. Said tube is oriented generally parallel
to the spin axis and is concave toward the spin axis. A viscous
fluid and spherical steel ball are sealed within the tube.
Inventors: |
Williams, deceased; Donald D.
(late of Inglewood, CA) |
Assignee: |
Hughes Aircraft Company (Culver
City, CA)
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Family
ID: |
22001727 |
Appl.
No.: |
05/056,034 |
Filed: |
June 29, 1970 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
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534251 |
Mar 14, 1966 |
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22733 |
Apr 18, 1960 |
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Current U.S.
Class: |
244/170;
74/573.1; 74/5.5 |
Current CPC
Class: |
B64G
1/281 (20130101); B64G 1/38 (20130101); Y10T
74/2122 (20150115); Y10T 74/1257 (20150115) |
Current International
Class: |
B64G
1/24 (20060101); B64G 1/38 (20060101); B64G
1/28 (20060101); B64c 017/00 (); G01c 019/04 () |
Field of
Search: |
;74/5.5,573,574
;244/155,15A,14 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Buchler; Milton
Assistant Examiner: Kunin; Stephen G.
Parent Case Text
This is a continuation of prior co-pending application Ser. No.
534,251, filed Mar. 14, 1966, now abandoned which is a continuation
application Ser. No. 22,733, filed Apr. 18, 1960 now abandoned.
Claims
What is claimed is:
1. A nutation damper for a spin-stabilized body comprising: a body
having a spin axis passing through the center of gravity thereof;
arcuate containing means disposed within said body at a distance
from said axis and arranged substantially parallel thereto; the
curvature of said containing means being concave toward said axis;
the center of said containing means being on a line normal to said
axis passing through the center of gravity of said body; means
providing viscosity disposed within said containing means; and a
movable weight disposed within said containing means.
2. A nutation damper for a spin-stabilized body comprising: a body
having a spin axis passing through the center of gravity thereof;
an arcuate tube disposed within said body at a distance from said
spin axis and arranged substantially parallel to said spin axis;
the curvature of said tube being concave toward said spin axis; the
center of said tube being on a line normal to said spin axis
passing through the center of gravity of said body, the radius of
curvature of said tube being equal to the distance of said tube
from said spin axis divided by 1.4 and multiplied by (I.sub.x
/I.sub.z - I.sub.x).sup.2 ; where I.sub.x is equal to the moment of
inertia of said body around any axis normal to said spin axis
passing through the center of gravity of said body and I.sub.z is
the moment of inertia of said body around said spin axis; a viscous
fluid sealed within said tube; and a ball disposed within said
tube.
3. A nutation damper for use in a spin-stabilized body having a
spin axis passing through the center of gravity thereof comprising:
elongated containing means disposed within said body at a distance
from said axis and arranged substantially parallel thereto; the
center of said containing means being on a line normal to said axis
and passing through the center of gravity of said body; a mass
disposed within said containing means and adapted to move along
said containing means in response to nutation of said body whereby
to damp said nutation.
4. A nutation damper for use with a spin-stabilized body having a
spin axis passing through the center of gravity thereof comprising:
arcuate containing means for attachment to said body at a distance
from said axis and substantially parallel thereto; said containing
means to be disposed with the curvature thereof being concave
toward said axis; the center of aid containing means to be disposed
on a line normal to said axis and passing through the center of
gravity of said body; means providing viscosity disposed within
said containing means; and a mass disposed within said containing
means and adapted to move along said containing means in response
to nutation of said body whereby to damp said nutation.
5. A nutation damper for use with a spin-stabilized body having a
spin axis passing through the center of gravity thereof comprising:
arcuate containing means for attachment to said body at a distance
from said axis and substantially parallel thereto; said containing
means to be disposed with the curvature thereof being concave
toward said axis; the center of said containing means to be
disposed on a line normal to said axis and passing through the
center of gravity of said body; the radius of curvature of said
containing means to be substantially equal to the distance of said
containing means from said axis divided by 1.4 and multiplied by
(I.sub.x /I.sub.z - I.sub.x).sup.2 ; where I.sub.x is equal to the
moment of inertia of said body around any axis normal to said spin
axis and I.sub.z is the moment of inertia of said body around said
spin axis; means providing viscosity disposed within said
containing means; and a mass disposed within said containing means
and adapted to move along said containing means in response to
nutation of said body whereby to damp said nutation.
