U.S. patent number 3,711,046 [Application Number 05/870,077] was granted by the patent office on 1973-01-16 for automatic missile guidance system.
Invention is credited to Hamilton Barhydt, Spencer D. Howe.
United States Patent |
3,711,046 |
Barhydt , et al. |
January 16, 1973 |
AUTOMATIC MISSILE GUIDANCE SYSTEM
Abstract
The automatic missile guidance system comprises a sight with
which a gunner establishes a line-of-sight from the gun position to
the target. When the missile is launched, a source of pulsating,
radiant energy on the rear of the missile is detected by a guidance
unit at the sight. The guidance unit produces steering commands
related to the deviation of the missile from the line-of-sight.
Means interconnecting the guidance unit and the missile transmits
the guidance signals to the missile to direct it along the
line-of-sight. This guidance unit to missile connection may be
wires which unreel from the missile as it proceeds towards its
target.
Inventors: |
Barhydt; Hamilton (Playa del
Rey, CA), Howe; Spencer D. (Los Angeles, CA) |
Family
ID: |
25354753 |
Appl.
No.: |
05/870,077 |
Filed: |
October 22, 1969 |
Current U.S.
Class: |
244/3.12;
244/3.16 |
Current CPC
Class: |
F41G
7/30 (20130101); F42B 15/04 (20130101) |
Current International
Class: |
F41G
7/20 (20060101); F42B 15/00 (20060101); F42B
15/04 (20060101); F41G 7/30 (20060101); F42b
015/04 (); F41g 007/02 (); F41g 007/14 () |
Field of
Search: |
;244/3.12,3.16 |
Primary Examiner: Borchelt; Benjamin A.
Assistant Examiner: Webb; Thomas H.
Claims
What is claimed is:
1. An automatic missile guidance system for guiding a missile along
a line-of-sight to a target established by a human operator, said
missile having guidance means comprising:
a guidance unit for developing steering signals related to the
deviation of said missile from said line-of-sight;
means for transmitting said guidance signals from said guidance
unit to said missile; and
a source of pulsating radiant energy on said missile and directed
to the rear thereof, said source of radiant energy having a
predetermined pulsation frequency, said guidance unit being
selectively responsive to said radiant energy at said predetermined
pulsation frequency.
2. The automatic missile guidance system of claim 1 wherein said
guidance unit has means for narrowing its field of view when said
missile is proximate to said line-of-sight.
3. The automatic missile guidance system of claim 2 wherein said
guidance unit has a radiant energy receiver thereon adapted to
receive radiant energy from said source of pulsating radiant energy
on said missile, said radiant energy receiving means having a large
radiant energy detector disposed therein to receive radiant energy
in any part of the field of view and a small radiant energy
detector disposed therein, said small radiant energy detector being
positioned to receive radiant energy in the central portion of the
field of view of said radiant energy receiver, and switch means for
switching between said large and said small radiant energy
detectors in response to position of missile source image in field
of view.
4. The automatic missile guidance system of claim 3 wherein said
small radiant energy detector means has an output to said switch
means to switch said switch means so that its small radiant energy
detector means is switched to control said missile when the input
radiance to said small radiant energy detection means exceeds a
predetermined value.
5. The automatic missile guidance system of claim 4 wherein said
source of pulsating radiant energy on said missile operates at a
substantially constant frequency.
6. The automatic missile guidance system of claim 5 wherein said
radiant energy receiving means on said guidance unit comprises a
telescope having a wide field of view about an optical axis, a
reticle disposed at the focus of said telescope concentric with and
transverse to the axis thereof, means rotating said reticle about
the axis of said telescope at a predetermined speed, said reticle
having an opaque modulation pattern thereon to provide a
substantially sinusoidally varying average transmissivity gradient
circularly around said reticle and to provide a rapidly increasing
depth-of-modulation gradient outward from the center of said
reticle, said large radiant energy detector being disposed to
receive radiant energy passed by said reticle, said small radiant
energy detector being disposed in a position to intercept radiant
energy passing through the central portion of said reticle.
7. The automatic missile guidance system of claim 3 wherein the
output of said switch is connected to first and second phase
detectors, for respectively detecting pitch and yaw of said
missile, said phase detectors being connected to said missile to
guide said missile.
8. The automatic missile guidance system of claim 7 wherein said
means for transmitting said guidance signals from said guidance
unit to said missile comprises at least one wire interconnected
between said guidance unit and said missile.
9. An automatic missile guidance system for guiding a missile along
a line-of-sight to a target established by a human operator, said
missile having guidance means comprising:
a guidance unit for developing steering signals related to the
deviation of said missile from said line-of-sight;
means for transmitting said guidance signals from said guidance
unit to said missile; and
a source of pulsating radiant energy on said missile and directed
to the rear thereof, said source of radiant energy having a
predetermined substantially constant pulsation frequency, said
guidance unit being selectively responsive to said radiant energy
at said predetermined substantially constant pulsation frequency,
said guidance unit having means for narrowing its field of view
when said missile is proximate to said line-of-sight.
10. A system for guiding a self-propelled missile having steering
means comprising:
means disposed on said missile for emitting radiant energy
pulsating at a predetermined frequency rearwardly thereof;
means for intercepting radiant energy having a wide field of
view;
means disposed within said interception means for amplitude
modulating intercepted radiant energy at a predetermined
frequency;
a large radiant energy detector disposed to receive radiant energy
intercepted by said interception means and modulated by said
modulation means;
a small radiant energy detector disposed within said interception
means in a position to receive radiant energy in the central
portion of the field of view of said interception means and
modulated by said modulation means;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing to its output only signals
appearing at the output of said large radiant energy detector, said
switch means being responsive to a switching signal applied at a
control input for passing to its output only signals appearing at
the output of said small radiant energy detector;
first frequency selective means having its input electrically
coupled to said small radiant energy detector and having a narrow
frequency passband centered around said predetermined frequency,
said switch means having its control input electrically coupled to
the output of said first frequency selective means;
second frequency selective means having its input coupled to the
output of said switch means and having a narrow frequency passband
centered around said predetermined frequency;
an amplitude modulation demodulator having its input coupled to the
output of said second frequency selective means;
phase detecting means having an input coupled to the output of said
demodulator;
reference means associated with said modulation means for
developing periodic reference signals, the phase of which is
related to the phase of said modulation means, said reference
signals being coupled to the reference input of said phase
detecting means; and
signal transmission means having its input coupled to the outputs
of said phase detecting means and its output coupled to said
missile.
