U.S. patent number 3,883,091 [Application Number 03/600,858] was granted by the patent office on 1975-05-13 for guided missile control systems.
This patent grant is currently assigned to Bell Telephone Laboratories, Incorporated. Invention is credited to Jacob W. Schaefer.
United States Patent |
3,883,091 |
Schaefer |
May 13, 1975 |
Guided missile control systems
Abstract
1. In an antiaircraft system, target tracking means for
continuously establishing the position of a target aircraft in
space, missile tracking means for similarly establishing the
position of a missile, computing means responsive to position
information from both said target and missile tracking means to
produce control signals for said missile to insure interception of
the target thereby and means for substituting the target tracking
means for said missile tracking means as a source of missile
position information for said computing means during the final
phases of an engagement.
Inventors: |
Schaefer; Jacob W. (Watchung,
NJ) |
Assignee: |
Bell Telephone Laboratories,
Incorporated (Murray Hill, NJ)
|
Family
ID: |
24405339 |
Appl.
No.: |
03/600,858 |
Filed: |
July 30, 1956 |
Current U.S.
Class: |
244/3.13;
244/3.19 |
Current CPC
Class: |
F41G
7/30 (20130101) |
Current International
Class: |
F41G
7/20 (20060101); F41G 7/30 (20060101); F41g
009/00 (); F41g 011/00 (); F41g 007/14 () |
Field of
Search: |
;244/14,3.13,3.14 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Borchelt; Benjamin A.
Assistant Examiner: Webb; Thomas H.
Attorney, Agent or Firm: Adams, Jr.; E. W.
Claims
What is claimed is:
1. In an antiaircraft system, target tracking means for
continuously establishing the position of a target aircraft in
space, missile tracking means for similarly establishing the
position of a missile, computing means responsive to position
information from both said target and missile tracking means to
produce control signals for said missile to insure interception of
the target thereby and means for substituting the target tracking
means for said missile tracking means as a source of missile
position information for said computing means during the final
phases of an engagement.
2. In an antiaircraft system a tracking radar for establishing the
position of a target aircraft in space, a second tracking radar for
similarly establishing the position of a missile launched against
said target aircraft, computing means responsive to position
information from both of said radars to produce control signals for
said missile to insure interception of the target aircraft thereby,
means for detecting the approach of the missile to within a
predetermined separation from said target aircraft and means
responsive to said detecting means for substituting the target
tracking radar for said missile tracking radar as a source of
missile position information.
3. In an antiaircraft system target tracking means for continuously
establishing the azimuth, elevation and range to a target aircraft,
missile tracking means for similarly establishing the position of a
missile, computing means responsive to the position information
from both target and missile tracking means to control the path of
said missile to insure interception of the target thereby and means
for substituting the target tracking means for said missile
tracking means as the source of missile azimuth and elevation
information during the final phases of an engagement.
4. In an antiaircraft system a missile tracking radar operating at
a first frequency, a target tracking radar comprising a transmitter
operated at a second frequency and receiving means capable of
separate response to echo signals of both said first and second
frequencies, computing means normally responsive to target and
missile position information derived from said first and second
radars respectively to control the path of said missile to insure
interception of the target thereby and means operative during the
final phase of an engagement to substitute as a source of missile
position information for said computer the output of the target
tracking receiving means produced in response to echoes of said
second frequency in place of the output of the missile tracking
radar.
5. In an antiaircraft system a target tracking radar for
continuously establishing the position of a target aircraft in
space, a missile tracking radar for similarly establishing the
position of a missile launched against said target aircraft,
computing means responsive to position information from both said
target and missile tracking radars to produce control signals for
said missile to insure interception of the target aircraft thereby,
means for determining the angular separation in both elevation and
azimuth of the tracking beams of said radars and means operative
when both of said angular separations are reduced below
predetermined values for substituting the target tracking radar for
said missile tracking radar as a source of missile position
information for application to said computing means.
6. Apparatus for controlling a guided missile relative to a target
comprising missile tracking radar apparatus, target tracking radar
apparatus and computer apparatus connected with said respective
missile tracking and target tracking radar apparatus, said target
tracking radar apparatus comprising means independently receiving
radio signals from said missile and said target for simultaneously
supplying said computer apparatus with information signals in
accordance with the angular position of said missile and the range
and angular position of said target.