6. A nutation damper for use with a spin-stabilized body having a
spin axis passing through the center of gravity thereof comprising:
elongated containing means for attachment to said body at a
distance from said axis and substantially parallel thereto; the
center of said containing means to be disposed on a line normal to
said axis and passing through the center of gravity of said body;
and a mass disposed within said containing means and adapted to
move along said containing means in response to nutation of said
body whereby to damp said nutation.
7. In spin-stabilized apparatus of the class wherein a
spin-stabilized body having a spin axis passing through the center
of gravity thereof has nutation induced in the motion of the body
through the action of pulsed jet thrust means, the combination with
said body of a nutation damper having elongated containing means
for attachment to said body at a distance from said axis and
substantially parallel thereto; the center of said containing means
to be disposed on a line normal to said axis and passing through
the center of gravity of said body; and a mass disposed within said
containing means and adapted to move along said containing means in
response to nutation of said body whereby to damp said nutation.
Description
The present invention relates to space vehicles such as satellites
and, more particularly, to a method and apparatus for controlling
the velocity, and the orientation of the spin axis, of a
spin-stabilized space vehicle.
Man-made satellites placed into orbit around the earth are often
provided with equipment requiring the satellites to be accurately
placed into specific orbits and to be oriented in a predetermined
manner. For example, a satellite for use as a radio communication
relay may have to be accurately placed into a west-to-east circular
orbit in the plane of the earth' s equator and having a period of
24 hours. Such an orbit is desirable because the satellite hovers
above a single point on the earth, inasmuch as both the satellite
and the earth have the same angular velocity. A satellite that
hovers above a single point on the earth must be accurately placed
into an orbit having a 22,750 nautical mile radius from the center
of the earth and the satellite must travel around the orbit with a
linear velocity of 10,090 feet per second.
An active communication satellite, that is, one which is equipped
to receive and retransmit radio waves, is provided with an antenna
and may be provided with solar cells as a source of power, both of
which require accurate orientation of the satellite to be useful.
To increase the antenna gain, the satellite should be spinning
about the antenna axis and this axis should be parallel to the
earth's axis. In this way, the antenna radiation pattern may be
omnidirectional about the antenna axis but have a narrow beam width
about a plane extending through the center of the satellite
perpendicular to the antenna axis. This system provides increased
antenna gain, 8 decibels, for example, in the direction of the
earth and permits the use of a radio transmitter having a
relatively low power output, thus reducing size and weight
requirements. Further, solar cells need be placed only on those
surfaces of the satellite that intercept maximum light from the
sun.
It may be found that the satellite drifts relative to a precise
stationary orbit and requires correction over a period of time. The
possible drift due to errors in the velocity of the satellite has
been determined to be 39.4 .degree. per year per foot per second of
velocity error.
To keep the cost of the satellite and the launching rocket as low
as possible, the satellite and its apparatus for orientation and
stabilization should be simple in operation, light in weight, and
small in size.
Accordingly, it is an object of the present invention to provide
apparatus for orienting a spin-stabilized vehicle, such as a
satellite, and also for controlling the velocity of the satellite
for establishing a desired orbit.
Another object of the invention is the provision of apparatus for
correcting the orbit of a spin-stabilized vehicle, such as a
satellite.
Another object of the invention is the provision of apparatus for
precessing the spin axis of a spin-stabilized body from a first
position to a second position perpendicular to the first
position.
Another object of the invention is the provision of means for
damping the nutation of a spin-stabilized body.
Yet another object of the present invention is to provide increased
antenna gain in a spinning vehicle, such as a space satellite.
A further object of the invention is the provision of optimum solar
cell illumination in an orbiting space satellite.
Still another object of the present invention is to provide
apparatus for orienting a spin-stabilized satellite that is simple
in form, reliable in operation, small in size, light in weight,
inexpensive, and low in power consumption.