11. A guided missile system comprising:
a self-propelled missile having steering means;
a source of infrared energy pulsating at a substantially constant
predetermined frequency disposed on said missile and directed
rearwardly thereof;
a visual telescope having an optical axis;
an infrared telescope having an optical axis substantially aligned
with the optical axis of said visual telescope and having a wide
field of view;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a predetermined speed, said reticle being transparent
to said pulsating infrared energy and having an opaque coating
arranged in a modulation pattern thereon, said coating being
disposed in a sinusoidally varying concentration circularly around
said reticle to provide a sinusoidal average transmissivity
gradient circularly around said reticle, said coating being
disposed in a varying radial concentration increasing outwardly to
provide a rapidly increasing depth-of-modulation gradient near the
center of said reticle;
a condensing lens disposed adjacent said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said reticle
and said condensing lens;
a light pipe disposed in the center of said condensing lens to
intercept radiant energy passing through the central portion of
said reticle;
a small radiant energy detector disposed in said light pipe in a
position to receive radiant energy intercepted thereby;
an electromagnetic relay having a normally open contact;
a normally closed contact;
a switch arm and an actuating coil, said small radiant energy
detector being electrically coupled to the normally open contact of
said relay, said large radiant energy detector being electrically
coupled to the normally closed contact of said relay;
a first narrowband amplifier having its input electrically coupled
to said small radiant energy detector and having a narrow passband
centered around said predetermined frequency;
a rectifier and filter connected to the output of said first
narrowband amplifier, said filter having a predetermined charge and
discharge time constant;
a switch amplifier having its input electrically coupled to the
output of said rectifier and filter and having its output connected
to the actuating coil of said relay, said switch amplifier being
biased to have a predetermined input threshold level above which
said relay is energized and below which said relay is
de-energized;
a source of voltage that exceeds said threshold a predetermined
time after the launching of said missile electrically coupled to
the input of said switch amplifier;
a second narrowband amplifier having its input coupled to the
switch arm of said relay and having a narrow passband centered
around said predetermined frequency;
an automatic gain control circuit coupled to said second narrowband
amplifier for maintaining signals at the output thereof at a
constant amplitude;
an amplitude modulation demodulator having its input coupled to the
output of said second narrowband amplifier;
a variable gain amplifier having its input coupled to the output of
said demodulator, said variable gain amplifier having a gain that
varies as a function of an applied control voltage;
a gain control voltage source coupled to said variable gain
amplifier and applying a control voltage thereto that gradually
increases the gain of said variable gain amplifier from the time
said missile is launched;
a pair of phase detectors, each having signal inputs coupled to the
output of said variable gain amplifier;
a pair of coil and magnet assemblies associated with said reticle
for developing periodic reference signals having phases related to
the instantaneous angular position of said reticle, said reference
signals being in phase quadrature with each other;
a pair of squaring circuits, each having its input individually
coupled to the output of a different one of said coil and magnet
assemblies, the output of each of said squaring circuits being
individually coupled to the reference input of a different one of
said phase detectors;
a pair of compensation networks, each having its input coupled to
the output of a different one of said phase detectors; and
a signal transmission circuit having its input coupled to the
outputs of said compensation networks and its output coupled to
said missile.
12. A system for guiding a self-propelled missile having steering
means comprising:
a source of infrared energy pulsating at a predetermined frequency
disposed on said missile and directed rearwardly thereof;
an infrared telescope having a wide field of view about an optical
axis;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a predetermined speed, said reticle being transparent
to said pulsating infrared energy and having an opaque coating
arranged in a modulation pattern thereon, said coating being
disposed in a sinusoidally varying concentration circularly around
said reticle to provide a sinusoidal average transmissivity
gradient circularly around said reticle, said coating being
disposed in a varying radial concentration increasing outwardly to
provide a rapidly increasing depth-of-modulation gradient near the
center of said reticle;
a condensing lens disposed adjacent said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said reticle
and said condensing lens;
a light pipe disposed in the center of said condensing lens to
intercept radiant energy passing through the central portion of
said reticle;
a small radiant energy detector disposed in said light pipe in a
position to receive radiant energy intercepted thereby;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing only signals appearing at the
output of said large radiant energy detector, said switch means
being responsive to an applied switching signal for passing only
signals appearing at the output of said small radiant energy
detector;
a first narrowband amplifier having its input electrically coupled
to said small radiant energy detector and having a narrow passband
centered around said predetermined frequency;
a rectifier and filter connected to the output of said first
narrowband amplifier, said filter having a predetermined charge and
discharge time constant;
a switch amplifier having its input electrically coupled to the
output of said rectifier and filter and having its output connected
to the control input of said switch means, said switch amplifier
being biased to have a predetermined input threshold level above
which said switch means is responsive and below which said switch
means is nonresponsive;
a source of voltage that exceeds said threshold a predetermined
time after the launching of said missile electrically coupled to
the input of said switch amplifier;
a second narrowband amplifier having its input coupled to the
output of said switch means and having a narrow passband centered
around said predetermined frequency;
an automatic gain control circuit coupled to said second narrowband
amplifier for maintaining signals at the output thereof at a
constant amplitude;
an amplitude modulation demodulator having its input coupled to the
output of said second narrowband amplifier;
a variable gain amplifier having its input coupled to the output of
said demodulator, said variable gain amplifier having a gain that
varies as a function of an applied control voltage;
a gain control voltage source coupled to said variable gain
amplifier and applying a control voltage thereto that gradually
increases the gain of said variable gain amplifier from the time
said missile is launched;
a pair of phase detectors, each having signal inputs coupled to the
output of said variable gain amplifier;
reference means associated with said reticle for developing a pair
of periodic reference signals having phases related to the
instantaneous angular position of said reticle, said reference
signals being in phase quadrature with each other, each of said
reference signals developed by said reference means being
individually coupled to the reference input of a different one of
said phase detectors;
a pair of compensation networks, each having its input coupled to
the output of a different one of said phase detectors; and
a signal transmission circuit having its input coupled to the
outputs of said compensation networks and its output coupled to
said missile.
13. A system for guiding a self-propelled missile having steering
means comprising:
a source of infrared energy pulsating at a substantially constant
predetermined frequency disposed on said missile and directed
rearwardly thereof;
an infrared telescope having a wide field of view about an optical
axis;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a substantially constant predetermined speed, said
reticle being transparent to said pulsating infrared energy and
having an opaque coating arranged in a modulation pattern thereon,
said coating being disposed in a sinusoidally varying concentration
circularly around said reticle to provide a sinusoidal average
transmissivity gradient circularly around said reticle, said
coating being disposed in a varying radial concentration increasing
outwardly to provide a rapidly increasing depth-of-modulation
gradient near the center of said reticle;
a condensing lens disposed adjacent said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said reticle
and said condensing lens;
a light pipe disposed in the center of said condensing lens to
intercept radiant energy passing through the central portion of
said reticle;
a small radiant energy detector disposed in said light pipe in a
position to receive radiant energy intercepted thereby;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing to its output only signals
appearing at the output of said large radiant energy detector, said
switch means being responsive to a switching signal applied at a
control input for passing only signals appearing at the output of
said small radiant energy detector;
a first narrowband amplifier having its input electrically coupled
to said small radiant energy detector and having a narrow passband
centered around said predetermined frequency;
a rectifier and filter connected to the output of said first
narrowband amplifier, said filter having a predetermined charge and
discharge time constant, said switch means having its control input
electrically coupled to the output of said rectifier and
filter;
a source of voltage that exceeds said threshold a predetermined
time after the launching of said missile electrically coupled to
the control input of said switch means;
a second narrowband amplifier having its input coupled to the
output of said switch means and having a narrow passband centered
around said predetermined frequency and having an automatic gain
control for maintaining signals at the output thereof at a
substantially constant amplitude;
an amplitude modulation demodulator having its input coupled to the
output of said second narrowband amplifier;
a variable gain amplifier having its input coupled to the output of
said demodulator, said variable gain amplifier having a gain that
gradually increases from the time said missile is launched;
a pair of phase detectors, each having signal inputs coupled to the
output of said variable gain amplifier;
reference means associated with said reticle for developing a pair
of periodic reference signals having phases related to the
instantaneous angular position of said reticle, said reference
signals being in phase quadrature with each other, each of said
reference signals developed by said reference means being
individually coupled to the reference input of a different one of
said phase detectors; and
a signal transmission circuit having its input coupled to the
outputs of said phase detectors and its output coupled to said
missile.