7. In an antiaircraft system target tracking means for continuously
establishing the position of a target aircraft in space, missile
tracking means independent thereof for similarly establishing the
position of a missile launched against said target aircraft, a
computer, means for applying the output of said target tracking
means to said computer, means for applying the output of said
missile tracking means to said computer after correction for
parallax between said missile and target tracking means, said
computer being arranged to produce orders for said missile to
insure interception of the target aircraft by said missile and
means operative when the target aircraft and missile approach one
another to within a predetermined separation for substituting the
output of said target tracking means as a source of missile
position information in place of the missile tracking means and the
means for correcting for parallax as to the information for which
such substitution is made.
Description
This invention relates to guided missile systems and more
particularly to improvements in those missile systems employing
command guidance techniques.
In a copending application, Ser. No. 449,396, filed Aug. 12, 1954
in the names of E. L. Norton and the present inventor and assigned
to the assignee of the present application, which matured into U.S.
Pat. No. 3,156,435 on Nov. 10, 1964, there was disclosed a guided
missile system designed for the interception and destruction of
high altitude, high performance bombing aircraft against which
adequate defense with conventional antiaircraft artillery is
inoperative. The extreme ranges at which engagement must occur to
prevent a successful bombing attack make it necessary to provide
some means for adjusting the course of a projectile launched
against the bomber during the rather considerable time of flight
between launching and interception.
As disclosed in the above-mentioned copending application, a
practical guided missile system for such applications may include a
missile, normally self-propelled, which may be controlled in
accordance with commands issued by a ground based guidance
equipment. This guidance equipment includes precision radars for
individually tracking both the target and the missile launched
against the target to obtain data as to the present positions of
both. These data are supplied to a computer which predicts the
future position of the target at an assumed time of interception
and generates orders for transmission to the missile during its
flight to control the course thereof in such a way as to insure
interception of the target at the assumed time. Equipment aboard
the missile provides a frame of reference traveling with the
missile and identifiable at the location of the ground guidance
equipment with respect to which control orders may be produced.
Although the command guidance system is considered highly efficient
there are other systems of missile guidance which have attractive
features. One of these which is now well known is the co-called
beam riding system of missile guidance. Here a tracking device,
usually a ground based radar is employed to track the target
continuously and the missile is launched in such a way as to
intercept the beam of the radar shortly after launch. Detection
devices aboard the missile respond to the beam when intercepted and
provide inputs for a missile-borne computer which determines those
adjustments in the controls of the missile which are required to
maintain the missile in the beam as the beam is swept to follow the
flight of the target. This system has the inherent advantage that a
single radar or guidance device is employed. In more detail the use
of a single radar eliminates the need for precise corrections for
parallax between the locations of the two tracking radars of the
command system. In addition, since only a single radar is employed
bore-sighting errors are not compounded. Normally such matters do
not constitute so great a disadvantage as to outweigh the greater
advantages of the command system. However, at extreme ranges and
during the final phases of an engagement slight resultant
misalignments between the two radars may seriously affect
performance. On the other hand the beam riding system has the
serious disadvantage that the missile does not follow the most
efficient course (in terms of time and fuel) to the target.
It is an object of the present invention to so modify a command
guidance system as to make available therein the attractive
features of the beam riding missile system of guidance without
incurring the penalties inherent in the beam riding system.
In accordance with this object the command missile system of the
present invention is generally the same as that of the copending
application referred to above. In addition to the elements thereof
as outlined above, means are provided for determining when the
beams of the tracking radars for both missile and target have
approached one another to within predetermined limits. At this
time, which indicates effective proximity of the missile and the
target, means are actuated for transferring the missile tracking
function from the missile tracking radar to the target tracking
radar. For this purpose modifications are made in the target
tracking radar which enable it to receive both return pulses from
the target and those from the missile which occur at a different
frequency. The switching device referred to above transfers the
missile position inputs to the computer from the missile tracking
radar to the target tracking radar and thus makes available for the
final phases of an attack the advantageous single reference system
of guidance normally associated with a beam riding system.
The above and other features of the invention will be described in
detail in the following specification taken in connection with the
drawings, in which
FIG. 1 is a block schematic diagram of the complete command missile
guidance system of the invention;
FIG. 2 is a diagram in schematic form of the missile employed in
the system of FIG. 1 showing the equipment required aboard the
missile;
FIG. 3 is a vector diagram illustrating the missile command problem
and the manner in which the orders for transmission to the missile
are computed; and
FIG. 4 is a block diagram of the control system required for
transferring the missile tracking function from the missile
tracking radar of the target tracking radar and provides details of
the control circuit shown schematically in FIG. 1.
