CA1313698C - Spacecraft accelerometer auto-alignment - Google Patents
Spacecraft accelerometer auto-alignmentInfo
- Publication number
- CA1313698C CA1313698C CA000543321A CA543321A CA1313698C CA 1313698 C CA1313698 C CA 1313698C CA 000543321 A CA000543321 A CA 000543321A CA 543321 A CA543321 A CA 543321A CA 1313698 C CA1313698 C CA 1313698C
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- accelerometers
- spacecraft
- accelerometer
- pair
- rotor
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
- G01P21/00—Testing or calibrating of apparatus or devices covered by the preceding groups
Abstract
SPACECRAFT ACCELEROMETER AUTO-ALIGNMENT
ABSTRACT OF THE DISCLOSURE
A spin-stablized spacecraft has a pair of three-axis accelerometers (A, C) mounted on the spacecraft rotor (10) to provide for self-aligning, enduring and sophisticated navigational capability. A
second pair of accelerometers (B, D) is added for redundancy and added sensitivity. Each pair of accelerometers are disposed at diametrically opposed positions on the rotor and respective axes co-aligned. The orientation of spacecraft angular momentum ? is determined by star and planet sensors.
Radial alignment of each accelerometer relative to the spacecraft angular momentum is attained by zeroing non-radial accelerometer readings during spacecraft quiescence. Axial alignment is attained during periods of substantially axial thrust by isolating axial translational acceleration by summing readings across an accelerometer pair to cancel rotational, spin-periodic contributions to the readings and time-averaging or otherwise filtering to cancel other non-axial periodic contributions to the readings. Axial alignment is automatically effected upon the correct alignment of the radial and tangential axes of each accelerometer.
ABSTRACT OF THE DISCLOSURE
A spin-stablized spacecraft has a pair of three-axis accelerometers (A, C) mounted on the spacecraft rotor (10) to provide for self-aligning, enduring and sophisticated navigational capability. A
second pair of accelerometers (B, D) is added for redundancy and added sensitivity. Each pair of accelerometers are disposed at diametrically opposed positions on the rotor and respective axes co-aligned. The orientation of spacecraft angular momentum ? is determined by star and planet sensors.
Radial alignment of each accelerometer relative to the spacecraft angular momentum is attained by zeroing non-radial accelerometer readings during spacecraft quiescence. Axial alignment is attained during periods of substantially axial thrust by isolating axial translational acceleration by summing readings across an accelerometer pair to cancel rotational, spin-periodic contributions to the readings and time-averaging or otherwise filtering to cancel other non-axial periodic contributions to the readings. Axial alignment is automatically effected upon the correct alignment of the radial and tangential axes of each accelerometer.
Description
~ 313~98 SPACECRAFT ACCELE~OMETER AUTO-ALIGNMENT
BACKGROUND OF THE INVENTION
The present invention relates to spacecraft navigation, and, more particularly, to a system and method providing for on-board alignrnent of navigational accelerometers.
Spacecraft navigation is facing increased demands.
Formerly, spacecraft navigation involved attaining a desired orbit and making occasional corrections to maintain the orbit. More recently, however, rnultiple-year missions involve, for example, more distant objectives of autonomous orbit change or re-entry. These subsequent rnissions require navigational systems capable of meeting objectives far more complex than maintaining orbit and far more enduring than navigational systen~ designed for attaining orbit.
One challenge faced in developing such an enduring navigational system is maintaining its calibration or alignment with respect to an inertial frame of reference. This is particularly true where it is irnpractical to perform this alignment on a regular basis from a ground or other remote station.
One navigational approach has been ~o mount three-axis accelerometers on a gyro-stabilized platform mounted on a launch vehicle or spacecraft. The gyro-stabilized three-axis accelerometers provide a acceleration history which can be integrated once to obtain a velocity history and integrated twice to obtain a position history for the 3~
1313~98 --2~
spacecraft in a known coordinate system. The histories so obtained are used to attain elaborate navigational objectives. The alignment of the accelerometer is maintained by the characteristics of the gyro. Such an arrangement has been used effectively in guiding the ascent of a satellite to a predetermined orbit.
The mechanical limitations of such an arrangement in the face of launch stresses, environmental extremes in space, mechanical creep, other stresses and uncompensatable random drift rates and biases limit the long-term accuracy of the gyros of such an arrangement.
Accordingly, the accelerometer-on-a-gyro approach has not proved satisfactory when a regenerative mission objective requires elaborate navigation months or years after orbit is first attained. Thus, such systems require remote realignment of accelerometers if such realignment is provided for at all. Other systems require dedicated celestial body trackers physically tied to the gyro platform to achieve the realig~Dnent and calibration of the gyros. Still other systems discuss a "strap-down"
method as speculated in "Theory of Inertial Guidance " by Connie L.
McClure, Prentice Hall 1960, pp. 286-291.
Spinning spacecraft have re enduring orientation aids such as attitude and nutation sensors. However, while these are well-adapted for rnaintaining and adjusting attitude, they are not suited for extensive spacecraft navigation. In other words, they caMot readily yield a substitute for the instantaneous acceleration histories provided by accelerometers during periods of thrust.
What is needed is a system and method for realignment of a navigational inertial coordinate system. The system should provide for self-alignment and operate for extended periods essentially autonomously, without the constraints of navigation gyros. The system should be capable of managing demanding navigational objectives after an extended orbital period.
SUMI\~ARY O~ THE INVENTION
A spin-stabiliæed spacecraft includes a platform and a rotor. A pair of circumferentially spaced three-axis accelerometers located on the rotor provide acceleration histories for navigation. In accordance with the present invention, radial alignment of each accelerometer relative to the spacecraft angular momentum, h (~igure 1), is performed during periods of quiescence and axial alignment is performed during periods of thrust.
Using a pair of circumferentially spaced accelerometers permits mathematical cancellation of certain rotationally-induced i0 acceleration components. Other non-translational periodic components can be cancelled through time-averaging or otherwise filtering individual accelerometer readings. Accordingly, acceleration components of interest, namely centripetal and axial-translational components, can be isolated to permit alignment of the accelerometers.
The particulars of cancellation are dependent on the relative placement and orientation of the two accelerometers. The mathematical computations can be greatly simplified in the case where the accelerometers are located at diametrically opposed positions on the rotor and their respective axes co-oriented. With this arrangement, the accelerometer readings can be simply summed, averaged or otherwise filtered to cancel rotational contributions to individual accelerometer readings.