14. A system for guiding a self-propelled missile having steering
means comprising:
a source of infrared energy pulsating at a substantially constant
predetermined frequency disposed on said missile and directed
rearwardly thereof;
an infrared telescope having a wide field of view about an optical
axis;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a substantially constant predetermined speed, said
reticle having an opaque modulation pattern thereon to provide a
sinusoidal average transmissivity gradient circularly around said
reticle and to provide a rapidly increasing depth-of-modulation
gradient near the center of said reticle;
a condensing lens disposed adjacent said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said reticle
and said condensing lens;
a light pipe disposed in the center of said condensing lens to
intercept radiant energy passing through the central portion of
said reticle;
a small radiant energy detector disposed in said light pipe in a
position to receive radiant energy intercepted thereby;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing to its output only signals
appearing at the output of said large radiant energy detector, said
switch means being responsive to a switching signal applied at a
control input for passing only signals appearing at the output of
said small radiant energy detector;
first frequency selective means having its input electrically
coupled to said small radiant energy detector and having a narrow
frequency passband centered around said predetermined frequency,
said switch means having its control input electrically coupled to
the output of said first frequency selective means;
a source of voltage that exceeds a predetermined value a
predetermined time after the launching of said missile electrically
coupled to the control input of said switch means;
second frequency selective means having its input coupled to the
output of said switch means and having a narrow frequency passband
centered around said predetermined frequency and having an
automatic gain control for maintaining signals at the output
thereof at a substantially constant amplitude;
an amplitude modulation demodulator having its input coupled to the
output of said second frequency selective means;
a variable gain amplifier having its input coupled to the output of
said demodulator, said variable gain amplifier having a gain that
gradually increases from the time said missile is launched;
phase detecting means having an input coupled to the output of said
variable gain amplifier;
reference means associated with said reticle for developing a pair
of periodic reference signals having phases related to the
instantaneous angular position of said reticle, said reference
signals being in phase quadrature with each other, said reference
signals being coupled to the reference input of said phase
detecting means; and
a signal transmission circuit having its input coupled to the
outputs of said phase detecting means and its output coupled to
said missile.
15. A system for guiding a self-propelled missile having steering
means comprising:
a source of infrared energy pulsating at a predetermined frequency
disposed on said missile and directed rearwardly thereof;
an infrared telescope having a wide field of view about an optical
axis;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a predetermined speed, said reticle having an opaque
modulation pattern thereon to provide a sinusoidally varying
average transmissivity gradient circularly around said reticle and
to provide a rapidly increasing depth-of-modulation gradient
outward from the center of said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said
reticle;
a small radiant energy detector disposed in said infrared telescope
in a position to intercept radiant energy passing through the
central portion of said reticle;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing to its output only signals
appearing at the output of said large radiant energy detector, said
switch means being responsive to a switching signal applied at a
control input for passing only signals appearing at the output of
said small radiant energy detector;
first frequency selective means having its input electrically
coupled to said small radiant energy detector and having a narrow
frequency passband centered around said predetermined frequency,
said switch means having its control input electrically coupled to
the output of said first frequency selective means;
second frequency selective means having its input coupled to the
output of said switch means and having a narrow frequency passband
centered around said predetermined frequency;
an amplitude modulation demodulator having its input coupled to the
output of said second frequency selective means;
a variable gain amplifier having its input coupled to the output of
said demodulator and having a gain that gradually increases from
the time said missile is launched;
phase detecting means having an input coupled to the output of said
variable gain amplifier;
reference means associated with said reticle for developing
periodic reference signals, the phase of which is related to the
instantaneous angular position of said reticle, said reference
signals being coupled to the reference input of said phase
detecting means; and
a signal transmission circuit having its input coupled to the
outputs of said phase detecting means and its output coupled to
said missile.
16. A system for guiding a self-propelled missile having steering
means comprising:
a source of infrared energy pulsating at a predetermined frequency
disposed on said missile and directed rearwardly thereof;
an infrared telescope, having a wide field of view about an optical
axis;
a reticle disposed at the focus of said infrared telescope
concentric with and transverse to the axis thereof;
means for rotating said reticle about the axis of said infrared
telescope at a predetermined speed, said reticle having an opaque
modulation pattern thereon to provide a sinusoidally varying
average transmissivity gradient circularly around said reticle and
to provide a rapidly increasing depth-of-modulation gradient
outward from the center of said reticle;
a large radiant energy detector disposed to receive radiant energy
intercepted by said infrared telescope and passed by said
reticle;
a small radiant energy detector disposed in said infrared telescope
in a position to intercept radiant energy passing through the
central portion of said reticle;
switch means having inputs electrically coupled to said radiant
energy detectors and normally passing to its output only signals
appearing at the output of said large radiant energy detector, said
switch means being responsive to a switching signal applied at a
control input for passing to its output only signals appearing at
the output of said small radiant energy detector;
first frequency selective means having its input electrically
coupled to said small radiant energy detector and having a narrow
frequency passband centered around said predetermined frequency,
said switch means having its control input electrically coupled to
the output of said first frequency selective means;
second frequency selective means having its input coupled to the
output of said switch means and having a narrow frequency passband
centered around said predetermined frequency;
an amplitude modulation demodulator having its input coupled to the
output of said second frequency selective means;
phase detecting means having an input coupled to the output of said
demodulator;
reference means associated with said reticle for developing
periodic reference signals, the phase of which is related to the
instantaneous angular position of said reticle, said reference
signals being coupled to the reference input of said phase
detecting means; and
a signal transmission circuit having its input coupled to the
outputs of said phase detecting means and its output coupled to
said missile.
17. An interrupted radiation source comprising:
a radiating member having a radiating surface;
means adjacent said member for heating said radiating surface;
driving means disposed adjacent said radiating member and having a
shaft spaced away from and generally parallel to said radiating
surface; and
a plane interruption member of substantially the same size and
shape as said radiating member and mechanically coupled at its edge
to the shaft of said driving means for rotation thereby about an
axis parallel with the plane surfaces of said interruption member,
said interruption member being disposed to shield said radiating
member twice during each revolution about said axis.
18. An interrupted radiation source comprising:
a cylindrical housing sealed at one end by a heat-resistant
disk;
a pyrotechnic composition disposed inside said housing;
an electrically operated firing squib disposed inside said housing
in contact with said pyrotechnic composition;
at least one piece of iron within said housing and embedded in said
pyrotechnic composition;
a speed-regulated motor secured to the outside of said housing and
having its shaft spaced away from and parallel to the plane outer
surface of said heat-resistant disk; and
an interruption disk of substantially the same size as said
heat-resistant disk and mechanically coupled at its circumference
to the shaft of said motor for rotation thereby about an axis
parallel with the plane surfaces of said interruption disk, said
interruption disk being disposed to shield said heat-resistant disk
twice during each revolution about said axis.
19. An interrupted radiation source comprising:
a cylindrical housing sealed at one end by a molybdenum disk having
a zirconium carbide coating;
a pyrotechnic composition disposed inside said housing and composed
of 56 percent iron oxide, 14 percent aluminum powder, 3 percent
boron and 27 percent barium chromate;
an electrically operated firing squib disposed inside said housing
in contact with said pyrotechnic composition;
at least one piece of iron within said housing and embedded in said
pyrotechnic composition;
a speed-regulated motor secured to the outside of said housing and
having its shaft spaced away from and parallel to the plane outer
surface of said molybdenum disk; and
an interruption disk of substantially the same size as said
molybdenum disk and mechanically coupled at its circumference to
the shaft of said motor for rotation thereby about an axis parallel
with the plane surfaces of said interruption disk, said
interruption disk being disposed to shield said molybdenum disk
twice during each revolution about said axis.
20. A reticle for amplitude modulating radiant energy
comprising:
a disk transparent to said radiant energy;
a coating opaque to said radiant energy disposed on said disk in a
pattern, said pattern comprising concentric circular opaque lines
having a progressively varying width around said disk, increasing
from a minimum at a first radial zone on said disk to a maximum at
a second radial zone on said disk and decreasing to said minimum at
said first radial zone, said first and second zones being
180.degree. opposed.
21. A reticle for amplitude modulating radiant energy
comprising:
a disk transparent to said radiant energy;
a coating opaque to said radiant energy disposed on said disk in a
pattern, said pattern comprising concentric circular opaque lines
having a progressively varying width around said disk, increasing
from a minimum at a first radial zone on said disk to a maximum at
a second radial zone on said disk and decreasing to said minimum at
said first radial zone, said first and second zones being
180.degree. opposed, said concentric circular opaque lines having a
progressively increasing maximum width from the center of said disk
to the circumference thereof, the spacing between adjacent ones of
said concentric circular opaque lines in said second zone
progressively decreasing from the center of said disk to the
circumference thereof.
