U.S. patent number 3,614,027 [Application Number 04/839,829] was granted by the patent office on 1971-10-19 for pneumatic resolver for missile control.
This patent grant is currently assigned to The United States of America as represented by the Secretary of the Army. Invention is credited to Charles Lynn Lewis.
United States Patent |
3,614,027 |
Lewis |
October 19, 1971 |
PNEUMATIC RESOLVER FOR MISSILE CONTROL
Abstract
A pneumatic attitude control system for an artillery-type
missile which is pin stabilized during flight. The control system
includes a sensor portion which is roll stabilized and, thus, will
not roll with the missile. Thus, attitude angles measured by the
roll stabilized sensors will be space fixed. Control valves for
providing restoring torques on the missile are mounted on the
rotating part of the missile for rotation therewith. A resolution
system is provided for resolving the space-fixed signals from the
roll-stabilized sensors into the rolling missile's coordinate
system for desired operation of the control valves.
Inventors: |
Lewis; Charles Lynn
(Huntsville, AL) |
Assignee: |
The United States of America as
represented by the Secretary of the Army (N/A)
|
Family
ID: |
25280721 |
Appl.
No.: |
04/839,829 |
Filed: |
July 8, 1969 |
Current U.S.
Class: |
244/3.22;
137/805 |
Current CPC
Class: |
F42B
10/663 (20130101); Y10T 137/2071 (20150401) |
Current International
Class: |
F15D
1/02 (20060101); F15D 1/00 (20060101); G06F
1/00 (20060101); F42b 015/02 (); G06f 001/00 ();
F15d 001/02 () |
Field of
Search: |
;244/3.22 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Borchelt; Benjamin A.
Assistant Examiner: Webb; Thomas H.
Claims
I claim:
1. An attitude control system for a missile having spinning and
nonspinning portions, said attitude control system including:
a. sensing means carried in said nonspinning portion and disposed
for sensing the attitude of said missile and generating space-fixed
signals indicative thereof;
b. resolver means disposed for resolving said attitude signals into
the coordinate system of said spinning portion;
c. control means disposed for receiving said resolved signals for
utilization thereof to provide resorting torques on said
missile.
2. An attitude control system as set forth in claim 1 including
pulse duration modulation means connected between said resolver
means and said control means to transform said signals into pulses
for actuation of said control means.
3. An attitude control system as set forth in claim 2 wherein said
sensing means includes a pneumatic gyro disposed for generating
signals indicative of missile attitude.
4. An attitude control system as set forth in claim 3 wherein said
resolver means includes:
a. a plurality of fluid amplifiers connected to said gyro to
receive input signals therefrom; each of said amplifiers having a
pair of output channels;
b. cam means carried on said spinning portion of said missile for
coaction with the outputs of predetermined ones of said fluid
amplifiers for generating back pressures therein, said pressures
varying in response to rotation of said spinning portion.
5. An attitude control system as set forth in claim 4 wherein said
cam means includes a pair of tracks having surfaces of varying
depths, said surfaces varying in a sinusoidal manner.
6. An attitude control system as set forth in claim 5 wherein a
first plurality of fluid amplifiers are disposed for providing
signals for control of said missile in a first attitude plane, and
a second plurality of fluid amplifiers are disposed for providing
signals for attitude control in a second plane, said first
plurality of fluid amplifiers having predetermined ones thereof
provided with outputs disposed adjacent the first of said cam
tracks, said second plurality of fluid amplifiers having
predetermined ones thereof provided with outputs disposed adjacent
the second of said cam tracks.
7. An attitude control system as set forth in claim 6 wherein said
first and second plurality of amplifiers includes:
a. pickoff means connected to the output of said predetermined ones
of said fluid amplifiers intermediate said cam and the body of said
fluid amplifiers; and,
b. a plurality of fluid summing amplifiers having said pickoff
means connected across the inlet thereof;
c. an output summing amplifier having the outputs of said summing
amplifier connected across the inlet thereof.
8. An attitude control system as set forth in claim 7 wherein said
pulse duration modulation means is connected to the outputs of said
output summing amplifiers to receive signals therefrom.