During a period of quiescence, a spin-stabilized spacecraft is spinning in orbit with negligible non-gravitational acceleration, nutation or precession. In this quiescent state, the axial and tangential acceleration components are negligible. A non-zero tangential or axial acceleration reading by the accelerometers can be treated as an error signal to be zeroed through realignment to a plane perpendicular to the angular momentum vector, h.
13~3~98 During a period of thrust with substantial axial acceleration colrponents, the accelerations at the accelerometer locations can be mathematically combined so as to cancel non-translational, spin periodic accelerometer readings due, for example, to torquing. Other non-axial, periodic (e.g. nutational) conponents can be generally time-averaged to zero or digitally filtered. Thus, resultant non-zero readings can serve as axial and tangential accelerometer ali~nent error signals to be corrected through realignment about the radial direction and perpendicular to the angular momentum, h. Periods of thrust occur during initial ascent, north-south station keeping, orbit adjustment, and, if necessaryJ during maneuvers undertaken specifically for accelerometer alignment.
Redundancy and increased sensitivity can ~e provided by adding a second pair of diametrically opposed accelerometers, each located a quarter turn from each of the accelerometers of the first pair.
Thus, a force which may be small relative to the centripetal force of the spinning rotor and therefore difficult to detect as a radial component of an accelerometer of, for example, the first pair, can be readily detected as a tangential component by an accelerometer of the second pair. Also, if an accelerometer fails, one pair remains for navigational purposes.
In accordance with the foregoing description, a system and method for on-board alignment of an accelerometer-within-a-gyro-based navigational system is provided. The accelerometers are well-established for major navigational objectives and are less prone to long-term errors than gyro instruments. While remote calibration can supplement, the provision of on-board alignment ensures the availability of this autonomous navigation capability months or years after initial orbit is attained, depending primarily on the completeness of the on-board ephemeris-generating program and the accuracy of the sensor-based, time independent on-board attitude determination subsystem.
1~3698 -4a Other aspects of this invention are as follows:
A spacecraft with autonomous navigation accelerometer alignment comprising: a spin-stabilized spacecraft having a platform and a rotor with a spin axis; a pair of three-axis accelerometers circumferentially spaced about said rotor; means for applying force to said spacecraft to induce acceleration with a substantial translational axial component; means for detecting and correcting radial alignment error in each accelerometer when said spacecraft is in a quiescent state; and means for detecting and correcting axial alignment error in each accelerometer when said spacecraft is accelerating with a substantial component in said axial direction.
A method of realigning a three-dimensional reference coordinate system on board a thruster-controlled orbiting spin-stabilized spacecraft having a platform and a rotor which rotates about a spin axis, said spacecraft being subject to dynamic imbalance resulting in misalignment of the reference coordinate system, comprising the steps of:
(a) placing at least one pair of accelerometers on the rotor at respective positions radially spaced from said spin axis and circumferentially spaced with respect to each other, each accelerometer having three sensing axes for respectively 1313~98 -4b-sensinq acceleration of the rotor with respect to said spin axis; (b) aligning each of said accelerometers radially with respect to said spin axis while the spacecraft is not being acted on by a thrust force; and (c) aligning each of said accelerometers axially with respect to said spin axis while the spacecraft is being acted upon by a thrust force.
A spin-stabilized spacecraft comprising: a rotor;
a pair of three-axis accelerometers circumferentially spaced upon and fixed to said rotor; means for obtaining quiescent state readings from each of said accelerometers;
means for determining from said quiescent state readings errors in the radial alignment of each of said accelerometers; means for correcting said errors in radial alignment; means for obtaining thrust state readings from each of said accelerometers; means for mathematically combining and time-averaging the thrust state readings of said accelerometers so as to yield calculated axial translational components: and means for axially aligning each of said accelerometers so that its axial readings match said calculated axial translational components.
-4c-A method of calibrating a navigational system comprising the steps of: mounting a pair of three-axis accelerometers circumferentially spaced upon a rotor:
obtaining quiescent state readings from Pach of said accelerometers; determining from said quiescent state readings errors in radial alignment for each accelerometer;
correcting said errors in radial alignment; reading thrust state histories from each accelerometer; mathematically combining said thrust state histories to cancel rotational spin-periodic contributions to said accelerometer readings and filtering said thrust state histories to cancel other non-axial periodic components of acceleration to yield calculated axial translational components; for each accelerometer, determining the axial orientation which would have resulted in a thrust state history corresponding to said calculated axial translational components and align accordingly.
131'~98 BRIEF DESCRIPTION OF THE DRAWINGS
In the accompanying drawings:
FIG. 1 shows R rotor of a spin-stabilized spacecraft with accelerometers mounted thereon in accordanc~ with the present invention.
S FIG. 2 is a block diagram of a navigation~l system in accordance with the present invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
In accordance with the present invention, the rotor 10 of a spin-stabilized spacecraft has mounted thereon n pair of accelerometers A, C as shown in FIG. 1. The accelerometers A, C are aligned radially during quiescent orbit and axially during periods of thrust. Since the accelerometers are spaced circumferentially, their readings can be mathematically combined to cancel spin-periodic and centripetal acceleration components. Other non-axial periodic (e.g.
nutation) components can be generally time-averaged to zero or otherwise filtered to isolate variables of interest.
For computational simplicity, the accelerometers A, C
can be diametrically opposed, i.e., be spaced 180 about the circumference of the rotor 10 and their respective axes co-oriented.
However, the present invention provides for alternative circumferential spacings and relative orientations. A second pair of accelerometers B, D
can be included for increased sensitivity and redundancy.
While only the rotor l0 and accelerometers A, B, C, D
are illustrated, it is understood that a spin-stabilized spacecraft typically includes a platform upon which most of the mission equipment is placed.
The platform ~s relatively despun, whi1e the rotating rotor 10 supplies 1313~8 directional stability to the spacecraft. Of course, equipment, such as attitude sensors intended to be rotated, can be mounted on the rotor 10.
The translational motion of the spacecraft is indicated by velocity vector v, while the position of the spacec~aft having rotor 10 is identified by the displacement vector Ro of the spacecraft center of mass cm relative to the origin of an inertial frame of reference X, Y, Z.