In accordance with these and other objects of the invention, an
orbiting satellite having a radio antenna and solar cells is
oriented with respect to the earth and the sum to optimize the
satellite antenna gain and the solar cell illumination. The
satellite enters its orbit around the earth with its spin-axis
initially oriented perpendicular to the earth's axis. The desired
orientation of the satellite spin axis is parallel to the earth's
axis and means is provided for precessing the spin axis of the
satellite to the desired orientation by applying a reactive force
to the satellite in a proper plane. Nutation of the satellite is
damped by a viscous damper provided therein. Means is provided for
sensing the orientation of the satellite relative to the earth, the
sun, or both.
Deviations from the correct orbital period, eccentricity, and phase
are determined from observations made from the earth. Means is
provided to correct orbital deviations by applying a reactive force
to the satellite with a predetermined amount of force and in the
proper direction.
The following specification and the accompanying drawings
respectively describe and illustrate an exemplification of the
present invention. Consideration of the specification and the
drawings will provide a complete understanding of the invention,
including the novel features and objects thereof. Like reference
characters are used to designate like parts through the figures of
the drawings.
FIG. 1 is an elevation of an exemplary embodiment of a satellite
launching vehicle in accordance with the present invention showing
the relationship of the propulsion rockets to the satellite;
FIG. 2 is an enlarged elevation of a portion of the launching
vehicle of FIG. 1 showing the relationship of fourth, fifth, and
sixth stage rockets to the satellite and enclosing heat shield;
FIG. 3 is a plan of the satellite showing an antenna provided
thereon in a folded position;
FIG. 4 is a perspective of the satellite illustrating the antenna
in its erected position and the solar sensing arrangement;
FIG. 5 is a sectional elevation of a portion of the satellite
illustrating the arrangement of an attitude and velocity control
system including compressed gas tanks, valves and nozzles and
showing a nutation damping arrangement;
FIG. 6 is a diagram of the satellite radio and electrical
system;
FIG. 7 is a diagram of a satellite trajectory from the earth to a
predetermined orbit;
FIG. 8 is a diagram illustrating the initial relationship of the
spin axis of the satellite to the earth's axis and the application
of a reactive force producing a precessional torque;
FIG. 9 is a diagram illustrating the change in satellite
orientation resulting from the reactive force depicted as being
applied in FIG. 8;
FIG. 10 is a diagram illustrating further reactive force being
applied to the satellite; and
FIG. 11 is a diagram depicting the satellite after orientation by
precession has aligned the spin axis of the satellite with the
earth's axis.
Although the present invention does not embrace a vehicle for
conveying a satellite to an orbit, a brief description of a
representative rocket vehicle useful for this purpose is included.
Reference is hereby made to the book "Rocket Encyclopedia
Illustrated," edited by J. W. Herrick and E. Burgess, Aero
Publishers, Inc., Los Angeles, California, 1959, and to the
bibliography therein for details of rocket vehicles and definitions
of terms.
There is shown in FIG. 1 an exemplary embodiment of a rocket
vehicle, indicated generally by the numeral 25, for accurately
launching a space vehicle, such as a satellite 26, into a
predetermined orbit around the earth. In the present example, the
orbit into which the satellite 26 is to be placed is a so-called
stationary or synchronous orbit which is a circular west-to-east
equatorial orbit having a period of 24 hours. This orbit has a
radius of 22,750 nautical miles from the center of the earth and
the satellite 26 travels at a velocity of 10,090 feet per second
(fps) in the direction of the earth's rotation and appears to stand
still or hover over one point on the earth's surface. The satellite
26 may then be used, for example, to relay radio communications
over long distances.
The primary or booster portion of the rocket vehicle 25 is a
multistage rocket power plant designated by the numeral 27 in FIG.