9. An attitude control system as set forth in claim 8 including
fluidic slipring means connected between said spinning and
nonspinning portions of said missile to transfer signals from said
pulse duration modulation means to said control means.
Description
SUMMARY OF THE INVENTION
A pneumatic control system for a missile having a portion thereof
which is spun prior to launch and a nonspinning (roll-stabilized)
portion. Tubes are mounted on the roll-stabilized portion for
carrying pneumatic signals therefrom representing space-fixed
attitude angles. The signals are obtained from a pneumatic gyro and
carried by the tubes to exhaust perpendicularly onto a cam or plate
having concentric grooves, the depths of which vary in a sinusoidal
manner around the circumference. The cam is mounted to a rotating
portion of the missile. Thus, the distance between the ends of the
tube and the surfaces of the plate varies as the plate rotates
relative to the ends of the tubes. The variable depths of the
grooves change the back pressure in each tube and thereby change
the pressure in a pickoff line upstream of the exhaust point, or
end of the tubes. The pickoff pressure is proportional to the
distance between the cam surface and the tube and will change the
pressures in the pickoff lines. The shape of the cam provides for
the occurrence of pressure variations in a sinusoidal manner.
The outputs of the back pressure pickoffs are summed with necessary
bias signals to provide analog signals which provide the required
resolution of the space-fixed attitude angles into rolling
coordinates.
It is, therefore, an object of the present invention to provide a
pneumatically controlled attitude control system for a missile.
It is another object of the present invention to provide a missile
attitude control system having a sensor portion thereof which is
space fixed and a control portion thereof which is rotating during
missile flight.
It is a further object of this invention to provide such a control
apparatus with means to resolve signals from the space-fixed sensor
into the missiles rolling coordinate system for utilization by the
control portion of the system.
These and other objects of the present invention will be more
readily apparent from the following detailed description and
accompanying drawings .
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a diagrammatic representation of the control system of
the present invention.
FIG. 2 is a diagrammatic representation of the control system
including the attitude angle resolver system as shown in FIG.
1.
FIG. 3 is a sectional view along line 3--3 of FIG. 1 showing the
resolver cam and slipring arrangement with the tubes entering the
slipring assembly removed for clarity.
FIG. 4 is a diagrammatic view of the space-fixed and rolling body
coordinate systems.
FIG. 5 is a diagrammatic view of mechanism for coordinate
transformation of the space-fixed attitude angle in a first
attitude plane.
FIG. 6 is a diagrammatic view of mechanism for coordinate
transformation of the space-fixed attitude angle in a second
attitude plane.
FIG. 7 is a graphical representation illustrating pickoff pressure
variation versus plate rotation.
FIG. 8 is a graphical representation illustrating output pressure
versus plate rotation.
FIG. 9 is a pictorial view illustrating diagrammatically, the
missiles roll-stabilized and rotating portions and the slipring
connection therebetween.
DESCRIPTION OF THE PREFERRED EMBODIMENT
As shown in FIG. 1, a missile 10 is shown to include a
roll-stabilized portion 12 and a rotating portion 14 which are
secured together for relative rotation therebetween by means well
known in the art. A control system 16 carried in the missile
includes a rotating control portion 18 and a space-fixed sensor
portion 20. Sensor portion 20 is shown to include a plurality of
tubes 22-29 which are disposed in communication with a rotating cam
15 which is carried by, and disposed for rotation with rotating
portion 14 of the missile. Cam 15 is provided with a pair of
concentric surfaces 11 and 13 (see FIG. 3) of varying depth. Tubes
22-29 communicate between a pair of attitude angle resolver systems
31 and 33 and cam 15 for generation of signals which are resolved
into the missile's rolling coordinate system by resolver systems 31
and 33 and transferred to a pulse duration modulation mechanism 17
and through a fluidic slipring 19 to a plurality of control valves
21 mounted on the periphery of the missile. Input signals to
resolver systems 31 and 33 are received from a pneumatic gyro
9.