Passing through the center of mass cm are the principal or geometric axes of the spacecraft x, y, z, the last being the spin axis z of the spacecraft.
The instantaneous spin vector ~ is shown non-coincident with the spin axis z, as would be the case during nutation.
Nutation, or more colloquially, wobble, is generally undesirable and so nutation damping is widely employed. In the absence of nutation and other forms of perturbation, ~.) and z are coincident.
During such periods of quiescence, i.e. while the rotor is spinning, but in the absence of nutation, precession or thrust, the angular momentum h is colinear with the spin axis z and the spin vector ~
However, h can diverge from the spin axis z and the spin vector ~ due to the application of a thrust force F whose direction is offset from the instantaneous center of mass cm.
The angular momentum h is particularly important because its orientation can be very precisely determined usin~ celestial, e.g., star, sun, moon, earth feature sensors mounted on the rotor. For example, a star sensor can determine the spacecraft angular mornentum direction relative to inertial coordinates to within 2 arcseconds. Thus, in accordance with the present invention, the alignrnent of accelerometers in inertial space can be precisely determined by aligning the accelerometers with respect to the angular momentum h of the spacecraft.
Before explaining the alignment of the accelerometers A, B, C and D, their use in a navigational mode is explained with reference to FIG. 2. The star and planet sensors 12 provide attitude and angular momentum data while the accelerometers provide acceleration histories. This data is processed by an on-board data processor 14 to determine orbit changes and attitude. The raw data can also be sent to ground processing stations via a tracking, telemetry and comrand (TT+C) mod ule 16 .
A ground station can program the TT~C 16 module with mission objectives which are compared by the data processor 14 with present orbit and attitude data to determine appropriate commands to the propulsion system 18. The action of the propulsion system 18 is fed back to the on-board processor 14 to obtain a better coordination of propulsion commands and their effects on orbit and attitude. The effect of the propulsion system 18 on the spinning spacecraft dynamics 20 are reflected in the readings of the accelerometer pairs A, C and B, D.
Conveniently, the accelerometers are disposed in diametrically opposed pairs. This will cause the centripetal and spin-period components of acceleration due to rotor rotation rather than spacecraft translation to cancel. Thus, by combining the data of an accelerometer pair and computationally compensating or filtering other (e.g. nutation) periodic components, translational acceleration histories due to thrust forces can be obtained in a straightforward manner. An acceleration history so obtained can be integrated once to obtain a velocity history and twice to obtain a positional history. In this way, each accelerometer pair, A and C, B and D, can provide information sufficient for complex navigation.
In addition to providing redundancy, the second pair B, D of accelerometers provide additional sensitivity. Since the pairs are orthogonally disposed with respect to one another, a force which is radial with respect to one pair is tangential with respect to the other. An orbit-correction radial thrust pulse might be very sma~l in magnitude 1313~9~
compared to ongoing radial forces and thus difficult to detect by an accelerometer positioned radially with respect to the pulse. However, an accelerometer disposed one quarter circumference from the radially positioned accelerometer could assess the pulse with much greater precision. This is because the ongoing tangential force is nominally zero and should not mask the effect of the pulse to the extent that the inherent radial force would.
However, the usefulness of the accelerometers A, B, C, D is premised on their orientations being precisely known with respect to an inertial frame. The invention provides for precise alignment relative to an inertial frame as explained below.
Each accelerometer A, B, C, D is R three-axis accelerometer, comprising three mutually orthogonal one-dimensional accelerometers, a nominally radial axis ar, a nominally tangential axis at, and a nominally axial axis aa. The three axes are reliably and rigidly disposed relative to each other so that it can be assumed that if two are properly aligned, then the third is properly aligned. Thus, the challenge reduces to aligning two of the axes of each accelerometer with respect to the angular momentum h, the direction of which is known in inertial coordinates from the readings of the star and planet sensors 12.
The two axes to be aligned are the radial and the axial axes, with tangential alignment automatically resulting. Radial alignment is effected during spacecraft quiescence and the axial alignment is effected during periods of thrust. Since quiescence is the usual state, it is convenient to consider the radial alignment as occurring first and the axial, and concomitantly, the tangential alignment occurring afterwards.
lIowever, conceptually, and in accordance with the present invention, axial alignment can occur prior to radial alignment and resultant tangential alignment.
-9- 1313~98 Radial alignment takes place during quiescence, which, as stated above7 means that the rotor is rotating, but free of perturbations, nutation or precession. In~this case, the spin axis z, the spin vector w and the angular momentum h are aligned and the axial and tangential acceleration components at each accelerometer position are zero. Alternatively, the readings of opposing accelerometers can be summed so that the radial co~onent is doubled and the tangential co~onents cancel. A properly aligned pair of accelerometers should give a substantial radial reading and zero readings in the tangential and axial directions.
Any non-zero tangential or axial reading is treated as an error signal and converted by the data processor 14 into an alignment correction command to be implemented through the TT+C module. This correction can be implemented by issuing commands to move mechanical pivots upon which an accelerometer is mounted. However, electronic, rather than mechanical, compensation can be advantageously applied using the memory of the on-board data processor 14.
It is not necessary that the spacecraft be coml?letely quiescent to effect radial alignment. The component cancelling techniques described below for isolating axial acceleration can be applied to compensate for deviations from quiescence. For example, nutation can be accounted for by averaging over a nutation cycle or by compensating for nutation by measuring it and adjusting the acceleration data accordingly. However, in practice, quiescence can be attained to the extent required for purposes of radial alignment.
Axial alignment is effected during periods of thrust, F ~ 0. These may be initiated by ground comnand for orbit and attitude changes, or otherwise. The com~ands are sent via the TT~C module 16 to the propulsion system 18, which in turn affects the spacecraft dynamics 20, which effects are read by the accelerometers A, B, C and D.
The readings of each pair of accelerometers are summed, to cancel and/or compensate periodic acceleration components due to spacecraft precession ~ 3~36~8 and/or nutation, and then integrated over several revolutions, to obtain velocity change dat~ 2( ~Vi) along each of the three corresponding pairs of accelerorneter axes, i = 1, 2, 3.