1 and comprising first, second, third, and fourth stage rockets 31,
32, 33, and 34, respectively, arranged in tandem. The first,
second, and third stage rockets 31, 32, and 33 are provided with
guidance elements, such as jet vanes (not shown) which are disposed
in the jet stream from the rocket exhaust nozzles to deflect the
jet and thus obtain a turning force to control the direction and
attitude of the vehicle 25. Such an arrangement is shown on page
248 of the above-referenced Rocket Encyclopedia. The fourth stage
rocket 34 is provided with means, such as spin nozzles or spin
rockets (not shown) to rotate the fourth stage rocket 34 about its
longitudinal axis for stabilization and such arrangements are shown
and described on pages 456 and 457 of the Rocket Encyclopedia. The
ignition of the rocket engines, jettisoning of burned out rockets,
guidance, and spin stabilization of the multistage rocket power
plant 27 is automatically controlled by a guidance unit 28, which
may be one of several types of programming and control systems that
are well known in the art of rocketry.
If desired, the multistage rocket power plant 27 may be, for
example, the "Scout" rocket developed for the National Aeronautics
and Space Agency by Vought Astronautics Division of Chance Vought
Aircraft, Inc., Dallas, Texas. Reference is made to the publication
entitled "Space Research Vehicle Systems Developed from NASA
SCOUT," publication number E9R-12402, dated August 1959 and
published by Chance Vought Aircraft, Inc., for details of the Scout
rocket. In particular, on page 23 of the Chance Vought publication,
there is a description of a guidance system developed by the
Minneapolis-Honeywell Regulator Co. which may be used as the
guidance unit 28 in the present rocket vehicle 25.
The multistage rocket power plant 27 propels a gross payload 29 to
a point near the perigee of the transfer ellipse; that is, near the
lowest point of the elliptical trajectory from the outer atmosphere
of the earth to the desired orbit. The gross payload 29, best seen
in FIGS. 1 and 2, comprises a fifth stage rocket 35, a sixth stage
rocket 36 and the satellite 26. The gross payload 29 and the fourth
stage rocket 34 are covered by a nose shell or cylindrical heat
shield 37 until the launching vehicle 25 leaves the earth's
atmosphere. The heat shield 37 is then automatically separated from
the gross payload 29 and the fourth stage rocket 34. The
construction and method of separation of a typical nose shell is
illustrated on page 6 of the Chance Vought publication.
The fifth stage solid-propellant rocket 35, which is secured in
tandem to the fourth stage rocket 34, provides the additional
thrust required for the satellite 26 to reach the perigee of the
transfer ellipse. The fifth stage rocket 35 is provided with means,
such as an electrically fused annular explosive charge 38 (FIG. 2)
which encircles the case of the fifth stage rocket 35 for reducing
the thrust to zero. When detonated, the explosive charge 38
ruptures the combustion chamber of the fifth stage rocket 35.
The sixth stage rocket 36 provides the additional thrust to
establish the satellite 26 in the circular orbit from the apogee or
highest point of the transfer ellipse. The sixth stage rocket 36 is
secured to the satellite 26 on the side opposite the side to which
the fifth stage rocket 35 is secured, and is oriented in the
opposite direction. That is, the sixth stage rocket 36 is oriented
so that the rocket exhaust nozzle extends out from the satellite 26
in the opposite direction from that in which the exhaust nozzle of
the fifth stage rocket 35 extends out from the satellite 26. As
will be fully apparent hereinafter, this is necessary so that the
thrust of the sixth stage rocket 36 will be applied in the correct
direction.
As may be seen in FIGS. 3, 4 and 5, the satellite 26 is of a
cylindrical configuration and may have a diameter of 29 inches, a
thickness of 121/2 inches, and a weight on the order of 25 pounds,
for example. An antenna 50 is secured to one of the plane faces of
the satellite 26. After the satellite 26 has reached its
predetermined orbit, the antenna 50 extends outwardly from the
satellite along an axis extending through the center thereof
coaxial with the longitudinal axis of the rocket vehicle 25. This
is the spin axis and the antenna axis of the satellite 26. The
antenna 50 provides a radiation pattern that is substantially
omnidirectional about the antenna axis, but has a narrow beam width
about a plane extending through the center of the satellite 26
perpendicular to the axis of the antenna 50. The cylindrical
surface of the satellite 26 parallel to the axis of the antenna 50
is provided with solar cells, indicated by the numeral 52, for
converting sunlight into electrical energy.