FIG. 2 is a block diagram of the control system of FIG. 1 and
illustrates the mechanism of the attitude angle resolver systems as
generally shown in FIG. 1. As shown in FIG. 2, the attitude angle
resolver system includes a plurality of outputs 30-37 of a
plurality of center dump amplifiers 40-47, respectively, which
communicate, respectively, with tubes 48-55 are connected to tubes
22-29 intermediate the input center dump amplifiers 40-47 and
rotating cam 15. Tubes 48, 49, 50 and 51 communicate respectively,
with tubes 22, 23, 24 and 25 and a plurality of summing amplifiers
58, 59, and 60, respectively Lines 52, 53, 54, and 55 connect,
respectively intermediate lines 26, 27, 28 and 29 and a second
plurality of summing amplifiers 61, 62, and 63.
Amplifiers 41, 43, 45 and 47, are provided with a second plurality
of output tubes (FIGS. 2, 5 and 6) 68-70, 72-74, 76-78, and 80-82,
respectively, (FIGS. 5 and 6) which are provided with an attenuator
or orifice 84 to reduce pressure in these tubes by three-fourths.
This serves to provide the input sine wave with a polarity. The
outputs from the bank of output summing amplifiers (FIG. 2) are
transmitted through tubes 86, 88, 90 and 92 to a pulse duration
modulation mechanism 17 to produce a pulse duration modulation
signal which is transferred through "slipring" type mechanism 19
(FIG. 9). The transferred control signals are then used to drive
the plurality of control valves 21 carried around the periphery of
the rotating portion of the missile.
As more clearly seen in FIG. 5, the mechanism for transformation of
space-fixed signals for control in one plane .GAMMA..theta. is
illustrated, assuming this to be the missle's body-fixed pitch
control plane. In this control mechanism, a first pair of input
amplifiers 40 and 41 are connected to input sources 100 and 102
which provide a control signal across the power jet flowing through
entrance 104, which is received from a missile carried source, not
shown, of amplifiers 40 and 41. The jet is disposed for flow
through tubes 24 and 25 of amplifier 40 and through tubes 68 and 70
of amplifier 41. Tubes 24 and 25 terminate adjacent the cam surface
15. However, pickoffs 50 and 51 connect into tubes 24 and 25,
respectively, upstream of the ends of the tubes which lie adjacent
cam 15 and connect to the power jet entrance 104 of a summing
amplifier 58. Tubes 68 and 70 of amplifier 41 connect adjacent
power jet inlet 104 of amplifier 58, as shown in FIG. 5.
Outputs 106 and 108 of amplifier 58 connect across opposite sides
of entrance 104 of another output summing amplifier 60 which is
provided with output tubes 90 and 92 which connect to the pulse
duration modulation mechanism 94 (FIGS. 1 and 2).
In like manner, input amplifier 42, as shown in FIG. 5, is
connected to inputs 110 and 112 which provide a control signal
across the power jet flowing through entrance 104 of amplifiers 42
and 44. The jet is disposed for flow through tubes 22 and 23 of
amplifier 42 and through tubes 72 and 74 of amplifier 43. Tubes 22
and 23 terminate adjacent cam surface 15. However, pickoffs 48 and
49 connect into tubes 22 and 23 upstream of the ends of the tubes
which lie adjacent cam 15 and these pickoffs connect adjacent the
entrance 104 of a summing amplifier 59. Outputs 110 and 112 of
summing amplifier 59 connect across entrance 104 of output summing
amplifier 60.
Similarly, to provide control in the plane .GAMMA..psi., assuming
this to be the missile's body-fixed yaw control plane attitude in
yaw, the control mechanism as illustrated in FIG. 6 is utilized. In
this control mechanism the pair of input center dump amplifiers 44
and 45 are connected to input sources 114 and 116 which provide a
signal across the power jet flowing through the power jet inlet 104
of amplifiers 44 and 45. The jet is disposed for flow through tubes
26 and 27 of amplifier 44. Tubes 26 and 27 terminate adjacent cam
surface 15. However, pickoffs 52 and 53 connect into tubes 26 and
27 upstream of the ends of the tubes which lie adjacent cam 15 and
these pickoffs connect adjacent the power jet inlet 104 of output
summing amplifier 61. Outputs 118 and 120 of summing amplifier 61
connect across inlet 104 of an output summing amplifier 63 which is
provided with a pair of output tubes 90 and 92 which connect into
pulse duration modulation mechanism 17.