Since non-translational cornponents of ~Vi for the most part cancel, their values are used to indicate the ef~ective axial component Pa of thrust F. Radial components of thrust F are substantially spin-averaged to zero. Fa is parallel to the average direction of h. Thus, aa is aligned for each accelerometer by zeroing the radial and tangential readings. Assuming radial alignrnent is previously effected, each aa axis is aligned in the process of aligning the respective tangential axes at.
Radial alignment can be conveniently perforrned as often as desired. The axial and tangential alignrnent should be done during north-south station-keeping maneuvers to maximize alignment sensitivity.
Similar results can be achieved from di~ferent approaches. Axial alignrnent can involve orienting the axial axes to maximal values rather than zeroing the radial and tangential axes although this is a less sensitive method. Also, diametrically opposed accelerometers need not be co-oriented. For example, both accelerometers of a pair can have their radial axes directed toward the spin axis and their tangential axes co-parallel. In this case, adding readings would cancel non-translational centripetal acceleration components, although this too is a less sensitive method.
In accordance with the foregoing detailed description, a self-alignment spacecraft navigational system with improved duration and navigational capability is provided. Many modifications and variations of the disclosed embodiment are provided for by the present invention. For example, different accelerometer relative positions and orientations involving different mathematical approaches can be applied readily.
Therefore, the present invention is to be limited only by the scope of the fol10wing claims.
;
BACKGROUND OF THE INVENTION
The present invention relates to spacecraft navigation, and, more particularly, to a system and method providing for on-board alignrnent of navigational accelerometers.
Spacecraft navigation is facing increased demands.
Formerly, spacecraft navigation involved attaining a desired orbit and making occasional corrections to maintain the orbit. More recently, however, rnultiple-year missions involve, for example, more distant objectives of autonomous orbit change or re-entry. These subsequent rnissions require navigational systems capable of meeting objectives far more complex than maintaining orbit and far more enduring than navigational systen~ designed for attaining orbit.
One challenge faced in developing such an enduring navigational system is maintaining its calibration or alignment with respect to an inertial frame of reference. This is particularly true where it is irnpractical to perform this alignment on a regular basis from a ground or other remote station.
One navigational approach has been ~o mount three-axis accelerometers on a gyro-stabilized platform mounted on a launch vehicle or spacecraft. The gyro-stabilized three-axis accelerometers provide a acceleration history which can be integrated once to obtain a velocity history and integrated twice to obtain a position history for the 3~
1313~98 --2~
spacecraft in a known coordinate system. The histories so obtained are used to attain elaborate navigational objectives. The alignment of the accelerometer is maintained by the characteristics of the gyro. Such an arrangement has been used effectively in guiding the ascent of a satellite to a predetermined orbit.
The mechanical limitations of such an arrangement in the face of launch stresses, environmental extremes in space, mechanical creep, other stresses and uncompensatable random drift rates and biases limit the long-term accuracy of the gyros of such an arrangement.
Accordingly, the accelerometer-on-a-gyro approach has not proved satisfactory when a regenerative mission objective requires elaborate navigation months or years after orbit is first attained. Thus, such systems require remote realignment of accelerometers if such realignment is provided for at all. Other systems require dedicated celestial body trackers physically tied to the gyro platform to achieve the realig~Dnent and calibration of the gyros. Still other systems discuss a "strap-down"
method as speculated in "Theory of Inertial Guidance " by Connie L.
McClure, Prentice Hall 1960, pp. 286-291.
Spinning spacecraft have re enduring orientation aids such as attitude and nutation sensors. However, while these are well-adapted for rnaintaining and adjusting attitude, they are not suited for extensive spacecraft navigation. In other words, they caMot readily yield a substitute for the instantaneous acceleration histories provided by accelerometers during periods of thrust.
What is needed is a system and method for realignment of a navigational inertial coordinate system. The system should provide for self-alignment and operate for extended periods essentially autonomously, without the constraints of navigation gyros. The system should be capable of managing demanding navigational objectives after an extended orbital period.
SUMI\~ARY O~ THE INVENTION
A spin-stabiliæed spacecraft includes a platform and a rotor. A pair of circumferentially spaced three-axis accelerometers located on the rotor provide acceleration histories for navigation. In accordance with the present invention, radial alignment of each accelerometer relative to the spacecraft angular momentum, h (~igure 1), is performed during periods of quiescence and axial alignment is performed during periods of thrust.
Using a pair of circumferentially spaced accelerometers permits mathematical cancellation of certain rotationally-induced i0 acceleration components. Other non-translational periodic components can be cancelled through time-averaging or otherwise filtering individual accelerometer readings. Accordingly, acceleration components of interest, namely centripetal and axial-translational components, can be isolated to permit alignment of the accelerometers.
The particulars of cancellation are dependent on the relative placement and orientation of the two accelerometers. The mathematical computations can be greatly simplified in the case where the accelerometers are located at diametrically opposed positions on the rotor and their respective axes co-oriented. With this arrangement, the accelerometer readings can be simply summed, averaged or otherwise filtered to cancel rotational contributions to individual accelerometer readings.
During a period of quiescence, a spin-stabilized spacecraft is spinning in orbit with negligible non-gravitational acceleration, nutation or precession. In this quiescent state, the axial and tangential acceleration components are negligible. A non-zero tangential or axial acceleration reading by the accelerometers can be treated as an error signal to be zeroed through realignment to a plane perpendicular to the angular momentum vector, h.
13~3~98 During a period of thrust with substantial axial acceleration colrponents, the accelerations at the accelerometer locations can be mathematically combined so as to cancel non-translational, spin periodic accelerometer readings due, for example, to torquing. Other non-axial, periodic (e.g. nutational) conponents can be generally time-averaged to zero or digitally filtered. Thus, resultant non-zero readings can serve as axial and tangential accelerometer ali~nent error signals to be corrected through realignment about the radial direction and perpendicular to the angular momentum, h. Periods of thrust occur during initial ascent, north-south station keeping, orbit adjustment, and, if necessaryJ during maneuvers undertaken specifically for accelerometer alignment.
Redundancy and increased sensitivity can ~e provided by adding a second pair of diametrically opposed accelerometers, each located a quarter turn from each of the accelerometers of the first pair.
Thus, a force which may be small relative to the centripetal force of the spinning rotor and therefore difficult to detect as a radial component of an accelerometer of, for example, the first pair, can be readily detected as a tangential component by an accelerometer of the second pair. Also, if an accelerometer fails, one pair remains for navigational purposes.