As shown in FIG. 2, the fifth and sixth stage rockets 35 and 36 are
secured to the plane faces of the satellite 26 and are adapted to
be separated from the satellite when desired. The antenna 50, which
is a dielectric-loaded coaxial transmission line having
circumferential slots spaced a half wavelength apart, is pivoted at
the surface of the satellite 26. Thus, initially the antenna 50 is
folded against the surface of the satellite 26 beneath the fifth
stage rocket 35 (see FIGS. 2 and 3). The antenna 50 is spring
loaded at the pivot point for erection thereof after separation of
the fifth stage rocket 35 from the satellite 26 (FIG. 4).
The satellite 26 is provided with a source of reactive power such
as two tanks 53 of gas or condensed liquid and its associated vapor
under pressure, for example, compressed nitrogen gas at a pressure
of 3,000 pounds per square inch. In the present embodiment, the
tanks 53 are of a toroidal configuration and are disposed
immediately within the cylindrical surface of the satellite 26. A
velocity control valve 54 and an attitude control valve 55 are
connected to the tanks 53 (best seen in FIG. 5). The valves 54 and
55 are quick acting, low-leakage solenoid-controlled valves such as
the type AF56C-37A manufactured by the Eckel Valve Company of San
Fernando, California. The velocity control valve 54 is connected to
a nozzle 56 disposed at the periphery of the satellite 26 and
oriented to provide a jet of nitrogen gas directed radially outward
from the center of the satellite 26 along a line perpendicular to
the spin axis passing through the center of gravity of the
satellite 26. The attitude control valve 55 is connected to a
nozzle 57 disposed near the periphery of the satellite 26 and
oriented to provide a jet of nitrogen gas directed outward from the
satellite 26 along a line parallel to the spin axis of the
satellite 26.
Means is provided for absorbing nutation energy such as a nutation
damper 58 comprising a sealed arcuate tube 60 which may be four
inches in length disposed near the periphery of the satellite 26
(FIG. 5) and oriented generally parallel to the spin axis, the
curvature of the tube 60 being concave toward the spin axis. The
tube 60 is filled with a viscous fluid 61 such as oil or silicone
fluid and contains a spherical steel ball 62, 1/4 inch in diameter,
for example. The radius of curvature of the tube 60 may be made
equal to its distance from the spin axis divided by 1.4 and
multiplied by
(I.sub.x /I.sub.z - I.sub.x).sup.2
where I.sub.x is the moment of inertia of the satellite 26 around
any axis normal to the spin axis passing through the center of the
satellite 26 and I.sub.z is the moment of inertia of the satellite
26 about the spin axis. The factor of 1.4 is necessary because the
ball 62 rolls rather than slides in the tube 60. This choice of
radius makes the damper resonant at the correct frequency for any
spin speed of the vehicle.
The satellite 26 is also provided with means for sensing its
orientation with respect to the sun. In FIG. 4 there is shown a
slit 70 provided in the cylindrical surface of the satellite 26
providing a fan-shaped angular field of view that extends parallel
to the antenna axis. A single orientation sensing solar cell 71 is
disposed within the satellite 26 adjacent the slit 70 so that when
the sun is within the field of view, which may be 140.degree., for
example, the solar cell 71 is energized to develop a potential at
its output terminals. A second slit 72 is provided on the same face
of the satellite 26 and provides a second fan-shaped angular field
of view that intersects the field of view of the first slit 70 at
an angle which may be 35.degree., for example. Similarly, a second
orientation sensing solar cell 73 disposed within the satellite 26
develops a potential when the sun is in the field of view of the
second slit 72.
As may be seen in FIG. 6, the satellite 26 is provided with a radio
transmitter and receiver 90 that is electrically connected to the
antenna 50 as indicated. A source of potential, such as a storage
battery 91, applies electrical power to the radio transmitter and
receiver 90. The solar cells 52 disposed on the outer surface of
the satellite 26 are connected through rectifiers or charging
diodes 92 to the battery 91 to maintain it in a charged condition.