In the broadest sense the antiaircraft guided missile system of the
invention comprises a target tracking device 10, a missile tracking
device 12, a computer 14 and a missile 16, all as shown in the
block diagram of FIG. 1. Target tracking device 10 which preferably
comprises a precision automatic tracking radar is ranged
continuously to provide data as to the present position of a target
aircraft 18. This tracking radar may, for example, be similar to
the well known SCR-584 radar which is described in detail in
"Electronics" for November 1945 beginning at page 104, for December
1945 beginning at page 104 and for February 1946 beginning at page
110. Briefly this radar is an automatic tracking radar employing
conical lobing whereby azimuth and elevation error signals are
produced by the receiver. These signals may be applied to servo
systems to cause the radar antenna to continuously track the target
in both elevation and azimuth. In addition this radar includes a
range unit also responsive to the reflected radar pulses which
automatically maintains itself adjusted to represent the slant
range to the target.
As shown in FIG. 1 target tracking radar 10 includes a transmitter
100 operating at a frequency f.sub.3 which is used only by this
radar in the present system, a receiver unit 102 tuned to receive
echo signals at a frequency f.sub.3 and a second receiver unit 104
also associated with the same radar antenna and tuned to receive
echoes at a frequency f.sub.2 which as will be described below
identifies those return signals emanating from the missile.
Receiver units 102 and 104 may be identical with the exception of
the frequency to which they are tuned and may conveniently be
connected to the same radio frequency components of the radar
transmit-receive equipment. This can be accomplished at radio
frequency through the use of a conventional branching network, the
output of which is applied to two preamplifiers each associated
with its own mixer, intermediate frequency amplifier, and other
necessary receiver circuitry. As will appear below, receiver unit
104 need not include a range unit.
The normal output of tracking radar 10 is derived from information
indicating the elevation and azimuth errors and the range to the
target as indicated by the error signals received by receiver 102
and translated into position data appropriate for transmission to
the computer in the range unit forming a part of the receiver and
in servo unit 106 which derives the azimuth and elevation signals
while orienting the antenna to reduce the error signals to zero.
This servo unit may be considered to be a typical unit and is
assumed for the present purposes that the azimuth and elevation
error signals and the range signals are converted to electrical
quantities corresponding to the azimuth range and elevation for
individual transmission over a connector 20 to a predictor 22. The
form in which the target position data are determined is not
significant since the means for converting data in one coordinate
system to another system are well known in the computer art.
Target radar 10 produces a second set of output quantities for
application to transmission link 108 which quantities are of the
same nature as those applied to transmission link 20 and are
representative of azimuth and elevation data. In this instance,
however, they are derived from receiver 104 tuned to the frequency
f.sub.2 and converted in data unit 110. It should by noted that the
quantities obtained from data unit 110 are employed only for
application to transmission link 108 and are not employed in the
automatic tracking circuitry of target tracking radar 10.
The missile tracking device 12 may be similar to target tracking
device 10 and may in the same way produce output quantities
proportional to the slant range, the elevation, and the azimuth of
the missile as measured at the location of the missile tracking
device. As shown in FIG. 1, however, certain advantageous
modifications have been made in the missile tracking device to
improve the performance thereof. Basically these modifications
involve recognition of the fact that the antiaircraft missile
presents an extremely difficult target for a tracking radar.
Accordingly it has been found desirable to employ a so-called radar
beacon system rather than a conventional radar. For this purpose
pulses are radiated from a transmitter 24 at a radio frequency
f.sub.1. The missile as will be explained in greater detail
hereinafter carries a responder which is responsive to pulses of
frequencies f.sub.1 and radiated pulses of frequencies f.sub.2.
These pulses are picked up by antenna 25 and directed to a receiver
26 responsive to that radio frequency. The receiver 26 operates in
a manner identical to that of target tracking radar 10 to provide
the required output quantities as to azimuth and elevation for
transmission over a connector 28 by way of a transfer switch 112
operated by an actuator 114 and parallax correction unit 31 to a
second predictor 30. The output derived from the range unit of
receiver 26 and representative of the range to the missile is
applied directly to parallax correction unit 31 by way of connector
29.
Normally and during all phases of an engagement other than the end
game (the final increments of the engagement), transfer switch 112
remains in the position shown in FIG. 1 whereby the output
quantities from missile tracking radar 12 are applied after
correction for parallax to predictor 30 of computer 14.
Radar beacon systems of the type contemplated for use in the
missile tracking system are well known in the art and are discussed
in detail in "Radar Beacons" by Roberts, Vol. 3 of "The Radiation
Laboratory Series, McGraw-Hill, 1947." Modification of the SCR-584
radar referred to above as illustrative of the ground based missile
tracking device, for this type of performance may easily be
accomplished merely by tuning of the receiver to frequency f.sub.2
rather than f.sub.1. The missile borne equipment will be considered
hereinafter.