In accordance with the foregoing description, a system and method for on-board alignment of an accelerometer-within-a-gyro-based navigational system is provided. The accelerometers are well-established for major navigational objectives and are less prone to long-term errors than gyro instruments. While remote calibration can supplement, the provision of on-board alignment ensures the availability of this autonomous navigation capability months or years after initial orbit is attained, depending primarily on the completeness of the on-board ephemeris-generating program and the accuracy of the sensor-based, time independent on-board attitude determination subsystem.
1~3698 -4a Other aspects of this invention are as follows:
A spacecraft with autonomous navigation accelerometer alignment comprising: a spin-stabilized spacecraft having a platform and a rotor with a spin axis; a pair of three-axis accelerometers circumferentially spaced about said rotor; means for applying force to said spacecraft to induce acceleration with a substantial translational axial component; means for detecting and correcting radial alignment error in each accelerometer when said spacecraft is in a quiescent state; and means for detecting and correcting axial alignment error in each accelerometer when said spacecraft is accelerating with a substantial component in said axial direction.
A method of realigning a three-dimensional reference coordinate system on board a thruster-controlled orbiting spin-stabilized spacecraft having a platform and a rotor which rotates about a spin axis, said spacecraft being subject to dynamic imbalance resulting in misalignment of the reference coordinate system, comprising the steps of:
(a) placing at least one pair of accelerometers on the rotor at respective positions radially spaced from said spin axis and circumferentially spaced with respect to each other, each accelerometer having three sensing axes for respectively 1313~98 -4b-sensinq acceleration of the rotor with respect to said spin axis; (b) aligning each of said accelerometers radially with respect to said spin axis while the spacecraft is not being acted on by a thrust force; and (c) aligning each of said accelerometers axially with respect to said spin axis while the spacecraft is being acted upon by a thrust force.
A spin-stabilized spacecraft comprising: a rotor;
a pair of three-axis accelerometers circumferentially spaced upon and fixed to said rotor; means for obtaining quiescent state readings from each of said accelerometers;
means for determining from said quiescent state readings errors in the radial alignment of each of said accelerometers; means for correcting said errors in radial alignment; means for obtaining thrust state readings from each of said accelerometers; means for mathematically combining and time-averaging the thrust state readings of said accelerometers so as to yield calculated axial translational components: and means for axially aligning each of said accelerometers so that its axial readings match said calculated axial translational components.
-4c-A method of calibrating a navigational system comprising the steps of: mounting a pair of three-axis accelerometers circumferentially spaced upon a rotor:
obtaining quiescent state readings from Pach of said accelerometers; determining from said quiescent state readings errors in radial alignment for each accelerometer;
correcting said errors in radial alignment; reading thrust state histories from each accelerometer; mathematically combining said thrust state histories to cancel rotational spin-periodic contributions to said accelerometer readings and filtering said thrust state histories to cancel other non-axial periodic components of acceleration to yield calculated axial translational components; for each accelerometer, determining the axial orientation which would have resulted in a thrust state history corresponding to said calculated axial translational components and align accordingly.
131'~98 BRIEF DESCRIPTION OF THE DRAWINGS
In the accompanying drawings:
FIG. 1 shows R rotor of a spin-stabilized spacecraft with accelerometers mounted thereon in accordanc~ with the present invention.
S FIG. 2 is a block diagram of a navigation~l system in accordance with the present invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
In accordance with the present invention, the rotor 10 of a spin-stabilized spacecraft has mounted thereon n pair of accelerometers A, C as shown in FIG. 1. The accelerometers A, C are aligned radially during quiescent orbit and axially during periods of thrust. Since the accelerometers are spaced circumferentially, their readings can be mathematically combined to cancel spin-periodic and centripetal acceleration components. Other non-axial periodic (e.g.
nutation) components can be generally time-averaged to zero or otherwise filtered to isolate variables of interest.
For computational simplicity, the accelerometers A, C
can be diametrically opposed, i.e., be spaced 180 about the circumference of the rotor 10 and their respective axes co-oriented.
However, the present invention provides for alternative circumferential spacings and relative orientations. A second pair of accelerometers B, D
can be included for increased sensitivity and redundancy.
While only the rotor l0 and accelerometers A, B, C, D
are illustrated, it is understood that a spin-stabilized spacecraft typically includes a platform upon which most of the mission equipment is placed.
The platform ~s relatively despun, whi1e the rotating rotor 10 supplies 1313~8 directional stability to the spacecraft. Of course, equipment, such as attitude sensors intended to be rotated, can be mounted on the rotor 10.
The translational motion of the spacecraft is indicated by velocity vector v, while the position of the spacec~aft having rotor 10 is identified by the displacement vector Ro of the spacecraft center of mass cm relative to the origin of an inertial frame of reference X, Y, Z.
Passing through the center of mass cm are the principal or geometric axes of the spacecraft x, y, z, the last being the spin axis z of the spacecraft.
The instantaneous spin vector ~ is shown non-coincident with the spin axis z, as would be the case during nutation.
Nutation, or more colloquially, wobble, is generally undesirable and so nutation damping is widely employed. In the absence of nutation and other forms of perturbation, ~.) and z are coincident.
During such periods of quiescence, i.e. while the rotor is spinning, but in the absence of nutation, precession or thrust, the angular momentum h is colinear with the spin axis z and the spin vector ~
However, h can diverge from the spin axis z and the spin vector ~ due to the application of a thrust force F whose direction is offset from the instantaneous center of mass cm.
The angular momentum h is particularly important because its orientation can be very precisely determined usin~ celestial, e.g., star, sun, moon, earth feature sensors mounted on the rotor. For example, a star sensor can determine the spacecraft angular mornentum direction relative to inertial coordinates to within 2 arcseconds. Thus, in accordance with the present invention, the alignrnent of accelerometers in inertial space can be precisely determined by aligning the accelerometers with respect to the angular momentum h of the spacecraft.
Before explaining the alignment of the accelerometers A, B, C and D, their use in a navigational mode is explained with reference to FIG. 2. The star and planet sensors 12 provide attitude and angular momentum data while the accelerometers provide acceleration histories. This data is processed by an on-board data processor 14 to determine orbit changes and attitude. The raw data can also be sent to ground processing stations via a tracking, telemetry and comrand (TT+C) mod ule 16 .