Approximately 2200 solar cells 52 may be provided and
interconnected in banks. The cells 52 in each bank are connected in
a series-parallel arrangement, and although there may be different
numbers of cells 52 in each bank, the number of cells in series in
each series-parallel arrangement is identical to provide the proper
voltage for the battery 91. The charging diodes 92 are
nonconductive during periods that the voltage developed by any bank
of cells 52 drops below that of the battery 91.
A radio control circuit 93 is also connected to the radio
transmitter and receiver 90 and to the battery 91. The radio
control circuit 93 is responsive to control signals received by the
radio transmitter and receiver 90 for applying a potential from the
battery 91 to the various electrically controlled devices
associated with the satellite 26 and with the fifth and sixth stage
rockets 35 and 36. The particular radio remote control system
utilized may be one of several systems well known in the art, for
example, one utilizing subcarrier signals transmitted on a carrier
wave. The radio control circuit 93 may, for example, include a
number of filters for separating the various subcarrier signals and
actuating relays in response thereto, as is well known in the
art.
The explosive charge 38 for rupturing the combustion chamber of the
fifth stage rocket 35 is connected to the output of the radio
control circuit 93. The igniter for the sixth stage rocket 36 is
also connected to the output of the radio control circuit 93. The
velocity control valve 54 and the attitude control valve 55 are
each individually connected to the output of the radio control
circuit 93.
The first and second orientation sensing solar cells 71 and 73 are
respectively connected to first and second orientation signal
oscillators 95 and 96. The orientation sensing solar cells 71 and
73 supply electrical power for the oscillators 95 and 96 so that
orientation signals are developed when the solar cells 71 and 73
are illuminated by the sun. The output signals from the oscillators
95 and 96 are applied to the radio transmitter and receiver 90 for
transmission to a satellite control point (not shown). Thus, if the
sun passes through the field of view of the second slit 72, a
signal will be developed by the second oscillator 96; and if the
sun passes through the field of view of the first slit 70, a signal
will be developed by the first oscillator 95.
In the construction of the satellite 26, the mass of the units
associated with the satellite 26 is accurately determined and the
equipment is distributed within the satellite 26 so that the center
of gravity is made to coincide with the center of the satellite 26;
and the axis of maximum moment of inertia is made to coincide with
the antenna axis.
Referring now to FIG. 7, the earth is represented by the circle
designated 100 and is rotating in the counterclockwise direction,
indicated by the arrow 101, around an axis 102 represented as going
into the drawing through the north pole. The rocket vehicle 25 is
fired from a point 103 on the equator of the earth 100, which point
may be, for example, Jarvis Island in the Pacific Ocean that is
located at 23 minutes south latitude and 160 west longitude. Prior
to firing, the battery 91 is completely charged, the nitrogen tanks
53 are pressurized, and the radio transmitter and receiver 90 and
the radio control circuit 93 are placed in operation. After firing,
the first four rockets 31-34 of the propulsion system 27 are
automatically fired in sequence and guided by the guidance unit 28.
After burnout, each empty rocket case is jettisoned, as is well
known in the art.
After the rocket vehicle 25 has attained considerable altitude, the
rocket vehicle is automatically turned in an easterly direction to
coincide with the direction of rotation of the earth 100, by the
guidance unit 28. Prior to ignition of the fourth stage rocket 34,
and separation of this rocket from the third stage rocket, the
fourth stage rocket 34 has a spin about its longitudinal axis
imparted to it at a rate of 2.7 revolutions per second (rps) to
provide spin stabilization. The multistage rocket power plant 27
propels the gross payload 29 to a point above the earth 100 near
the lowest point or perigee 104 of an elliptical trajectory or
transfer ellipse 105.
The fifth stage rocket 35 provides the additional thrust necessary
to cause the satellite 26 to reach the perigee 104 and to traverse
the transfer ellipse 105. Because the velocity of the satellite 26
at the perigee 104 is quite critical in order to achieve the
correct apogee, the velocity is determined from the earth 100 in a
well known manner, such as by means of a radio interferometer, a
tracking antenna, or by doppler frequency shift measurement
techniques, for example. These measurements may make use of the
radio transmitter and receiver 90 in the satellite 26 as a radio
repeater.