In like manner, input amplifier 46, as shown in FIG. 6, is
connected to inputs 122 and 124 which provides a control signal
across the power jet flowing through inlet 104 of amplifiers 46 and
47. The jet is disposed for flow through tubes 28 and 29 of
amplifier 46 through tubes 80 and 82 of amplifier 47. Tubes 28 and
29 terminate adjacent cam surface 15. However, pickoffs 54 and 55
connect into tubes 28 and 29 upstream of the ends of the tubes
which lie adjacent cam 15 and these pickoffs connect adjacent the
entrance 104 of output summing amplifier 62. Outputs 126 and 128 of
output summing amplifier 62 connects across inlet 104 of output
summing amplifier 63.
The resolution required in the control system of the present
invention is illustrated in FIG. 4, where the solid axis represent
the space-fixed coordinate system and the dotted axes represent the
body-fixed coordinate system. The angle through which the missile
with roll frequency .omega. has rolled in time t is denoted
.omega.t.
Assuming .theta. is the missile's space-fixed attitude angle in
pitch and .psi. is the space-fixed attitude angle in yaw, then the
rolling (body-fixed) signals are seen to be sin-cos functions of
these angles and are given by:
.GAMMA..theta.=.theta. cos .omega.t+.psi. sin .omega.t
.GAMMA..psi.=.psi. cos .omega.t-.theta. sin .omega.t.
To mechanize this pneumatic resolution, tubes which are mounted on
the roll-stabilized portion of the control system and which carry
analog (smoothly varying) pneumatic signals representing the
space-fixed attitude angles .theta. and .psi. are positioned so
they exhaust perpendicularly into concentric grooves cut into a
plate that is affixed to the rolling portion of the missile. This
is shown in FIG. 2. The depth of the grooves cut into the plate
varies in a sinusoidal manner around the circumference.
From fluid flow theory, it is known that the back pressure in the
tube (which is picked off upstream of the exhaust point) is
proportional to the distance x between the tube end and the bottom
of the groove. The variable depth of the groove will change the
pressure in the pickoff lines as follows. Assume that in FIG. 5,
which shows the mechanization of one of the resolution equations,
line 29 has a constant freeflow pressure of X p.s.i. When the line
end is located at the 0.degree. position of the plate, the pressure
at pickoff 55 will also be X p.s.i. This indicates that the
distance between groove bottom and tube end at the 0.degree.
location should be zero. As the plate rotates until the 90.degree.
position is under line 29, the pickoff pressure drops to 3X/4
p.s.i. Then, the pressure at the pickoff continues to drop as the
plate rotates toward 180.degree.. At 180.degree., the pressure at
the pickoff is X/2 p.s.i. The pressure then begins to build up
through the next 180.degree. of motion until it is again X p.s.i.
after one complete cycle of the plate. This ratio of decrease and
increase of the pressure from its original value during a cycle of
the plate will hold for all of the lines impinging upon the plate.
The depth of the grooves is to be cut such that the above pressure
variations will occur in a sinusoidal manner. This is shown in FIG.
7.
By locating the impingement points (points where tubes exhaust into
the grooves) in a certain manner, as shown in FIGS. 5 and 6, and
summing the outputs of the back pressure pickoffs with necessary
bias signals, analog signals can be generated which are the
required resolution of the space-fixed attitude angles .theta. and
.psi. into coordinates. The summation of the pickoff signals and
the bias signals can be performed by any commercial pneumatic
summing device.