A ground station can program the TT~C 16 module with mission objectives which are compared by the data processor 14 with present orbit and attitude data to determine appropriate commands to the propulsion system 18. The action of the propulsion system 18 is fed back to the on-board processor 14 to obtain a better coordination of propulsion commands and their effects on orbit and attitude. The effect of the propulsion system 18 on the spinning spacecraft dynamics 20 are reflected in the readings of the accelerometer pairs A, C and B, D.
Conveniently, the accelerometers are disposed in diametrically opposed pairs. This will cause the centripetal and spin-period components of acceleration due to rotor rotation rather than spacecraft translation to cancel. Thus, by combining the data of an accelerometer pair and computationally compensating or filtering other (e.g. nutation) periodic components, translational acceleration histories due to thrust forces can be obtained in a straightforward manner. An acceleration history so obtained can be integrated once to obtain a velocity history and twice to obtain a positional history. In this way, each accelerometer pair, A and C, B and D, can provide information sufficient for complex navigation.
In addition to providing redundancy, the second pair B, D of accelerometers provide additional sensitivity. Since the pairs are orthogonally disposed with respect to one another, a force which is radial with respect to one pair is tangential with respect to the other. An orbit-correction radial thrust pulse might be very sma~l in magnitude 1313~9~
compared to ongoing radial forces and thus difficult to detect by an accelerometer positioned radially with respect to the pulse. However, an accelerometer disposed one quarter circumference from the radially positioned accelerometer could assess the pulse with much greater precision. This is because the ongoing tangential force is nominally zero and should not mask the effect of the pulse to the extent that the inherent radial force would.
However, the usefulness of the accelerometers A, B, C, D is premised on their orientations being precisely known with respect to an inertial frame. The invention provides for precise alignment relative to an inertial frame as explained below.
Each accelerometer A, B, C, D is R three-axis accelerometer, comprising three mutually orthogonal one-dimensional accelerometers, a nominally radial axis ar, a nominally tangential axis at, and a nominally axial axis aa. The three axes are reliably and rigidly disposed relative to each other so that it can be assumed that if two are properly aligned, then the third is properly aligned. Thus, the challenge reduces to aligning two of the axes of each accelerometer with respect to the angular momentum h, the direction of which is known in inertial coordinates from the readings of the star and planet sensors 12.
The two axes to be aligned are the radial and the axial axes, with tangential alignment automatically resulting. Radial alignment is effected during spacecraft quiescence and the axial alignment is effected during periods of thrust. Since quiescence is the usual state, it is convenient to consider the radial alignment as occurring first and the axial, and concomitantly, the tangential alignment occurring afterwards.
lIowever, conceptually, and in accordance with the present invention, axial alignment can occur prior to radial alignment and resultant tangential alignment.
-9- 1313~98 Radial alignment takes place during quiescence, which, as stated above7 means that the rotor is rotating, but free of perturbations, nutation or precession. In~this case, the spin axis z, the spin vector w and the angular momentum h are aligned and the axial and tangential acceleration components at each accelerometer position are zero. Alternatively, the readings of opposing accelerometers can be summed so that the radial co~onent is doubled and the tangential co~onents cancel. A properly aligned pair of accelerometers should give a substantial radial reading and zero readings in the tangential and axial directions.
Any non-zero tangential or axial reading is treated as an error signal and converted by the data processor 14 into an alignment correction command to be implemented through the TT+C module. This correction can be implemented by issuing commands to move mechanical pivots upon which an accelerometer is mounted. However, electronic, rather than mechanical, compensation can be advantageously applied using the memory of the on-board data processor 14.
It is not necessary that the spacecraft be coml?letely quiescent to effect radial alignment. The component cancelling techniques described below for isolating axial acceleration can be applied to compensate for deviations from quiescence. For example, nutation can be accounted for by averaging over a nutation cycle or by compensating for nutation by measuring it and adjusting the acceleration data accordingly. However, in practice, quiescence can be attained to the extent required for purposes of radial alignment.
Axial alignment is effected during periods of thrust, F ~ 0. These may be initiated by ground comnand for orbit and attitude changes, or otherwise. The com~ands are sent via the TT~C module 16 to the propulsion system 18, which in turn affects the spacecraft dynamics 20, which effects are read by the accelerometers A, B, C and D.
The readings of each pair of accelerometers are summed, to cancel and/or compensate periodic acceleration components due to spacecraft precession ~ 3~36~8 and/or nutation, and then integrated over several revolutions, to obtain velocity change dat~ 2( ~Vi) along each of the three corresponding pairs of accelerorneter axes, i = 1, 2, 3.
Since non-translational cornponents of ~Vi for the most part cancel, their values are used to indicate the ef~ective axial component Pa of thrust F. Radial components of thrust F are substantially spin-averaged to zero. Fa is parallel to the average direction of h. Thus, aa is aligned for each accelerometer by zeroing the radial and tangential readings. Assuming radial alignrnent is previously effected, each aa axis is aligned in the process of aligning the respective tangential axes at.
Radial alignment can be conveniently perforrned as often as desired. The axial and tangential alignrnent should be done during north-south station-keeping maneuvers to maximize alignment sensitivity.
Similar results can be achieved from di~ferent approaches. Axial alignrnent can involve orienting the axial axes to maximal values rather than zeroing the radial and tangential axes although this is a less sensitive method. Also, diametrically opposed accelerometers need not be co-oriented. For example, both accelerometers of a pair can have their radial axes directed toward the spin axis and their tangential axes co-parallel. In this case, adding readings would cancel non-translational centripetal acceleration components, although this too is a less sensitive method.
In accordance with the foregoing detailed description, a self-alignment spacecraft navigational system with improved duration and navigational capability is provided. Many modifications and variations of the disclosed embodiment are provided for by the present invention. For example, different accelerometer relative positions and orientations involving different mathematical approaches can be applied readily.
Therefore, the present invention is to be limited only by the scope of the fol10wing claims.