22. A radiant energy telescope having dual, concentric fields of
view comprising:
a cylindrical outer case having an entrance window at one end
thereof;
first and second objective lenses having a wide field of view and
disposed within said case adjacent said window;
a transparent image surface disposed within said case at the focus
of said objective lens;
a condensing lens disposed within said case adjacent said image
surface;
a large radiant energy detector disposed within said case at a
location to receive radiant energy focused on said image surface
and passed by said condensing lens;
a light pipe disposed in the center of said condensing lens to
intercept radiant energy passing through the central portion of
said image surface; and
a small radiant energy detector disposed within said light pipe in
a position to receive radiant energy intercepted thereby.
23. In an infrared tracker, apparatus for developing a signal
indicative of the instantaneous angular position of a modulating
reticle, said apparatus comprising:
a cylindrical reticle holder of nonmagnetic material rotatably
disposed in said infrared tracker and containing said modulating
reticle;
driving means mechanically coupled to said reticle holder for
rotation thereof;
a band of magnetic material fixed to the outside of said reticle
holder, said band varying linearly in width from a maximum width to
a minimum width;
a pair of magnetic pole pieces, each disposed with an edge adjacent
said band, the width of the adjacent edges of said pole pieces
being substantially equal to the maximum width of said band;
a permanent magnet having each of its poles individually
magnetically coupled to a different one of said pole pieces;
and
a coil of wire wound around said permanent magnet.
Description
CROSS REFERENCE
This application is a substitute for U.S. Pat. application, Ser.
No. 107,861, filed May 4, 1961, for "AUTOMATIC MISSILE GUIDANCE
SYSTEM", Hamilton Barhydt, George T. Hahn and Spencer D. Howe,
Inventors.
BACKGROUND
The present invention relates to apparatus for remotely controlling
the flight of self-propelled rockets and, more particularly, to a
guidance system for automatically providing guidance signals to a
missile, enabling it to follow a line-of-sight path to a
target.
Guidance of small, self-propelled missiles by manual operation of
controls that generate guidance signals is difficult, due to the
relatively slow reflexes of human operators. An additional
disadvantage of manual guidance is the necessity that the operator
be highly skilled or highly trained.
In automatic guidance of a missile, it is desirable that the
equipment for tracking the target and providing guidance signals
remain at the launching point, rather than be incorporated into the
missile. By having the tracking and guidance equipment at the
launching point, the missile is lighter in weight and, therefore,
easier to handle. Furthermore, the missile is less likely to
malfunction, because it is less complex and the missile is also
less expensive.
When the target tracking and missile guidance equipment is to be
portable, it must be small, light and compact. On the other hand,
it must operate efficiently, in spite of close proximity to the
ground. Radar tracking equipment, for example, is subject to
reflection from irregularities in the terrain, usually known as
ground clutter. Infrared tracking equipment is subject to unwanted
radiation from background and surrounding objects, as well as from
the sun, if it is in the field of view.
SUMMARY
The automatic missile guidance system comprises a sight with which
a gunner establishes a line-of-sight from the gun position to the
target. When the missile is launched, a source of pulsating,
radiant energy on the rear of the missile is detected by the sight.
A guidance unit connected to the sight produces steering commands
related to the deviation of the missile from the line-of-sight.
Means interconnecting the guidance unit and the missile transmit
the guidance signals to the missile.
Accordingly, it is an object of the present invention to provide an
automatic missile guidance system that is small in size, light in
weight and relatively inexpensive. Another object of the invention
is the provision of a missile guidance system that is automatic in
nature and does not require guidance signals to be manually
generated by a human operator. Yet another object of the present
invention is to provide an automatic missile guidance system that
includes missile tracking equipment of a simple, yet effective
type. A further object of the invention is the provision of a
missile guidance system that includes infrared tracking equipment
that is relatively insensitive to background radiation from the sun
and other intense sources of radiation.
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 throughout the figures
of the drawing.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a pictorial representation of an embodiment of an
automatic missile guidance system, in accordance with the present
invention;
FIG. 2 is a functional block diagram of the missile guidance system
of FIG. 1;
FIG. 3 is a side view of a missile suitable for use in the missile
guidance system of FIGS. 1 and 2;
FIG. 4 is a rear view of the missile of FIG. 3;
FIG. 5 is a side view, partly in section, of an embodiment of an
interrupted infrared radiation source, in accordance with the
invention, for use in the missile guidance system of FIGS. 1 and
2;
FIG. 6 is an end view of the interrupted radiation source of FIG.
5;
FIG. 7 is a side view, partly in section, of an embodiment of an
infrared telescope, in accordance with the invention, for use in
the missile guidance system of FIGS. 1 and 2;
FIG. 8 is a reduced transverse sectional view of the infrared
telescope taken substantially as indicated by line 8--8, FIG. 7,
and illustrating the reference signal generators associated with
the reticle;
FIG. 9 is an enlarged view of a reticle modulation pattern for use
in the infrared telescope of FIG. 7;
FIG. 10 is a graph illustrating the modulation characteristic of
the reticle pattern of FIG. 9;
FIG. 11 is a circuit diagram in block form of a signal processing
circuit, in accordance with the present invention and used in the
missile guidance system of FIGS. 1 and 2; and
FIG. 12 is a graph illustrating the gain characteristic of the
variable gain amplifier in the signal processing circuit of FIG.
11.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
In accordance with these and other objects of the invention, a
guidance unit is provided that may be similar in size and
appearance to a rifle. The guidance unit is provided with a visual
sight, or telescope, through which an operator views the target,
thus establishing a line-of-sight between the launching point and
the target. A remotely-controlled missile launched toward the
target is viewed by an infrared telescope disposed on the guidance
unit and aligned with the visual telescope. The missile is provided
with a periodically interrupted infrared source that emits a
distinctive frequency, pulsating radiation that is intercepted by
the infrared telescope, enabling subsequent circuits to
discriminate against background radiation. A rotating reticle is
provided in the infrared telescope to amplitude modulate the
intercepted pulsating radiant energy to provide information as to
the magnitude and direction of deviations of the missile from the
line-of-sight. The infrared telescope is provided with dual
infrared detection cells to convert the modulated radiant energy
into a modulated carrier wave that is supplied to a signal
processing circuit. The use of two detection cells permits the
infrared telescope to selectively have both a wide field of view
and a narrow field of view, the latter having greater sensitivity.
The detection cell for the large field of view is not saturated by
radiation from intense sources of background energy in its field of
view. Provisions are made for automatically switching from the wide
field cell to the narrow field cell when the target appears in the
narrow field of view, thus providing greater sensitivity and
discrimination against sources of extraneous radiation that are not
in the narrow field of view. Furthermore, the signal processing
circuit has a narrow bandpass characteristic to exclude extraneous
signals due to radiation sources other than that on the missile.
The signal processing circuit develops steering signals that are
transmitted to the missile, causing it to follow the line-of-sight
to the target.
An embodiment of an automatic missile guidance system, in
accordance with the present invention, is shown pictorially in FIG.
1 and functionally in FIG. 2. A portable guidance unit 20 is
provided that may be similar in size and appearance to a rifle
stock. The guidance unit 20 is pointed or aimed toward a target 21
and automatically provides steering signals to a ground-launched,
self-propelled missile 22 to direct it toward the target 21. The
guidance unit 20 includes a visual telescope 23 mounted to a gun
stock 24 for use by a human operator to establish a line-of-sight
25 between the guidance unit 20 and the target 21. An infrared
telescope 26 is also disposed on the gun stock 24 and aligned with
the visual telescope 23 for intercepting pulsating infrared
radiation from the missile 22 and detecting deviations of the
missile 22 from the line-of-sight 25. Inasmuch as the infrared
telescope 26 and the visual telescope 23 are side by side, their
optical axes cannot be exactly coincident. However, the few inches
difference between the axes make little difference in the operation
of the system and the axes may actually intersect at or near the
target 21. The forward portion 27 of the guidance unit 20 contains
a signal processing circuit that develops steering signals and
transmits them to the missile 22 via a cable 30, a box 31
containing missile launching and signal translation circuits, and
via wires 32 trailing behind the missile 22.