One representative cycle of the pneumatic resolver would go as
follows: First, it is assumed that with no input signal (P.sub.1
=P.sub.2 , P.sub.1 =P.sub.2 ), there will be no pressure in the
lines that exhaust into the grooves, and thus the output will
correctly remain zero (P.sub.1 =P.sub.2) regardless of the
orientation of the plate. This is accomplished by the use of the
center dump input amplifiers. For purposes of explaining the
operation of the system, in FIG. 5, let P.sub.1 =P.sub.2 =Y p.s.i.
for no input. The actual magnitude is not critical to the operation
of the resolver. Next, assume the input is a positive pitch
attitude angle .theta. which causes a differential pressure of
P.sub.1 -P.sub.2 =+X p.s.i. across the center dump input
amplifiers. In addition, assume the yaw input is zero (P.sub.1
-P.sub.2 =0). The sign of the input is given by the input
amplifiers. The sign changes inherent in the sin-cos terms as the
plate rotates
+ in 1, 2 quadrant Sin .omega.t= - in 3, 4 quadrant
+ in 1, 4 quadrant Cos .omega.t= - in 2, 3 quadrant
is taken care of by the bias, which is set to the value of the
pickoff pressure at the 90.degree. and 270.degree. points. This
value, as previously defined is 3X/4 p.s.i., where X is the input
pressure.
The change in the output pressure P.sub.1 and P.sub.2 of FIG. 5 as
the plate rotates through one cycle is shown in FIG. 8. Examination
of the .GAMMA..theta. equation (with .theta.=positive quantity,
.psi.=zero) indicates that .GAMMA..theta.=positive value (P.sub.1
>P.sub.2) for 0.degree. to 90.degree., zero at 90.degree.
(P.sub.1 =P.sub.2), negative for 90.degree. to 270.degree. (P.sub.2
>P.sub.1), zero at 270.degree., and positive for 270.degree. to
360.degree. (P.sub.1 >P.sub.2). This is precisely what FIG. 8
shows the pneumatic resolver does. The exact magnitude of the
output differential pressure as compared with the input
differential pressure magnitude will depend upon the gain set in
the pneumatic amplifiers and summers.
The function in the pneumatic resolver is to generate the
.GAMMA..theta. and .GAMMA..psi. signals from .theta. and .psi.
signals. Once these signals are obtained, they may be summed with a
high-frequency carrier signal to produce a pulse duration
modulation control signal. This pneumatic signal, now in the form
of a pulse train, is then transferred physically in a "slipring"
fashion from the roll-stabilized portion of the missile to the
rolling portion via a fluidic switching slipring 19, as shown in
FIG. 9.
The fluidic slipring is utilized to transfer the PDM control pulse
train from the roll-stabilized portion to the rolling missile
section of operation of the control valves. The axle shaft 129 of
the roll-stabilized section is mounted in bearings 130 as shown in
FIG. 9 to an opening 131 of the rolling portion of the missile. The
axle shaft has a plurality of grooves 132 cut circumferentially
therearound which are fed by the outputs 133, 134, 135, 136, of the
PDM control mechanism 17. Tubes 137, 138, 139 and 140 mounted in
the rolling portion of the missile are disposed for communication
with the grooves and are disposed to pick up pressure in the
grooves of the shaft, and transfer these pressures to control
valves 21.
It is to be understood that the cam 15 may be a plate having a pair
of grooves cut therein and provided with surfaces that vary
sinusoidally, or, if desired, the cam may be a member wherein the
cam surfaces protrude above a main surface in a manner to provide
the sinusoidal output from the tubes. Additionally, the gyro for 9
referred to in this application may be any of the many commercially
available type pneumatic gyros. One such type gyro, for example,
may be obtained from General Precision, Inc., Singer-Kearfott
Division, Little Falls, New Jersey.
The pulse duration modulation principle as utilized in the present
invention is similar to that disclosed in the patent to Kenneth C.
Evans, entitled "Pure Fluid Amplifier and Pure Fluid Amplifier
Attitude Control System for Missiles" No. 3,278,140, issued Oct.
11, 1966. Additionally, pulse duration modulation principle is
disclosed by Hancock, in "An Introduction to the Principle of
Communication Theory", McGraw-Hill 1961.
* * * * *