;
Claims (9)
1. A spacecraft with autonomous navigation accelerometer alignment comprising:
a spin-stabilized spacecraft having a platform and a rotor with a spin axis;
a pair of three-axis accelerometers circumferentially spaced about said rotor;
means for applying force to said spacecraft to induce acceleration with a substantial translational axial component means for detecting and correcting radial alignment error in each accelerometer when said spacecraft is in a quiescent state;
and means for detecting and correcting axial alignment error in each accelerometer when said spacecraft is accelerating with a substantial component in said axial direction.
a spin-stabilized spacecraft having a platform and a rotor with a spin axis;
a pair of three-axis accelerometers circumferentially spaced about said rotor;
means for applying force to said spacecraft to induce acceleration with a substantial translational axial component means for detecting and correcting radial alignment error in each accelerometer when said spacecraft is in a quiescent state;
and means for detecting and correcting axial alignment error in each accelerometer when said spacecraft is accelerating with a substantial component in said axial direction.
2. The spacecraft of claim 1 further characterized in that said accelerometers are in substantially diametrically opposed locations with respect to said spin axis.
3. The spacecraft of claim 2 further characterized in that the respective axes of said accelerometers are co-oriented.
4. The spacecraft of claim 2 further comprising at least a second pair of accelerometers in substantially diametrically opposed locations on said rotor with respect to each other, each of said second pair of accelerometers being located substantially one quarter circumference from each of the accelerometers of said first pair.
5. A method of realigning a three-dimensional reference coordinate system on board a thruster-controlled orbiting spin-stabilized spacecraft having a platform and a rotor which rotates about a spin axis, said spacecraft being subject to dynamic imbalance resulting in misalignment of the reference coordinate system, comprising the steps of:
(a) placing at least one pair of accelerometers on the rotor at respective positions radially spaced from said spin axis and circumferentially spaced with respect to each other, each accelerometer having three sensing axes for respectively sensing acceleration of the rotor with respect to said spin axis;
(b) aligning each of said accelerometers radially with respect to said spin axis while the spacecraft is not being acted on by a thrust force; and (c) aligning each of said accelerometers axially with respect to said spin axis while the spacecraft is being acted upon by a thrust force.
(a) placing at least one pair of accelerometers on the rotor at respective positions radially spaced from said spin axis and circumferentially spaced with respect to each other, each accelerometer having three sensing axes for respectively sensing acceleration of the rotor with respect to said spin axis;
(b) aligning each of said accelerometers radially with respect to said spin axis while the spacecraft is not being acted on by a thrust force; and (c) aligning each of said accelerometers axially with respect to said spin axis while the spacecraft is being acted upon by a thrust force.
6. The method of claim 5 with step (a) being further characterized in that said accelerometers are placed in substantially diametrically opposed locations on said rotor.
7. The method of claim 6 with step (a) further involving placing at least a second pair of accelerometers in substantially diametrically opposed locations on said rotor, each accelerometer of said second pair being spaced circumferentially one quarter circle from each accelerometer of said first pair.
8. A spin-stabilized spacecraft comprising:
a rotor;
a pair of three-axis accelerometers circumferentially spaced upon and fixed to said rotor;
means for obtaining quiescent state readings from each of said accelerometers;
means for determining from said quiescent state readings errors in the radial alignment of each of said accelerometers;
means for correcting said errors in radial alignment;
means for obtaining thrust state readings from each of said accelerometers;
means for mathematically combining and time-averaging the thrust state readings of said accelerometers so as to yield calculated axial translational components; and means for axially aligning each of said accelerometers so that its axial readings match said calculated axial translational components.
a rotor;
a pair of three-axis accelerometers circumferentially spaced upon and fixed to said rotor;
means for obtaining quiescent state readings from each of said accelerometers;
means for determining from said quiescent state readings errors in the radial alignment of each of said accelerometers;
means for correcting said errors in radial alignment;
means for obtaining thrust state readings from each of said accelerometers;
means for mathematically combining and time-averaging the thrust state readings of said accelerometers so as to yield calculated axial translational components; and means for axially aligning each of said accelerometers so that its axial readings match said calculated axial translational components.
9. A method of calibrating a navigational system comprising the steps of:
mounting a pair of three-axis accelerometers circumferentially spaced upon a rotor;
obtaining quiescent state readings from each of said accelerometers;
determining from said quiescent state readings errors in radial alignment for each accelerometer;
correcting said errors in radial alignment;
reading thrust state histories from each accelerometer;
mathematically combining said thrust state histories to cancel rotational spin-periodic contributions to said accelerometer readings and filtering said thrust state histories to cancel other non-axial periodic cornponents of acceleration to yield calculated axial translational components;
for each accelerometer, determining the axial orientation which would have resulted in a thrust state history corresponding to said calculated axial translational components and align accordingly.
mounting a pair of three-axis accelerometers circumferentially spaced upon a rotor;
obtaining quiescent state readings from each of said accelerometers;
determining from said quiescent state readings errors in radial alignment for each accelerometer;
correcting said errors in radial alignment;
reading thrust state histories from each accelerometer;