As assumed above the quantities applied to predictor 22 indicate
the present position of target 18 in spherical coordinates with
respect to the location of the target tracking radar while those
applied to predictor 30 represent similar information as to the
position of the missile with reference to the location of the
missile tracking radar. For ease in computation it is considered
desirable to convert this information into rectangular coordinates
with the origin at the location of the target tracking radar. Such
coordinate conversion is well known in the art and may be
accomplished as described, for example, in U.S. Pat. No. 2,408,081
to Lovell et al. which issued, Sept. 24, 1946. Conveniently each
predictor 22 and 30 includes a coordinate converter acting to
convert input quantities to rectangular coordinates (X, Y and H
where X and Y are orthogonal axes in the horizontal ground plane
and H is the vertical distance from the XY plane) with origins at
the locations of the respective tracking radars. The necessary
offset or parallax corrections referred to above and required to
convert the data as to missile position to the coordinate system
having its origin at the location of the target tracking radar may
be set in manually in unit 31 associated with predictor 30 as
potentials of suitable polarity along the three rectangular
coordinates. These corrections are constants and once determined
with adequate precision at the time at which the two tracking
radars are emplaced, need not be changed unless the emplacement of
the guidance equipment is changed. As pointed out above, however,
the present invention permits substantial relaxation of the degree
of precision required in these corrections. The errors in
performance resulting from parallax are not of great importance in
the early phases of an engagement. In the final phase control is
switched according to the invention to eliminate the parallax
correction problem in its entirety.
Computer 14 which includes predictors 22 and 30 acts on the basis
of an assumed time of flight for the missile to reach a point of
interception with the target. Using this time of flight the future
positions of the target and missile at the predicted time of
interception are computed independently. These positions are
compared coordinate by coordinate to find the predicted position
error at the predicted time of interception. The corrections in
either or both the time of flight assumed at the outset and the
course of the missile required to insure interception may be
determined from this information. Depending upon the nature of the
missile either or both of these quantities may be altered to
eliminate the position error by the predicted time of interception.
This general philosophy of computation is similar in certain
respects to that employed in antiaircraft gun direction systems of
the type disclosed in the patent referred to above. In the computer
herein disclosed, however, many of the operations are duplicated to
provide information not only as to the target but also as to the
missile since in the present case some measure of control remains
after the missile is fired.
Assuming a time of flight, predictors 22 and 30 determine the
future positions of both missile and target at the predicted time
of intercept (T=0 where T is the time until interception). In each
predictor the present position data is differentiated to obtain
velocity and this velocity is multiplied by the time of flight to
obtain future position. These operations are carried out in each
instance in terms of components along the orthogonal X, Y and H
coordinates which are ground coordinates with the origin at the
location of the target tracking radar as indicated in FIG. 3 of the
drawing.
As in the reference patent each predictor 22 and 30 provides three
coordinate output data representing the predicted position of the
target or the missile as the case may be when the predicted time of
intercept occurs. These quantities are identified in FIG. 1 as
X.sub.T, Y.sub.T and H.sub.T for the target and X.sub.M, Y.sub.M
and H.sub.M for the missile.
These quantities are applied to a comparator 32 in which are
subtracted coordinate by coordinate to obtain predicted position
error quantities. The error outputs shown at the output of
comparator 32 are E.sub.H showing the position error in the H
direction, and E.sub.X and E.sub.Y showing the position error in
the X and Y directions, respectively.
A better understanding of the significance of these quantities may
be obtained by reference to the diagram of FIG. 3. Here the present
heading of the missile is represented by the arrow labeled n.sub.M
and that of the target by the arrow n.sub.t. If it is assumed that
at the predicted time of intercept (T=0) the missile will have
reached a position a specified a position specified with respect to
the X, Y and H axes by the quantities X.sub.M, Y.sub.M and H.sub.M
referred to above and the target a position b similarly specified
by the quantities X.sub.T, Y.sub.T and H.sub.T. The total position
error at T=0 for these assumptions is measured by the vector E, the
components of which E.sub.X, E.sub.Y and E.sub.H in the X, Y and H
directions respectively are as shown in FIG. 3. It is apparent that
these quantities are measured with reference to a set of
coordinates fixed with respect to the target tracking radar and do
not indicate directly what maneuvers must be performed by the
missile to assure interception. It is necessary, therefore, to
convert these error quantities into quantities measured with
respect to a frame of reference traveling with the missile so that
appropriate commands for missile guidance may be produced.