The gun stock 24 may be of the type normally used for rifles, but
modified for the purpose at hand. A folding bipod 28, about which
the assembly can pivot, provides a forward support. It will be
apparent that other convenient arrangements may be used in place of
the gun stock 24. For example, the guidance unit 20 may be mounted
to a tripod in a manner such that it may be easily carried and
easily pointed in any direction. The visual telescope 23 may have a
variable power and may be of the type normally used as a telescopic
sight on a rifle, having cross hairs in the center of the field of
view.
The missile 22 may be any remotely guided type, such as the Nord
Aviation SS-11-A--1, for example, which is described in the Nord
Aviation brochure "SS-11 Teleguided Missile Type 5210". This
missile 22 receives its guidance signals over wires 32 trailing
behind the missile 22. However, the automatic missile guidance
system of the present invention is not restricted to use with this
particular missile 22, and may be easily adapted for use with a
missile that receives its guidance signals in some other way as,
for example, by radio transmission.
Although the missile 22 itself is not part of the present
invention, the outside configuration of the missile 22 is shown in
FIGS. 3 and 4 to illustrate how the present invention is adapted
for use therewith. The missile 22 is initially propelled for a
short time by a booster rocket motor which discharges through
nozzles 40 and 41 disposed on opposite sides of the missile 22 near
the aft end thereof, and is subsequently propelled for the
remainder of the flight by a sustainer rocket motor. The sustainer
motor nozzle is not shown in FIGS. 3 and 4, but is centrally
located in the aft end of the missile 22. Control moments result
from solenoid-actuated spoilers operating in the exit area of the
sustainer nozzle. The spoilers oscillate continuously and, by
controlling the spoiler duty cycle, proportional control is
achieved. Guidance signals are received by the missile 22 through
the wires 32 connected to the box 31, which may be representative
of the Nord T-9 generator, for example, and which contains missile
firing circuits and circuits for converting guidance signals into
the form best adapted for transmission to the missile 22. The box
31 may also contain manual guidance arrangements. On the missile
22, the wires 32 are initially contained in two bobbins disposed in
bobbin cases 42 and 43 disposed on opposite sides of the missile 22
near the aft end thereof. The wires 32 are unreeled during flight
of the missile 22. Since the missile 22 is intentionally rolled
during flight, the guidance signals are resolved into pitch and yaw
commands by a vertical gyro and a four-segment commutator. The near
ends of the guidance wires 32 are attached to an anchor block 44
initially attached to the rear of the missile 22, as seen in FIGS.
3 and 4, but left behind as the missile 22 is launched, as shown in
FIG. 1.
In accordance with the present invention, the missile 22 is
provided with an interrupted radiation source 50, disposed on the
aft end thereof, that generates distinctive infrared radiation
pulsating at a constant frequency. As will be fully described
hereinafter, pulsation is caused by an interrupter operated by a
battery 51, also located on the aft end of the missile 22. The
battery 51 is connected to the source 50 when the missile 22
separates from the block 44. The interrupted infrared radiation
source 50 is shown in FIGS. 5 and 6 and comprises a cylindrical
housing 52 filled with a pyrotechnic composition 53 that produces a
thermite reaction when ignited. A suitable mixture for the
pyrotechnic composition 53 may be, for example, iron ozide, 56
percent (by weight); aluminum powder, 14 percent; boron, 3 percent;
and barium chromate, 27 percent. Another suitable mixture is
molybdic trioxide, 80 percent and aluminum powder, 20 percent.
Other mixtures may also be found to be satisfactory. An
electrically operated firing squib 54 is disposed inside the case
52 in contact with pyrotechnic composition 53. The radiating
surface is a molybdenum disk 55, having a zirconium carbide coating
and which is disposed near one end of the housing 52. At least one
heavy metal bar 56 is embedded in the pyrotechnic composition 53 to
produce a hot metal mass that heats the molybdenum disk 55 when the
radiation source 50 is ignited. The bar 56 may be made of iron or
may be made of 90 percent tungsten and 10 percent copper or nickel,
for example.
The radiant energy from the molybdenum disk 55 is interrupted by a
rotating disk 58, similar to a butterfly valve, and rotatably
secured adjacent the molybdenum disk 55. The butterfly disk 58 is
mechanically coupled at the circumference thereof to the shaft of a
direct current motor 57, having a regulated speed which, in the
present example, is 100 revolutions per second. The motor 57 is
secured to the housing 52 of the radiation source 50. As the
butterfly disk 58 rotates, it interrupts the radiation from the
molybdenum disk 55 twice each revolution to produce interrupted
infrared radiation pulsating at a frequency of 200 cycles per
second.
It will be understood that other types of sources of interrupted
radiation may be utilized, if desired. For example, an intense
electric lamp, having an interrupter, produces pulsating radiation
suitable for use in a system, according to the present invention.
Additionally, mechanical devices may be employed to cause pulsation
of the plume of exhaust gases from the rocket motor or a portion of
these gases that is bypassed to a separate outlet.
The infrared telescope 26, shown in FIG. 7, has a generally
cylindrical outer case 60, with an entrance aperture 61 at one end
to admit radiant energy. A window 62 is disposed in the entrance
aperture 61 and may, if desired, have a multilayer interference
filter deposited on the inner surface to pass only radiation having
wavelengths within a desired band. Adjacent the window 62 are first
and second objective lenses 63 and 64 that focus intercepted
radiant energy onto a flat surface of a disk-shaped reticle 65
disposed transversely to and concentric with the optical axis of
the objective lenses 63 and 64.
Adjacent to the reticle 65 is a condensing lens 66. A large radiant
energy detector 67 is disposed adjacent and generally parallel
thereto in the end of the infrared telescope case 60, opposite the
entrance aperture 61. The condensing lens 66, in effect, focuses
the image of the entrance aperture 61, as seen through the first
and second objective lenses 63 and 64 and the reticle 65, onto the
large radiant energy detector 67, and, thus, the radiant energy
passing through the condensing lens 66 more or less uniformly
illuminates the large radiant energy detector 67. A large radiant
energy detector is required so that radiant energy from the sun or
other intense sources of background radiation, which may lie within
the large field of view of the infrared telescope 26 will not
saturate the signal generating capability of the detector and is
also required by a law of geometrical optics, known as the Abbe
sine condition, which states, in effect, that for any given field
of view, there is a minimum size, increasing with the size of the
field of view, for the image of the entrance aperture 61 that can
be focused by the condensing lens 66 onto a detector. Both these
conditions are met, in the present example, by the large radiant
energy detector 67, which is approximately 3 inches on a side. A
detection cell of this size is normally a cell blank from which
smaller cells are ordinarily cut, or several cell blanks placed
side by side.
In the center of the condensing lens 66, concentric with the
optical axis of the lenses 63, 64 and 66, is located a generally
cylindrical light pipe 68, having a small radiant energy detector
70 disposed at the end thereof. The internal shape of the light
pipe 68 is that of a hollow, truncated cone, the larger diameter
end being flush with the surface of the condensing lens 66 adjacent
the reticle 65. The smaller internal diameter end of the light pipe
68 extends away from the reticle 65 and through the opposite
surface of the condensing lens 66. The internal surface of the
light pipe 68 is highly polished and the small radiant energy
detector 70 is disposed at the small diameter end of the light pipe
68. Energy intercepted by the light pipe 68 is directed to the
small detector 70 by reflection from the inner surface of the light
pipe 68. Thus, the light pipe 68 acts as a condensing lens to
concentrate diverging energy passing through the reticle 65 onto
the small detector 70.
The optical elements of the infrared telescope 26 may be made of
any material that is transparent to the radiant energy of interest
and has a suitable index of refraction. For intercepted radiant
energy in the infrared spectrum, silicon or sapphire, for example,
may be used. Lenses for use in the infrared spectrum may be opaque
to visible light. The radiant energy detectors 67 and 70 may be
made of any suitable material for converting radiant energy in the
spectrum of interest into an electrical signal. In the present
embodiment of the invention, lead sulfide, for example, will be
found satisfactory. Although the present invention is described
with reference to operation with radiant energy in the infrared
portion of the radiant energy spectrum, it is to be expressly
understood that the apparatus may be easily adapted for use with
radiant energy in other parts of the radiant energy spectrum, such
as in the visible or ultraviolet portion, for example. This may be
done by proper selection of the lens materials and of the radiant
energy detectors 67 and 70, in accordance with well-known
principles.