mathematically combining said thrust state histories to cancel rotational spin-periodic contributions to said accelerometer readings and filtering said thrust state histories to cancel other non-axial periodic cornponents of acceleration to yield calculated axial translational components;
for each accelerometer, determining the axial orientation which would have resulted in a thrust state history corresponding to said calculated axial translational components and align accordingly.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US06/897,237 US4749157A (en) | 1986-08-18 | 1986-08-18 | Spacecraft accelerometer auto-alignment |
| US897,237 | 1986-08-18 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1313698C true CA1313698C (en) | 1993-02-16 |
Family
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| Application Number | Title | Priority Date | Filing Date |
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| CA000543321A Expired - Fee Related CA1313698C (en) | 1986-08-18 | 1987-07-29 | Spacecraft accelerometer auto-alignment |
Country Status (8)
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| EP (1) | EP0277217B1 (en) |
| JP (1) | JPH0774739B2 (en) |
| CN (1) | CN1010710B (en) |
| CA (1) | CA1313698C (en) |
| DE (1) | DE3772748D1 (en) |
| IL (1) | IL83328A (en) |
| WO (1) | WO1988001389A1 (en) |
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| FR2637565B1 (en) * | 1988-10-06 | 1991-01-11 | Aerospatiale | ACTIVE CONTROL SYSTEM BASED ON THREE AXES OF THE ATTITUDE OF A GEOSTATIONARY SATELLITE |
| US6282467B1 (en) * | 1997-10-14 | 2001-08-28 | The Boeing Company | Three-axis inertial attitude determination for spinning spacecraft |
| US6532419B1 (en) * | 1998-09-23 | 2003-03-11 | Magellan Dis, Inc. | Calibration of multi-axis accelerometer in vehicle navigation system |
| DE19843774C1 (en) * | 1998-09-24 | 1999-11-11 | Deutsch Zentr Luft & Raumfahrt | Space vehicle incorporating mechanical devices |
| EP1779659B1 (en) | 2004-08-12 | 2020-10-07 | Gracenote Inc. | Selection of content from a stream of video or audio data |
| CN1737580A (en) * | 2004-08-17 | 2006-02-22 | 皇家飞利浦电子股份有限公司 | Correction method and apparatus |
| CN100349018C (en) * | 2005-10-12 | 2007-11-14 | 北京航空航天大学 | Internal and external element correcting method of star sensor |
| ITTO20070335A1 (en) | 2007-05-15 | 2008-11-16 | Sequoia It S R L | WIDE BAND ACCELEROMETER WITH SELF-RECOGNITION OF CALIBRATION |
| CN101625571B (en) * | 2009-07-25 | 2010-12-29 | 大连理工大学 | Method for simulating six degrees of freedom movement of spinning aircraft |
| US8718963B2 (en) * | 2010-04-27 | 2014-05-06 | Memsic, Inc. | System and method for calibrating a three-axis accelerometer |
| CN101943584B (en) * | 2010-07-02 | 2012-05-09 | 哈尔滨工程大学 | Alignment method based on CCD (Charge Coupled Device) star sensor |
| CN102114918B (en) * | 2010-12-31 | 2013-03-20 | 北京航空航天大学 | Attitude control feedback loop based on combined fixed attitude of multi-rate sensor |
| CN103063870B (en) * | 2012-12-25 | 2014-09-24 | 中北大学 | Measuring method for initial velocity of ammunition body at gun muzzle |
| CN103256928B (en) * | 2013-04-28 | 2015-05-20 | 南京航空航天大学 | Distributed inertial navigation system and posture transfer alignment method thereof |
| US9128116B2 (en) * | 2013-07-05 | 2015-09-08 | Hti Ip, L.L.C. | Automatic alignment of a vehicle three-axes accelerometer |
| DE102013214020A1 (en) * | 2013-07-17 | 2015-02-19 | Stabilo International Gmbh | Digital pen |
| CA2943443C (en) | 2014-03-28 | 2022-11-01 | Gedex Systems Inc. | Absolute vector gravimeter and methods of measuring an absolute gravity vector |
| WO2016119056A1 (en) * | 2015-01-28 | 2016-08-04 | Gedex Inc. | Inertial sensing augmentation for navigation of spacecraft |
| CN105115501B (en) * | 2015-08-11 | 2018-06-05 | 中国航空工业集团公司西安飞机设计研究所 | 3-axis acceleration modification method and device |
| CN105467158B (en) * | 2015-12-28 | 2019-01-29 | 太原航空仪表有限公司 | The air speed modification method of Helicopter Maneuver Flight |
| US9885232B2 (en) * | 2016-06-09 | 2018-02-06 | Noralis Limited | Method for determining position with improved calibration with opposing sensors |
| RU2669164C1 (en) * | 2017-08-01 | 2018-10-08 | Публичное акционерное общество "Ракетно-космическая корпорация "Энергия" имени С.П. Королева" | Method for calibration of microacceleration sensor in space flight |
| CN109297507B (en) * | 2018-09-27 | 2021-11-12 | 南京邮电大学 | Human body limb movement capturing method free of alignment action based on inertial sensor |
| CN109635388B (en) * | 2018-11-29 | 2023-04-18 | 中国航空工业集团公司沈阳飞机设计研究所 | Whole-machine acceleration parameter measuring and calculating method |
| US11459133B2 (en) * | 2019-05-08 | 2022-10-04 | NovaWurks, Inc. | Method of aligning a spacecraft frame |
| RU2721589C1 (en) * | 2019-07-23 | 2020-05-20 | Акционерное общество «Информационные спутниковые системы» имени академика М.Ф. Решетнёва» | Space accelerometer |
| CN111664995B (en) * | 2020-07-08 | 2022-04-05 | 福州大学 | Satellite three-dimensional rotation quantity testing device and testing method |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2942864A (en) * | 1955-07-29 | 1960-06-28 | Stephen J Sikora | Device for measuring horizontal acceleration |
| US3269179A (en) * | 1959-05-29 | 1966-08-30 | Honeywell Regulator Co | Navigational instruments |
| US3350916A (en) * | 1961-06-01 | 1967-11-07 | Bosch Arma Corp | Accelerometer calibration on inertial platforms |
| US3442140A (en) * | 1964-12-24 | 1969-05-06 | North American Rockwell | Drift rate compensation for acceleration sensitivity of an inertial navigation platform |
| GB2130377B (en) * | 1982-11-15 | 1985-12-11 | Marconi Avionics | Inertial systems |
| GB2146776B (en) * | 1983-09-16 | 1986-07-30 | Ferranti Plc | Accelerometer systems |
-
1986
- 1986-08-18 US US06/897,237 patent/US4749157A/en not_active Expired - Lifetime
-
1987
- 1987-07-23 JP JP62505219A patent/JPH0774739B2/en not_active Expired - Lifetime
- 1987-07-23 WO PCT/US1987/001801 patent/WO1988001389A1/en active IP Right Grant
- 1987-07-23 EP EP87905680A patent/EP0277217B1/en not_active Expired
- 1987-07-23 DE DE8787905680T patent/DE3772748D1/en not_active Expired - Fee Related
- 1987-07-26 IL IL83328A patent/IL83328A/en not_active IP Right Cessation
- 1987-07-29 CA CA000543321A patent/CA1313698C/en not_active Expired - Fee Related
- 1987-08-18 CN CN87105637.2A patent/CN1010710B/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| CN1010710B (en) | 1990-12-05 |
| JPH01501015A (en) | 1989-04-06 |
| DE3772748D1 (en) | 1991-10-10 |
| US4749157A (en) | 1988-06-07 |
| EP0277217B1 (en) | 1991-09-04 |
| IL83328A (en) | 1991-03-10 |
| WO1988001389A1 (en) | 1988-02-25 |
| JPH0774739B2 (en) | 1995-08-09 |
| EP0277217A1 (en) | 1988-08-10 |
| CN87105637A (en) | 1988-04-20 |
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