The method of providing the required frame of reference traveling
with the missile and still identifiable at the location of the
guidance equipment will become apparent with reference to the
diagram of FIG. 2. Here the missile 16 is shown as carrying a
so-called free-free gyroscope 34, the rotor of which is suspended
in a conventional gimbal system. The outer gimbal 36 is journalled
for rotation about the longitudinal axis of the missile while the
inner gimbal 38 is journalled in the outer gimbal for rotation
about an axis normal to the longitudinal axis of the missile. The
gyroscope rotor is in turn journalled in the inner gimbal and spins
about an axis normal to the inner gimbal axis. This third axis is
hereinafter referred to as the gyro spin axis. In accordance with
well understood principles the gyroscope acts to maintain the gyro
spin axis G.sub.A at a fixed orientation in space regardless of the
maneuvers of the missile in which the gyroscope is mounted.
Conveniently, the gyro spin axis is initially oriented in the XY
(horizontal) plane and, if possible, normal to the plane of the
initial trajectory to the target. This axis and the longitudinal
axis of the missile define a reference plane attached to the
missile and identifiable at all times at the location of the
guidance equipment as will be pointed out below.
The initial orientation of the spin axis may be determined before
the missile is launched and the heading of the missile at any time
is determined as one of the necessary functions of predictor 30. It
will be recalled that the rates of change of each of the X, Y and H
quantities representing the present position of the missile are
determined in the prediction process. These components of the
missile velocity may be combined to give a quantity proportional to
the missile velocity in the direction of the missile flight path.
This combination of velocity components is performed by missile
heading resolver 40 which is a conventional coordinate resolution
device of the type employed generally in fire control computers. It
is noted that the frame of reference traveling with the missile is
based upon the gyroscope spin axis and the longitudinal axis of the
missile while the information available to the computer includes
the orientation of the gyro spin axis and the missile velocity. It
has been found that, because of the continuous control of the
missile afforded by the command system of the invention, the
missile velocity vector may be considered to have the same
orientation as the longitudinal axis of the missile. Any error so
introduced is within the range of correction of the command system.
It will be understood, however, that if sufficient information is
available as to the aerodynamic performance additional computation
equipment may be provided to determine the orientation of the
longitudinal axis of the missile from the available missile
velocity information.
It now becomes necessary to steer the missile with respect to the
frame of reference just considered. As shown in FIG. 2 the missile
is provided with two sets of paired steering fins 40 and 42
respectively. Fins 40 are mounted on a shaft 44 normal to the
longitudinal axis of the missile and fins 42 are mounted on a shaft
46 which is normal to both the longitudinal axis of the missile and
the shaft 44. Shafts 44 and 46 thus constitute a pair of steering
axes which will be referred to as the yaw and pitch axes
respectively and which control the orientation of the missile in
two orthogonal planes. As a matter of convenience the plane normal
to shaft 44 will be referred to as the yaw plane (which is the same
as the reference plane considered above) and that normal to shaft
46 as the pitch plane (which includes the longitudinal axis of the
missile and is normal to the reference plane). Fins 40 then
constitute the yaw steering fins and fins 42 the pitch steering
fins. These fins are positioned in response to steering orders
transmitted from computer 14 to control the course of the missile
after launching.
It is apparent that the course of the missile can be properly
controlled only if the steering axes represented by shafts 44 and
46 remain respectively normal to and in the plane of reference
considered above and defined by the gyro spin axis and the
longitudinal axis of the missile. This condition requires roll
stabilization of the missile. For this purpose the missile is
equipped with paired cruciform tail fins 48 and 50 and adjustable
ailerons 52 are provided upon the trailing edges of at least one
pair of tail fins (48 in FIG. 2). Conveniently gyroscope 34 is
employed to control ailerons 52 in such a way as to roll stabilize
the missile with respect to the reference plane.
The outer gimbal 36 of the gyroscope 34 is coupled to the movable
arm of a potentiometer 54 through a shaft 56 which forms an
extension of the outer gimbal axis. Potentiometer 54 comprises a
source of error signal for a servo system including an amplifier
56, a motor 58 and suitable linkage 60, 62 whereby opposite
deflections of upper and lower ailerons 52 may be produced by motor
58. Although these elements, together with the gyroscope 34 which
constitutes the error detecting element, may be connected to form
any of a large number of known kinds of servo system, it may be
assumed for the purposes of the present description that a simple
direct current servo system is employed.
Potentiometer 54 is provided with a center-tapped winding to which
direct current potentials are applied from a source such as a
battery (not shown to avoid undue complexity in the drawing). The
circuit is so arranged that no output is produced when shaft 44 is
normal to the reference plane. Whenever the potentiometer arm is
moved from this normal (null) position an output is developed, the
amplitude and polarity of which are indicative of the amount and
direction of the roll of the missile from the desired position.