The field of view of the infrared telescope 26 is 40 degrees.
Images of sources of radiant energy formed on the reticle 65 by the
objective lenses 63 and 64 are small and sharply defined in the
central portion of the reticle 65. However, sources near the outer
edge of the field of view form images near the circumference of the
reticle 65 that are slightly out of focus and, consequently, larger
and somewhat blurred. Inasmuch as accuracy is required only on or
near the optical axis of the infrared telescope 26, the blurring of
images distant from the optical axis is not deleterious to the
operation of the system.
It will be apparent that the major portion of the optical field of
view of the infrared telescope 26 is intercepted by the large
radiant energy detector 67. However, the small radiant energy
detector 70 intercepts the central 6 degrees of the optical field
of view. Thus, the infrared telescope 26 has dual concentric
radiant energy detectors 67 and 70, one providing a wide field of
view and the other providing a narrow field of view.
The reticle 65 amplitude modulates intercepted radiant energy in a
manner that will be fully described hereinafter. The reticle 65 is
a flat disk made of a material that is transparent to radiation in
the spectrum of interest, but has a partially opaque modulating
pattern superimposed thereon. The reticle 65 is rotated to produce
the modulation, and for this purpose, is mounted at its
circumference in a generally cylindrical holder 72. An internal
projection 73 of the infrared telescope case 60 rotatably supports
the reticle holder 72 by means of ball bearings 74 that engage the
inner surface of the holder 72. A ring gear 75 is disposed on the
outside of the reticle holder 72. An electric motor 76 disposed
outside the infrared telescope case 60 has a shaft 77 projecting
into the case 60, to which is attached a driving gear 78 that is
meshed with the ring gear 75. In this manner, the motor 76 causes
the reticle 65 to rotate about its center, which is on the optical
axis of the objective lenses 63 and 64. The reticle 65 is rotated
at a constant rate of 18 revolutions per second in the present
example.
The reticle 65 has associated with it an arrangement for producing
phase reference signals indicative of the instantaneous angular
position of the reticle 65. For this purpose, the reticle holder 72
is provided with a band 80 around the outside thereof, adjacent the
ring gear 75. The band 80 is made of magnetic material, such as
steel, while the reticle holder 72 is made of nonmagnetic material,
such as aluminum. The band 80 tapers in width from a maximum width
point 81 (FIG. 7), to a minimum width point 82. The band 80 tapers
linearly and the maximum point 81 is diametrically opposite the
minimum point 82. A first coil and magnet assembly 79 is provided,
with a first pole piece 83 fastened to the outer case 60 of the
infrared telescope 26 adjacent and extending toward the reticle
holder 72 and spaced slightly away from the band 80. A permanent
bar magnet 84 extends out from the pole piece 83 and has wound
around it a wire coil 85. Referring to FIG. 8, a second pole piece
86 is at the other side of the magnet 84 and also extends toward
the bank 80. The faces of the pole pieces 83 and 86 are as broad as
the width of the band 80 at the maximum point 81. The case 60 that
encloses the phase reference signal coil and magnet assembly 79 is
nonmagnetic.
Thus, the magnet 84 has a magnetic circuit that extends from one
pole of the magnet 84 through the associated pole piece 83, through
a portion of the band 80 on the reticle holder 72, through the
second pole piece 86 to the remaining pole of the magnet 84. As the
reticle holder 72 rotates, the width of the band 80 in the magnetic
circuit of the magnet 84 varies, so that a variable reluctance path
exists between the poles of the magnet 84. This, in turn, causes a
potential to be developed in the coil 85 around the magnet 84.
Hence, as the reticle holder 72 and reticle 65 rotate, a periodic
signal is generated in the coil 85, whose phase is directly related
to the instantaneous angular position of the reticle holder 72. A
second coil and magnet assembly 87 (FIG. 8), similar to the first,
is located in a position 90.degree. displaced from the first coil
and magnet assembly 79 and produces a second phase reference signal
that is displaced 90 degrees in phase from the first phase
reference signal.
The modulation pattern superimposed on reticle 65, and shown in
FIG. 9, is opaque to radiation in the spectrum of interest. The
pattern can be considered to consist of a series of adjacent,
concentric, circular bands. Each band is divided into two parts,
one part of which is opaque and the other part remaining
transparent, by a circle that is not concentric with the center of
the pattern. The area in each band between the outer boundary
circle and the dividing circle is opaque, and the area between the
dividing circle and the inner boundary circle is transparent. In
the outer part of the reticle 65, the dividing circle in each band
is tangent with the outer boundary circle in a maximum transmission
zone 90 and tangent with the inner boundary circle in a minimum
transmission zone 91, 180.degree. away from the maximum. In the
inner portion of the reticle 65, the dividing circle is not tangent
to either the inner or outer boundary circles, but the separation
between the dividing circle and the outer boundary circle in the
maximum transmission zone 90 is equal to the separation between the
dividing circle and the inner boundary circle in the minimum
transmission zone 91. Furthermore, in the inner portion of the
reticle 65, the circular bands themselves are narrower than in the
outer portion.
Thus, the pattern comprises concentric, circular, opaque lines that
progressively increase in width around the reticle 65 from a
minimum width in the maximum transmission zone 90 to a maximum
width in the minimum transmission zone 91 and decrease again to the
minimum width. The size of the image focused on the reticle 65 is
larger than the width of the circular bands. The bands near the
center of the reticle 65 are narrower than the outer bands, since
the image size is smaller near the center. Since the image is
larger than the bands, the radiant energy transmitted by the
reticle 65 is amplitude modulated in a sinusoidal fashion as the
reticle 65 is rotated.
The reticle pattern is not uniform over the entire surface of the
reticle 65, as may be seen in FIG. 9, but is divided into three
distinct concentric regions having slightly different modulation
characteristics. The central portion of the field of view of the
infrared telescope 26, extending from the optical axis to a circle
1.degree. away therefrom, falls on a region of the reticle 65 where
the pattern comprises closely spaced opaque lines that provide a
modulation percentage varying from zero percent modulation at the
center to 50 percent modulation at the circumference thereof. That
is, radiant energy producing an image at the center of the reticle
65 passes through the reticle 65 unmodulated, and radiant energy
producing an image 1.degree. away from the center of the reticle 65
is only 50 percent intercepted or modulated by the opaque pattern
in the minimum transmission zone 91. The portion of the optical
field of view that extends between 1.degree. and 6.degree. from the
optical axis falls on a second region of the reticle in which the
opaque lines are not quite as closely spaced as in the first or
central region, and in which the percent modulation varies from 50
percent to 100 percent as the angular distance from the center of
the pattern increases. The third or outer region has opaque lines
spaced farther apart than the other two regions, and in which the
modulation is 100 percent throughout. The modulation characteristic
of the reticle is illustrated graphically in FIG. 10, where percent
modulation is plotted along the ordinate, as a function of the
angular distance from the center of the reticle in degrees along
the abscissa.
It will be apparent from the graph of FIG. 10 that the percentage
of modulation of radiant energy from the radiation source 50 on the
guided missile 22 is a function of the angular distance from the
optical axis of the image thereof. Accordingly, the amplitude of
the modulation is indicative of the angular deviation of the
missile 22 from the line-of-sight 25. Furthermore, the closer the
image is to the optical axis of the infrared telescope 26, the
greater is the amount of modulation produced per unit of angular
deviation. This is indicated by the steep slope of the curve of
FIG. 10, as it passes through zero. Accordingly, the subsequent
control circuits do not need to have an extremely high gain and the
noise in the system is maintained at acceptable levels, due to a
relatively large deviation signal being provided, even when the
deviation is small.
