Optics: measuring and testing – By light interference – Rotation rate
Reexamination Certificate
2000-01-24
2002-04-23
Turner, Samuel A. (Department: 2877)
Optics: measuring and testing
By light interference
Rotation rate
C244S165000
Reexamination Certificate
active
06377352
ABSTRACT:
BACKGROUND
1. Field of the Invention
The present invention relates to an apparatus that senses angular rate and provides reaction torque, and more particularly to an apparatus that senses angular rate and provides reaction torque about the same axis.
2. Background of the Invention
FIG. 1
shows a conventional fiber optic gyroscope
10
. The fiber optic gyroscope includes a light source
12
, a coupler
14
, a polarizer
16
(and sometimes one or more depolarizers), a beam splitter
18
, a coil of optical fiber
20
, and a detector
22
. Light from the light source
12
is split by the beam splitter
18
into two phase matched waves that are fed into opposite ends of the coil of optical fiber
20
.
Electronics
28
at the detector measure the phase relationship between the two counter-rotating light waves by examining the interference pattern generated at the confluence. The difference in phase shifts experienced by the two beams provides a measure of the rate of rotation of the platform to which the instrument is fixed. For this reason, a fiber optic gyroscope is considered a rate gyroscope.
Fiber optic gyroscopes can be used to sense the rate of rotation of whatever object to which they are attached. These objects can range from underwater vehicles to land and air based vehicles and finally assorted spacecraft, including satellites, deep space probes, and space stations like Skylab and the International Space Station Freedom. However, a fiber optic rate gyroscope will only sense rotations about a single axis that is perpendicular to the plane of the fiber coil and intersects the center of the circle described by the fiber coil.
Typically, satellites require three orthogonal axes of rate knowledge in order to determine how their attitude, or orientation, is changing with time. For this reason, at least three rate gyroscopes are usually packaged together within a single housing.
FIG. 2
shows an example of three rate gyroscopes
10
being packaged into a single housing.
If the sensing axes of the gyroscopes
10
are appropriately skewed in three dimensional space, the rates that they sense can be resolved into any arbitrary set of three orthogonal vectors which fully define the variation in the attitude of the spacecraft. This package of at least three appropriately skewed rate gyroscopes is called a rate sensor assembly, or a rate gyro assembly.
While a conventional rate sensor assembly, such as the one shown in
FIG. 2
, allows for the measurement of an object's angular rate, the device, and as well the gyroscopes contained within the device, still has significant drawbacks. Each of these drawbacks are discussed below.
Even though the output rates from a rate sensor assembly can be resolved into a triad of orthogonal vectors that fully define the attitude rate of the spacecraft, the accuracy of this data is directly correlated to the accuracy with which the fiber optic rate gyroscopes are aligned with respect to each other. This requirement causes significant expense at the rate sensor assembly's design and test/integration phases.
As an additional complication, this triad of orthogonal vectors can have an arbitrary orientation in inertial space. Typically, the spacecraft designer needs to align this triad of orthogonal vectors to axes on the spacecraft that have important attitude characteristics or requirements. These axes may be related to payload equipment on the spacecraft such as telescopes and antennas, or may be related to guidance and control equipment such as star trackers and reaction and momentum wheel assemblies, or may be aligned along imaginary axes where the spacecraft designer believes the minimum and maximum moments of inertia of the vehicle exist.
In any case, the alignment of the rate sensor assembly to the spacecraft must be known with extreme accuracy so that errors in attitude do not occur. This entails substantial expense at the spacecraft's design and test/integration phases.
Another problem with the rate sensor assembly shown in
FIG. 2
is that the space environment in which the gyroscopes are used pose many difficulties for all types of equipment. Highly sensitive solid state devices like a fiber optic gyroscope can be affected or failed by high energy radiation that is present in space, but is filtered out by the atmosphere before it reaches the earth's surface.
Protection from radiation is dependent upon the mass of the material between the high-energy radiative particle and the circuitry of the fiber optic gyroscope. In other words, protection requires a barrier to exist between the radiation environment and the gyroscope.
The sufficiency of the barrier is related to thickness and the density of its composition. This weight is nonfunctional and, thus, the penalty imposed by this requirement reduces the weight allowable to the useful payload of the spacecraft. However, removal of the barrier would severely limit the life of the fiber optic gyroscope and, subsequently, the life of the spacecraft.
Another difficulty is that fiber optic gyroscopes
10
shown in
FIGS. 1 and 2
are also sensitive to electrical supply voltage and current variations. For this reason, a rate sensor assembly consisting of a set of skewed fiber optic rate gyroscopes must be attached to electrical supply circuits on the spacecraft that are stable in voltage and noise free.
This requirement adds complexity and cost to the fiber optic gyroscope implementation. One option to eliminate this electrical supply sensitivity would be to integrate power conditioning and regulating circuitry into the rate sensor assembly. However, this would simply move the complexity and cost penalty from the spacecraft bus level down to the subsystem level. In either case, this characteristic is a detriment to the attitude determination system design of a vehicle.
And still another problem with the conventional fiber optic gyroscope
10
shown in
FIGS. 1 and 2
is that the gyroscope's performance is directly proportional to the diameter of the coil of optical fiber. However, as the coil diameter increases, so must the housing that contains the set of gyroscopes grow larger.
As the housing of the rate sensor assembly becomes larger, so does the size and area of the mounting footprint. Optimally, the fiber optic rate gyroscopes are aligned orthogonally within the rate sensor assembly's housing. Because of this, the mounting footprint of the rate sensor assembly increases numerically as the square of the increase in the individual fiber optic rate gyroscope's diameter. This is a significant tradeoff, as spacecraft are designed to be as compact as possible and the mounting footprints of all componentry must be minimized.
Additionally, the volume displaced by the rate sensor assembly shown in
FIG. 2
increases numerically as the cube of the increase in the individual fiber optic rate gyroscope's diameter. This is a very strong impediment to the use of large fiber optic rate gyroscopes on spacecraft and is why most sensors have maximum diameters no more than six inches. However, the miniaturization of a fiber optic gyroscope significantly reduces performance unless expensive complications are applied to the basic gyroscope design.
FIG. 3
shows a conventional reaction and momentum wheel assembly
50
. The assembly
50
includes a flywheel
52
and electric motor assembly
54
, among other items, that are used to provide reaction torques to a vehicle and store momentum.
In the mechanism, the flywheel
52
is mounted on rotational bearings
56
, which may be ball, roller, or magnetic, that are secured to a housing
58
that is attached to the vehicle's structure.
The flywheel
52
is rotated by the electric motor assembly
54
that has a stator (nonrotating) attached to the housing, and a rotor (rotating) attached to the flywheel. Applying electricity to the motor
54
causes a torque to be developed within it and the flywheel begins to rotate. If a constant torque is generated, the flywheel accelerates at a constant rate, and the angular speed increases lin
Coronato Patrick
Pona Joseph
Honeywell International , Inc.
Turner Samuel A.
Yeadon Loria B.
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