When the correct velocity has been attained, a radio control signal
is transmitted from the earth 100 to the satellite 26 to fire the
explosive charge 38 and rupture the combustion chamber of the fifth
stage rocket 35 to reduce the thrust provided thereby to zero. The
fifth stage rocket 35 is then separated from the satellite 26.
A first arrow 106 (FIG. 7) indicates the orientation of the sixth
stage rocket 36 and the satellite 26 at the perigee 104 of the
transfer ellipse 105 with the arrow pointing in the direction of
thrust of the fifth stage rocket 35. Inasmuch as the satellite 26
is spin-stabilized about the rocket axis by the spin imparted by
the spin nozzles or spin rockets (not shown) on the fourth stage
rocket 34, the satellite 26 and the sixth stage rocket 36 maintains
its attitude in space (arrows 107, 108, and 110) as it traverses
the transfer ellipse 105 to the other side of the earth 100. As
described hereinbefore the sixth stage rocket 36 is secured to the
satellite 26 with an orientation such that it applies thrust in the
direction opposite to that of the thrust of the fifth stage rocket
35.
At the highest point or apogee 109 of the transfer ellipse 105, the
satellite 26 has attained the altitude of the desired circular
24-hour orbit 112, but is traveling at only 5200 fps, which is less
than that required for establishment of the satellite 26 in the
orbit 112. A radio control signal is transmitted to the satellite
26 to cause ignition of the sixth stage rocket 36 to provide the
additional velocity of 4890 fps to establish the satellite 26 into
the circular orbit 112. At the apogee 109 on the other side of the
earth 100 from the firing point 103, the direction of thrust of the
sixth stage rocket 36 is such as to cause the satellite 26 to enter
the circular orbit 112 due to the fact that the sixth stage rocket
36 has maintained its attitude in space while traversing the
transfer ellipse 105 and due to the fact that the orientation of
the sixth stage rocket 36 is opposite to that of the fifth stage
rocket 35.
As the satellite 26 traverses its orbit 112, it is spinning about
its spin axis with an angular velocity of 2.7 rps imparted to it by
the fourth stage rocket 34. However, the spin axis of the satellite
26 is perpendicular to the earth's axis 102, and thus the antenna
50 does not radiate efficiently toward the earth 100. Accordingly,
the satellite 26 is reoriented by precessing its spin axis through
90.degree.. Precession is accomplished by the reactive force
produced by a jet of nitrogen gas from the attitude control nozzle
57 which applies thrust parallel to the spin axis near the
periphery of the satellite 26. The jet of nitrogen gas is
controlled by the attitude control valve 55 to produce a net torque
around an axis perpendicular to the spin axis of the satellite 26.
By periodically pulsing the jet to be on during only a
predetermined portion of the spin cycle of the satellite 26, the
torque is applied in the correct plane to precess the spin axis
through 90.degree. until it is parallel to the earth's axis 102.
The attitude control valve 55 may be actuated during only
approximately 60.degree. of the spin cycle of the satellite 26, for
example.
The attitude control valve 55 is pulsed by radio control from the
earth 100. The correct phase of the spin cycle to actuate the valve
55 is determined from the earth 100 by means of the first
orientation sensing solar cell 71 adjacent the slit 70 in the
satellite 26 and its associated oscillator 95 which modulates the
radio transmission from the satellite 26. By correcting for the
two-way propagation delay, the jet is turned on during the correct
portion of the spin angle of the satellite 26 to cause precession
in the proper direction. This action is indicated in FIGS.
8-11.
The amount of precession of the antenna axis of the satellite 26 is
determined from the earth 100 by means of the orientation signal
developed by the second orientation oscillator 96 associated with
the slit 72 and orientation sensing solar cell 73. This
determination can be made only during the time of day that the
satellite 26 is in sunlight. However, in the equatorial orbit of
the present example, only rarely does the earth 100 come between
the sun and the satellite 26, and then only for intervals of short
duration. As the satellite 26 spins about its spin axis, the slit
72 periodically passes through sunlight. Thus a periodic
orientation signal is developed and transmitted to the earth 100
and the orientation of the antenna axis with respect to the sun may
be determined. This information tells when the necessary precession
has been completed.