After amplification this output may control motor 58 as in the
usual direct-current servo system, causing deflection of the
ailerons in the proper direction to return the missile to the
desired orientation as indicated by the null output from
potentiometer 54.
It will be understood that as a result of such roll stabilization
fins 40 are effective to produce steering forces in the yaw plane
and the other pair of steering fins 42 act to produce steering
forces in the pitch plane. Thus there is provided a set of
orthogonal reference coordinates traveling with the missile and
comprising the yaw axis y, the pitch axis p and the missile heading
h.sub.m. Further the orientation of this reference system in space
is continuously determinable at the location of the ground guidance
equipment.
It will be recognized that by the usual process of coordinate
conversion the total position error E shown in FIG. 3 may be
expressed in terms of components along the reference axes traveling
with the missile. Such conversion may be accomplished in coordinate
resolver 64, FIG. 1 as outlined beginning page 279 of "Electronic
Analog Computers" by Korn and Korn or in U.S. Pat. No. 2,658,674 to
Darlington et al., Nov. 10, 1953 which resulted from an application
filed Feb. 13, 1945 particularly in FIG. 38 and the specification
beginning at column 92 thereof. This coordinate resolver accepts
the three position error components E.sub.H, E.sub.X and E.sub.Y
shown in FIG. 3 and in addition accepts quantities from the output
of missile heading resolver 40 indicating the orientation of
missile heading axis in space and quantities representing the
position of the gyro spin axis G.sub.a which may be set into the
coordinate resolver as constants. The outputs of coordinate
resolver 64 are E.sub.y and E.sub.p representing position errors in
the yaw and pitch planes and measured along the pitch and yaw axes
respectively and E.sub.t representing a position error along the
missile path. These components are shown in FIG. 3 with respect to
the missile axes p, y and h.sub.m, (drawn with an origin at point
c, the present position of the missile.)
E.sub.t, the predicted position error along the missile path is a
measure of the adjustment which must be made in the time of flight
(or the velocity) of the missile to cause interception of the
target. If it be assumed that the velocity of the missile is not
subject to external control once the missile is launched, this
correction must be made by varying the time of flight originally
assumed at the outset of the computation process and employed as
one input to each of predictors 22 and 30. This quantity is,
therefore, applied to a time of flight servo mechanism 60 and
controls the setting of input quantities to the two predictors
possibly as a shaft rotation as in the predictors shown in U.S.
Pat. No. 2,408,081, referred to above. The quantity controlling
predictor 22 is applied directly thereto. However, the
corresponding quantity for predictor 30 which is associated with
the missile section of the computer is modified in accordance with
the ballistic characteristics of the missile before application to
predictor 30. Such modifications are accomplished by apparatus 62
wherein appropriate changes are made in the value of time of
flight. These changes are ordinarily accomplished by adding both
fixed and variable components to that corresponding to the time of
flight as shown for example in FIG. 8A of U.S. Pat. No. 2,408,081
to which reference has been made above.
It will be understood from the above that the necessary functions
for the continuous prediction process performed by the computer are
provided by way of the time of flight servo mechanism and the
target and missile tracking radars. The pitch and yaw error outputs
E.sub.y and E.sub.p from coordinate resolver 58 are employed for
the generation of steering orders for the missile. These quantities
depend of course upon the time of flight fed back to predictors 22
and 30 as discussed above. As a matter of convenience in control of
the missile it has been found desirable to convert these position
orders into acceleration orders, i.e., the quantities representing
the position errors measured along the yaw and pitch axes are
converted into lateral accelerations in the pitch and yaw planes
respectively such that the missile will be at the point of
predicted interception at the predicted time of interception. For
the generation of such orders the quantities E.sub.y and E.sub.p
are applied respectively to dividers 70 and 72 in which each is
divided twice by the time of flight produced as the output of unit
66. The yaw and pitch orders appearing at the outputs of dividers
70 and 72 respectively are thus proportional to the accelerations
required to cause the missile to reach the predicted point of
interception at the predicted time of interception. These orders
may be transmitted to the missile by any convenient means, for
example, by a high frequency radio communication channel.
Alternatively and as shown in FIG. 1 the acceleration orders for
the missile are transmitted by modulation of the repetition rate of
the missile tracking pulse transmitter. The necessary control
quantities may be transmitted on a time division basis or any other
convenient basis by the action of a modulator 74 associated with
transmitter 24. Various systems of signaling over the radar beam
are described in Section 11.2 of "Radar Beacons," Vol. 3 of the
"Radiation Laboratories Series". According to one such system the
two control signals are transmitted as audio frequency signals
frequency modulated upon the pulses from the track transmitter.