The opaque pattern may be applied to the reticle 65 by various
methods. One process that has been found satisfactory is a
photographic process in which the surface of the reticle 65 to
which the pattern is to be applied is first silvered and then
coated with a photo-resist material. An image of the pattern is
focused on the surface of the reticle 65, after which the reticle
65 is placed in an etchant bath, where portions of the silvered
area are etched away to leave the opaque pattern.
FIG. 11 shows a circuit diagram in block form of an embodiment of a
signal-processing circuit, in accordance with the present
invention. In this circuit, the radiant energy from the source 50
on the missile 22 intercepted by the infrared telescope 26 is
converted into guidance signals that are transmitted to the missile
22. As mentioned previously, the interrupted radiation source 50 on
the missile 22 emits radiant energy pulsating at 200 cycles per
second. Pulsating radiant energy intercepted by the infrared
telescope 26 is amplitude modulated by the reticle 65 at 18 cycles
per second and falls on the radiant energy detectors 67 and 70. The
modulated, pulsating radiant energy is converted to an
amplitude-modulated electrical carrier wave, or tracking signal, by
the detectors 67 and 70. The carrier wave frequency is 200 cycles
per second and the modulating wave frequency is 18 cycles per
second, so that the tracking signal occupies a frequency band from
182-218 cycles per second.
Electrical signals from the small radiant energy detector 70 are
applied to a preamplifier 100 for amplification. The preamplifier
100 may be any low noise amplifier circuit having a bandwidth of
approximately 150-300 cycles per second. Electrical signals from
the large radiant energy detector 67 are also applied to the input
of a preamplifier 101, which may be identical to the preamplifier
100 associated with the small detector 70. Output signals from the
preamplifiers 100, 101 are applied to a single-pole, double-throw
relay 102, the small cell preamplifier 100 being connected to the
normally open contact 103 and the large cell preamplifier 101 being
connected to the normally closed contact 104. Thus, either the
signal from the small detector 70 or the signal from the large
detector 67 appears at the switch arm 105, depending upon whether
the switching relay 102 is energized or de-energized. The
preamplifiers 100, 101 are phase compensated to provide identical
amounts of phase shift to signals amplified thereby. Accordingly,
when the relay 102 switches between the outputs of the
preamplifiers 100, 101, there will be no phase discontinuity in the
tracking signal.
A circuit is provided for automatically controlling the switching
of the relay 102. The output of the small cell preamplifier 100 is
connected to the input of a narrowband amplifier 107. The bandwith
of this amplifier 107 is from approximately 175 to 225 cycles per
second, so that it is responsive only to the tracking signal
derived from the radiation from the missile 22. Inasmuch as
background radiation is not pulsating at 200 cycles per second, the
electrical signal resulting therefrom is discriminated against by
the narrowband amplifier 107, because its frequency is outside the
passband thereof. The output signal of the narrowband amplifier 107
is applied to a rectifier and filter 108 that develops a
direct-current voltage which is applied to the input of a switch
amplifier 110. The relay 102 has its energizing coil connected to
the output of the switch amplifier 110.
The relay 102 is normally not energized and, in this condition, the
switch arm 105 is connected to the large detector 67. When a signal
of sufficient amplitude and having a frequency in the band from 175
to 225 cycles per second appears at the output of the small
detector 70, it is amplified, rectified and applied to the switch
amplifier 110, which energizes the relay 102. Subsequently, if the
signal from the small detector 70 falls below a predetermined
amplitude, the relay 102 is de-energized. Thus, when the missile 22
is in the narrow 6.degree. field of view of the small detector 70,
the relay 102 automatically switches to the small detector 70, and
switches back again when the missile 22 leaves the field of view of
the small detector 70.
To prevent rapid switching back and forth of the relay 102 when the
missile 22 momentarily passes through the field of view of the
small detector 70, the rectifier and filter 108 is provided with a
suitable time constant. Thus, the missile 22 must be in the field
of view of the small detector 70 for a predetermined length of time
before the voltage at the output of the rectifier and filter 108
can increase to the level at which switching takes place.
Similarly, when the missile 22 leaves the field of view of the
small detector 70, the voltage at the output of the rectifier and
filter 108 gradually decreases to the switching level. The switch
amplifier 110 is biased to have a suitable threshold level, so that
when the applied voltage is above the threshold, the relay 102 is
energized and, when the applied voltage is below the threshold, the
relay 102 is de-energized.
When the missile 22 is close to the target 21, the relay 102 is
locked into the energized position. This is accomplished by means
of a switch control voltage source 111 that applies a voltage that
exceeds the threshold level to the input of the switch amplifier
110 at the end of a predetermined time interval. The switch control
voltage source 111 is a capacitor-charging circuit that begins
operation at the time the missile 22 is fired. After 6.7 seconds,
the potential at the output of the switch control voltage source
111 exceeds the threshold level, causing the relay 102 to be locked
into the energized position.
The switch arm 105 of the relay 102 is connected to the input
circuit of a second narrowband amplifier 112 that is similar to the
narrowband amplifier 107 in the relay control circuit.
Specifically, the narrowband amplifier 112 has a bandwidth of 175
to 225 cycles per second to pass only the tracking signal while
excluding background signals. An AGC (Automatic Gain Control)
circuit 113 is connected to the narrowband amplifier 112. This
circuit 113 rectifies and filters the tracking signal appearing at
the output of the narrowband amplifier 112 to develop an AGC
voltage that is applied to the narrowband amplifier 112 to control
the gain thereof. Thus, as long as the missile 22 is in the field
of view of the infrared telescope 26, the tracking signal appearing
at the output of the narrowband amplifier 112 has a substantially
constant amplitude.
A demodulator 114 is also connected to the output of the narrowband
amplifier 112 and comprises a rectifier and filter that demodulates
the tracking signal to recover the 18-cycle-per-second modulating
wave or error signal introduced by the reticle 65. The error signal
is applied to the input circuit of a variable gain amplifier 115
that increases the amplitude of the missile error signal as the
missile 22 approaches the target 21 to account for the increasing
distance between the guidance unit 20 and the missile 22. If this
were not done, the missile 22 would undercorrect for errors as it
approached the target 21. Accordingly, the gain of the variable
gain amplifier 115 is gradually increased throughout the time of
flight, in accordance with the graphical representation shown in
FIG. 12, in which the gain of the amplifier 115 is plotted along
the ordinate as a function of elapsed missile flight time plotted
along the abscissa.
The variable gain amplifier 115 is an amplifier having a negative
feedback loop, including a resistive voltage divider network. A
pair of diodes connect a relatively low resistance in shunt with a
portion of the feedback network. The diodes are normally biased to
be nonconductive, at which time the negative feedback is large and
the gain of the amplifier 115 is low. When the diodes become
conductive, the impedance of the feedback network is decreased,
which decreases the negative feedback and the gain of the amplifier
115 increases. The gain of the variable gain amplifier 115 is
controlled by a gain control voltage source 116, which is similar
to the switch control voltage source 111 of the signal switching
circuit. The gain control voltage source 116 includes a capacitor
charging circuit that is set into operation when the missile 22 is
fired. The exponential charging voltage is applied to the diodes in
the feedback network of the variable gain amplifier 115 to cause
them to gradually become more and more conductive as the charging
voltage increases.
The error voltage from the output of the variable gain amplifier
115 is applied to a pair of circuits that develop steering signals.
These circuits are a pitch phase detector 117 and a yaw phase
detector 118, which may be conventional phase detector circuits.
Reference signals for the pitch and yaw phase detectors 117 and 118
are derived from the coil and magnet assemblies 79 and 87
associated with the reticle holder 72, previously described. These
coil and magnet assemblies 79 and 87 develop periodic wave
reference signals that are 90.degree. displaced in phase with
respect to each other, and have a fixed phase relationship to the
angular position of the reticle holder 72. The pitch and yaw
reference signals are applied to squaring circuits 120 and 121 that
amplify and clip the reference signals to develop square waves that
are then applied to the pitch and yaw phase detectors 117 and
118.