In as much as the satellite 26 may have velocity and altitude
errors, it traverses only an approximate 24-hour stationary or
synchronous orbit 112, and corrections are made by radio control of
the velocity control valve 54. The satellite 26 is tracked from the
earth 100 by means of radio signals transmitted to the satellite 26
and relayed back to the earth 100 to determine the drift of the
satellite 26 relative to the earth 100. The velocity of the
satellite 26 is increased or decreased by opening the velocity
control valve 54 for controlled time intervals during the proper
portion of the spin cycle by means of radio control signals.
Opening of the velocity control valve 54 results in jets of
compressed nitrogen issuing from the associated nozzle 56 to
provide a reactive force that changes the velocity of the satellite
26. This will correct the orbital period and will also reduce or
eliminate the eccentricity of the orbit. Inclination of the orbit
may be corrected by timing the attitude control valve 55 and nozzle
57 so that "on" periods of the jet produce no net precession. For
example, the valve 55 may be opened for one complete revolution of
the satellite 26 about the spin axis, which produces no net torque
about a single axis and hence no net precession.
As stated hereinbefore, the antenna axis of the satellite 26 is the
axis of the maximum moment of inertia. This provides stability
against the effects of vibration and associated energy loss that
would otherwise tend to orient the spin around the axis of the
largest moment of inertia, if it were other than the antenna axis.
In this way, the effects of such vibration are to cause the spin to
stabilize about the desired axis, that is, to damp the nutation. In
addition to the natural tendency of the satellite 26 to damp
nutation, the nutation damper 58 further reduces any nutational
motion by absorbing the nutation energy. The nutation damper 58 is
resonant at the correct frequency regardless of spin if the radius
of curvature of the tube 60 is chosen as indicated hereinbefore.
When precession is accomplished in a constant direction by a series
of pulses, the nutation resulting from a single pulse does not
build up so that only a small damper 58 is required.
Thus, the satellite 26 is made to orbit around the earth 100 with
the same angular velocity as the earth 100 and in the same
direction of rotation. The antenna axis of the satellite 26 is
parallel to the axis of rotation 102 of the earth 100 so that the
antenna 50 radiates signals to the earth 100 in a narrow beam and
the solar cells 52 are oriented to receive optimum light from the
sun.
As to the disclosed means for sensing the orientation of the
satellite 26, namely, the slits 70 and 72, orientation sensing
solar cells 71 and 73, and the orientation oscillators 95 and 96,
it will be understood that other means may be provided. For
example, asymmetry may be deliberately introduced into the antenna
radiation pattern of the satellite 26.
It will be obvious that more than one jet nozzle may be provided
for velocity control or for attitude control. Several jets of gas
may be produced at different locations on a spinning vehicle, each
jet being pulsed and timed or synchronized as described
hereinbefore with reference to a single jet. Further, a jet or jets
may be produced on a spinning body to achieve simultaneous change
in velocity and attitude so as to realize, for example, a
proportional navigation course. By this means a target seeking
vehicle may be constructed in accordance with the invention.
It will also be apparent that valves may be pulsed or modulated in
accordance with desired functions of time, such as portions of sine
waves, for example, to provide smoother control. In addition, jet
control systems in accordance with the invention may be used for
other types of satellites than communication satellites. For
example, meterological, astronomical or navigational satellites may
also be spin-stabilized and controlled by pulsed jets of gas to
change velocity or attitude of the spin axis in space, or both.
Furthermore, the principles embodied in the present invention may
also be applied to vehicles for probing into space and for
intercepting other space vehicles.
Thus, there has been described a method and apparatus for launching
a satellite into a particular orbit, and with a predetermined
orientation with respect to the earth, to provide optimum antenna
gain and optimum solar cell illumination. By using simple
spin-stabilization to orient the satellite, the weight and
complexity of the satellite have been minimized. A method and
apparatus has been described for precessing the spin axis of a
spin-stabilized body from a first position to a second position
perpendicular to the first position, and for damping of nutation.
Further, a method and apparatus has been described for correcting
the orbit of a spin-stabilized satellite.
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