Either one or both of the frequencies can thus be transmitted
depending upon the steering orders required at a particular time,
the modulating wave comprising either one or the sum of the audio
frequencies.
Also transmitted to the missile and conveniently by interruption of
all modulation upon the radar beam is the so-called burst order
which at a time related to the predicted time of intercept causes
the warhead of the missile to explode. Ordinarily this order is
transmitted a few microseconds prior to the time when T=0.
The remaining equipment carried aboard the missile may now be
considered. As has been stated above the missile carries a
transponder responsive to pulser from the ground base missile
tracking equipment. This transponder includes a receiver 68 tuned
to frequency f.sub.1 associated with antenna 70, and a microwave
pulse transmitter 72 which is triggered by the output of receiver
68 and which radiates pulses of frequency f.sub.2 from antenna 74.
These pulses when received at the location of the guidance
equipment permit tracking of the missile.
Radio receiver 68 also serves to receive the various orders
transmitted from the guidance equipment and intended to control the
steering fins of the missile and the bursting of the warhead at
appropriate times. Depending upon the nature of the modulation
employed to transmit these orders this portion of receiver 68 may
include a frequency modulation demodulator, pulse position
demodulators or decoders or a frequency division multiplex receiver
wherein the various order channels are distinguished upon a
frequency basis. In any event the receiver is designed with
reference to the particular transmitter 24, FIG. 1, employed in the
guidance equipment and produces three output signals corresponding
respectively to the pitch and yaw acceleration orders and the
warhead burst order.
As shown in FIG. 2 each pair of missile steering fins is driven by
an electric motor, in response to the appropriate orders occurring
at the output of receiver 68. Thus a motor 76 is geared to the
shaft 44 upon which steering fins 40 are mounted and a motor 80 is
geared to the corresponding shaft 46 upon which steering fins 42
are mounted. It will be recalled that the missile is to be made
responsive to acceleration orders. Accordingly for each set of
steering fins the appropriate order appearing at the output of
radio receiver 68 is applied to an amplifier to which is also
applied an output of an accelerometer. These two quantities are
applied in opposition and the motor is driven from the output of
the amplifier until the accelerometer indicates that the desired
angular acceleration has been introduced. When such a condition
occurs the output of the amplifier is reduced to zero and the motor
stops. If the angular acceleration increases, the output of the
accelerometer exceeds the order output of the receiver and the
motor is driven in the appropriate direction to return the
acceleration to the required value. The yaw steering fins 40 are
thus controlled by the output of the comparison amplifier 84 to
which is applied the output of an accelerometer 86 oriented in the
missile to indicate accelerations about the yaw axis. Similarly the
pitch steering fins 42 are controlled by a comparison amplifier 88
to which is applied the pitch order output of the receiver and the
output of an accelerometer 90 oriented in the missile to detect
accelerations about the pitch axis.
The remaining equipment in the missile includes a warhead 92
furnished with an appropriate detonator which may be actuated by an
electric impulse known as the burst order received from the third
output of receiver 68 and applied to the warhead over lead 94.
In the operation of the missile system as thus far described the
missile is launched and guided toward interception with the target
in the same manner as that disclosed in the copending application,
Ser. No. 449,396 referred to above. However, and in accordance with
the present invention the missile tracking function is transferred
from the missile tracking radar 12 to the target tracking radar
which is modified as described above so that during the final
phases of an engagement, referred to above in some instances as the
end game, both target and missile are tracked in elevation and
azimuth by the same radar, in this instance target tracking radar
10. Under these circumstances those data as to position of both
target and missile which are most sensitive to error are referred
to the same reference system and such matters as parallax between
the two tracking radars and the accurate bore-sighting of both as
required to insure appropriate corrections are substantially
eliminated. Conveniently the time in which transfer of control from
the missile tracking radar to the target tracking radar is effected
is determined by the reduction of the angular separation of the
tracking beams of target tracking radar 10 and missile tracking
radar 12 to a predetermined small amount. Although this criteria
for switching is considered the most desirable because of the fact
that in the normal engagement the missile and target tend to be in
approximate alignment in the final phases of the engagement, other
criteria for switching may well be employed. For example, switching
may be effected when the time to go to predicted interception has
been reduced to a predetermined value. In any event, switching
should not occur until the missile has approached sufficiently near
to the target to be illuminated by the beam of the target tracking
radar.