To establish the proper phase relationship between the modulation
pattern on the reticle 65 and the pitch and yaw reference signals,
an image of the radiation source 50 is focused on the reticle 65 at
a position vertically displaced from the center of the reticle 65.
The angular position of the reticle 65, with respect to the reticle
holder 72, is manually adjusted until the yaw phase detector 118
provides no steering signal output and the pitch phase detector 117
provides a maximum steering signal output when the guidance unit 20
is operating. When the image is moved to a position laterally
displaced from the center of the reticle 65, the pitch phase
detector 117 provides no steering signal output and the yaw phase
detector 118 provides a maximum steering signal output.
The output circuits of the pitch and yaw phase detectors 117, 118
are each connected to corresponding pitch and yaw compensation
networks 122, 123 that provide stability and damping, in accordance
with well-known principles of feedback control systems. That is,
the compensation networks 122 and 123 modify the natural response
of the guidance system to disturbances thereof by means of such
well-known techniques as basic lead compensation and error rate
damping, for example. The compensation networks 122, 123 are
specifically adapted to the type of missile 22 with which the
guidance system is to be used. This compensation provides guidance
system stability and satisfactory transient response, regardless of
displacements of the missile 22 from the line-of-sight 25 at launch
and internal system noise, for example. In addition, a gravity bias
may be provided to prevent collision of the missile 22 with the
ground during overshoots around the line-of-sight 25.
The output of the pitch and yaw compensation networks 122, 123 are
connected to a steering signal transmission circuit 124 which
applies the steering signal to the missile 22. The steering signal
transmission circuit is, in the present example, a box 31, shown in
FIG. 1, which is representative of the Nord T-9 generator and which
contains firing and signal conversion circuits and to which are
connected the wires 32 trailing behind the missile. It will be
understood that the steering signal transmission circuit 124 may be
any other form of guidance command link, such as a radio
circuit.
In operation, the portable guidance unit 20 is aimed toward the
target 21 by the operator who sights through the visual telescope
23. The operator maintains the portable guidance unit 20 trained on
the target 21 at all times during the flight of the missile 22,
even if the target 21 is moving. This establishes a line-of-sight
25 between the guidance unit 20 and the target 21. The missile 22
is then launched into the field of view of the infrared telescope
26. The launching process sets into operation the interrupted
radiation source 50 on the rear of the missile 22, which emits
infrared energy pulsating at 200 cycles per second. The launching
of the missile 22 also sets into operation the charging of
capacitors in the switch control voltage source 111 and the gain
control voltage source 116.
The pulsating radiant energy from the missile 22 is intercepted by
the infrared telescope 26 where it is focused on the reticle 65.
The reticle 65 modulates the intercepted radiant energy at 18
cycles per second, the phase of the modulating signal being
dependent upon the angular direction of the missile 22 from the
line-of-sight 25. The amplitude of the modulating signal is
dependent upon the radial distance of the missile 22 from the
line-of-sight 25. The modulated intercepted radiant energy is then
concentrated on the radiant energy detectors 67 and 70, which
convert the radiant energy into an amplitude-modulated electrical
carrier wave.
It will be apparent that background radiation from objects other
than the missile 22 and including the sun, do not result in a
similar electrical signal because the radiant energy emitted by
background objects is not pulsating at 200 cycles per second.
However, background radiation will be modulated to some extent at
18 cycles per second by the reticle 65. Signals from the radiant
energy detectors 67 and 70 are applied to the preamplifiers 100,
101, which have a bandwidth of 150 to 300 cycles per second and,
therefore, will not pass electrical signals at 18 cycles per second
that are derived from background radiation.
Even when intense sources of background radiation, such as the sun,
are in the field of view, the large detector 67 is not saturated.
Thus, the pulsating energy from the missile 22 continues to produce
electrical signals at the output of the large detector 67.
Accordingly, the missile signal is at all times distinguishable
from background signals.
The amplitude modulated electrical carrier wave, or tracking
signal, then passes through the contacts of the relay 102 to the
input of the narrowband amplifier 112, which further discriminates
against extraneous signals. The AGC circuit 113, associated with
the narrowband amplifier 112, maintains the tracking signal at a
substantially constant amplitude at the output of the narrow band
amplifier 112. However, it will be understood that the time
constant is sufficiently long so that the modulation will not be
suppressed by gain control action. The tracking signal is then
applied to the demodulator 114, where it is demodulated to recover
the 18-cycle-per-second modulating wave or error signal introduced
by the reticle 65. The phase of the error signal is a function of
the angular direction of the image from the center of the reticle
65 and the amplitude of the error signal is a function of the
radial distance of the image from the center of the reticle 65.
The error signal is applied to the input of the variable gain
amplifier 115, which adjusts the amplitude of the error signal to
account for the increasing distance of the missile 22 from the
guidance unit. The error signal is then applied to the pitch and
yaw phase detectors 117 and 118, which convert the error signal
from polar form into rectangular form by means of the phase
reference signals developed by the coil and magnet assemblies 79
and 87 associated with the reticle 65. This operation results in a
pitch steering signal at the output of the yaw phase detector 118.
The pitch and yaw steering signals are then applied to the pitch
and yaw compensation networks 122 and 123, which provide guidance
system stability. The steering signals are then applied to the
steering signal transmission circuit 124, where they are
transmitted to the missile 22 to correct for deviations thereof
from the line-of-sight 25.
Inasmuch as this is a closed loop system, the missile 22 always
tends to follow the line-of-sight 25. If the missile 22 is not on
the line-of-sight 25 at launch, the error signal produced steers
the missile 22 back onto the line-of-sight 25 because the system
operates to reduce the error signal and the steering signals to
zero.
Initially, the missile 22 may have a large heading error, with
respect to the line-of-sight 25. However, the extremely wide field
of view of the infrared telescope 26 intercepts the pulsating
radiation from the missile 22 and initiates the controlling signals
to bring it back to the line-of-sight 25. As the missile approaches
the line-of-sight 25, it enters the central 6.degree. of the field
of view of the infrared telescope 26, at which time the pulsating
infrared radiation falls on the small radiant energy detector 70.
The result is that the tracking signal is applied to the
narrow-band amplifier 107 in the signal switching circuit, which
discriminates against background signals and applies it to a
rectifier and filter 108 that, after a short time interval, builds
up a voltage that exceeds the threshold of the switch amplifier
110, causing the relay 102 to connect the output of the small
detector 70 into the guidance system.
The narrow field of view of the small detector 70 eliminates many
sources of background radiation and, therefore, provides increased
discrimination against background signals. The small detector 70
also provides greater sensitivity. Should the missile 22 leave the
central 6.degree. of the field of view of the infrared telescope
26, the voltage at the output of the rectifier and filter 108
decreases after a short time interval to a value less than the
threshold level of the switch amplifier 110, at which time the
large detector 67 is connected into the guidance loop by the relay
102. 6.7 seconds from the time of missile launch, the gradually
increasing voltage from the switch control voltage source 111
exceeds the threshold of the switch amplifier 110 and locks the
small detector 70 into the guidance loop as, at that time, the
missile 22 should be within the central 6.degree. of the field of
view of the infrared telescope 26.
When the image is on the central 1 degree of the reticle 65, the
error signal is larger per unit of angular deviation from the
center of the reticle 65. Thus, the signal processing circuit does
not need to have an extremely high gain.
As the missile 22 approaches the target 21, the gradually
increasing voltage at the output of the gain control voltage source
116 causes the gain of the variable gain amplifier 115 to gradually
increase, thereby increasing the amplitude of the error signal, so
that the missile 22 does not undercorrect for errors as it
approaches the target 21.
Thus, there has been described an automatic missile guidance system
that is small in size, light in weight, relatively inexpensive, and
simple to operate. The missile guidance system of the present
invention includes infrared tracking equipment of a simple, yet
effective type, that is relatively insensitive to background
radiation from the sun and other intense sources.
* * * * *