In any event information as to the reduction of the switching
criteria to a predetermined value is transmitted over leads 116 and
118 from target tracking radar 10 and missile tracking radar 12,
respectively, to a control circuit 120. When the predetermined
criteria has been realized control circuit 120 becomes effective to
operate actuator 114 switching the input of predictor 30 from the
output of parallax unit 31 associated with lead 38 from missile
tracking radar 10 to lead 108 from receiver 104 of the target
tracking radar. It will be noted that this transfer eliminates the
parallax correction unit 31 from the elevation and aximuth inputs
to predictor 30 since such correction is no longer required. From
this time onwards during the engagement both target and missile are
illuminated by the beam from target tracking radar 10 and the
automatic tracking equipment of this radar maintains it in such
orientation as to keep both target and missile within the beam. The
beacon signals from missile 16 may thus be received by receiver 104
associated with the target tracking radar and from these received
signals may be derived error quantities representative of the
failure of the target tracking antenna to track the missile exactly
in elevation and azimuth. In the conventional tracking radar, these
error signals are employed as inputs to servo devices arranged to
orient the antenna and adjust the range unit in such a way as to
reduce the errors in tracking effectively to zero. Here, however,
antenna orientation and range unit adjustment are and remain under
control of the target tracking radar and are effected in response
to the error signals from receiver 104. The error signals
.DELTA.A.sub.m, and .DELTA.E.sub.m, representing the azimuth and
elevation errors in the tracking of the missile by the target
tracking radar are therefore applied to an adder 110. This device
is of conventional type and includes three channels in which
.DELTA.A.sub.m and .DELTA.E.sub.m are algebraically added to
A.sub.T, and E.sub.T respectively as derived from servo unit 106.
The output of adder 110 as applied to transmitting channel 108
includes separate quantities representing A.sub.m and E.sub.m of
the missile position data as required by predictor 30 of the
computer. Missile range data for application to predictor 30
continues to be derived from the range unit of missile tracking
radar 12 and is applied by way of parallax correction unit 31 as
noted previously. These quantities representing angular position
are applied to predictor 30 in preference to those normally
supplied by missile tracking radar 12 and the computer treats them
in the same manner as though they were derived from missile
tracking radar 12 and the remainder of the system operates exactly
as previously described. The single reference for both missile and
target position data is thus achieved for the azimuth and elevation
quantities and the need for accurate parallax and bore-sight
adjustments is eased during that phase of the engagement for which
these quantities become of the greatest importance.
The control system by means of which the approach of angular
separation of the two tracking beams to within the predetermined
small quantity, for example 1.degree. in both elevation and
azimuth, may be detected will now be described with reference to
FIG. 4 of the drawings. Here the antenna platforms 122 and 124 of
target and missile tracking radars respectively are shown
schematically. These units may include any combination of synchros,
converters, motors, etc., as required according to well known
principles, to provide an output which is the differential between
two inputs. The exact nature of the units will be dependent upon
the nature of the inputs and outputs which may be electrical
quantities or mechanical shaft rotations as indicated in FIG. 4.
Each of these platforms will be understood to provide for rotation
of the antenna proper in elevation and azimuth. In the control
circuitry now to be considered the azimuth shaft of missile
tracking radar platform 124 is connected to a synchro transmitter
126 while the azimuth shaft is connected to a similar transmitter
128. In a similar manner the elevation shaft of target tracking
radar platform 122 is connected as one input to a differential
synchro unit 130 and the azimuth shaft is connected as one input to
a differential synchro unit 132. For ease of illustration, the
various components such as synchros are shown adjacent to the
antenna platforms to which they are linked. It will be understood
that with the exception of antenna platforms 122 and 124, all of
the elements shown in FIG. 4 are a part of the control circuit 120
shown schematically in FIG. 1. The electrical output of elevation
transmitter 126 is applied as the second input as the differential
synchro 130 and the electrical output of azimuth 128 is applied as
the second input to differential synchro 132. The output shaft of
differential synchro 130 is mechanically connected to a switch 134
and the corresponding shaft of differential synchro 132 is
connected to a switch 136. Switches 132 and 136 are simple single
pole switches through which may be completed a series circuit
including a source of potential such as a battery 138 and actuator
114 of FIG. 1.
The mechanical output of each of differential synchros 130 and 132
constitutes a measure of the angular separation about the
associated axes of radar platforms 122 and 124 thus switch 134 is
closed whenever the two elevation shafts approach within 1.degree.
of angular separation and switch 136 is closed whenever the azimuth
shafts approach within the same limit. When both switches are
closed the requirement for switching of the control from the
missile tracking radar to the target tracking radar is made and
actuator 114 closes transfer switch 112 and the required transfer
of control is effected.
* * * * *