JOURNAL OF GUIDANCE, CONTROL, AND DYNAMICS
`Vol. 36, No. 3, May–June 2013
`
`History of Key Technologies
`
`Blazing Gyros: The Evolution of Strapdown Inertial
`Navigation Technology for Aircraft
`
`Paul G. Savage
`Strapdown Associates, Inc., Maple Plain, Minnesota 55359
`
`DOI: 10.2514/1.60211
`
`Introduction
`
`I NERTIAL navigation is the process of autonomously calculating
`
`the position and velocity of a moving vehicle from measurements
`of angular rotation and linear acceleration provided by vehicle-
`mounted inertial sensors (gyros and accelerometers). The first
`inertial navigation system (INS) was developed at the Massachusetts
`Institute of Technology (MIT)
`Instrumentation Laboratory
`(eventually becoming the Charles Stark Draper Laboratory) for
`ballistic missile guidance [1]. (The INS includes velocity, attitude,
`heading, etc. outputs. In commercial application parlance, it also
`includes guidance steering outputs based on an input waypoint-
`defined flight profile.) Soon thereafter, the technology was applied to
`aircraft navigation, with four companies eventually dominating the
`U.S. aircraft inertial navigation industry in the 1960s: Honeywell
`Aerospace and Defense Group with gyro design and manufacturing
`in Minneapolis, Minnesota, and INS design, development, and
`manufacturing in Clearwater, Florida; Kearfott in Wayne, New
`Jersey; Litton Guidance and Control Division in Woodland Hills,
`California; and Delco Electronics Division of General Motors in
`Milwaukee, Wisconsin. Honeywell specialized in high-accuracy
`systems and introduced a new electrostatically suspended gyro
`(ESG) technology for precision applications. Delco concentrated
`on transoceanic commercial and military cargo/tanker aircraft
`applications using the Carousel IV system (a variation of the Titan II
`ballistic missile inertial guidance set). Litton and Kearfott focused on
`medium-accuracy military tactical aircraft and airborne missile
`applications. To achieve required gyro accuracy, each of the
`
`aforementioned inertial navigation systems was configured with
`gimbaled platforms to isolate the inertial sensors from aircraft
`angular rates.
`INS advanced development at Litton and Kearfott in the 1960s
`centered on improving accuracy, reducing size, weight, and cost,
`and improving reliability of gimbaled INS products (important
`requirements for expanding military aircraft and airborne missile
`applications). A key contribution was the introduction of dry
`tuned-rotor-gyro (TRG) technology. Delco focused on reliability
`improvement for transport applications. Honeywell focused on
`improved accuracy and reliability of ESG gimbaled systems.
`For future cost, reliability, and size reduction based in part on
`projected advances in computer technology, several companies (most
`prominently, Honeywell) focused a significant portion of company
`resources on a radically new strapdown approach to inertial naviga-
`tion: replacing the gimbaled platform with a computerized analytical
`equivalent and mounting (“strapping down”) the inertial sensors
`directly to the user vehicle. Based on technical books, journals,
`internet archives, discussions with past colleagues, but mostly from
`direct experience and personal records, this paper describes the
`curious and sometimes convoluted path by which the Honeywell
`strapdown program eventually led to development of the ring laser
`gyro (RLG) strapdown INS and conversion from gimbaled to
`strapdown technology throughout the airborne inertial navigation
`industry.
`For technical background, the paper first discusses the concept
`of inertial navigation using gimbaled versus strapdown system
`
`Paul G. Savage is an internationally recognized expert in the design and testing of strapdown inertial navigation
`systems, and is the President of Strapdown Associates, Inc., a company he founded in 1980. Strapdown Associates has
`provided software and engineering services to government agencies and aerospace companies for strapdown inertial
`system configuration definition, flight software development, system simulation, and testing. Mr. Savage has
`published and presented several papers on strapdown inertial navigation systems and associated computational
`elements. From 1974 to 2009, he served as an author/speaker on several NATO AGARD and Research and
`Technology Organisation technology transfer lecture series tours. From 1981 to 2009, Mr. Savage provided his
`Introduction to Strapdown Inertial Navigation Systems course to the aerospace industry. Since 2011, he has provided a
`two-day focused version of the Intro to Strapdown course onsite at host facilities in the continental United States. He
`has written and published the textbook Strapdown Analytics (available from Strapdown Associates), detailing the
`analytical aspects of strapdown inertial navigation system design. From 1963 to 1980, Mr. Savage was employed at
`Honeywell Avionics Division as Senior Principal Engineering Fellow, where he led engineering design teams and
`provided technical consultation to Honeywell engineering managers for system design, analysis, software
`development, simulation, and integration/test in the evolutionary development of laser gyro strapdown inertial
`navigation systems for military and commercial aircraft. From 1971 through 1975, he was the Engineering Manager
`and System Design Engineer for the strapdown Honeywell Laser Inertial Navigation System (LINS), the first to prove
`the readiness of laser gyro strapdown inertial navigation technology for aircraft applications, as demonstrated during
`a landmark flight-test series at Holloman Air Force Base in 1975. Mr. Savage is a graduate from the Massachusetts
`Institute of Technology, where he received his BS and MS degrees in aeronautical engineering in 1960. He is a Senior
`Member of the AIAA.
`
`Received 4 September 2012; revision received 10 September 2012; accepted for publication 4 September 2012; published online 19 March 2013. Copyright ©
`2012 by Strapdown Associates, Inc. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Copies of this paper may be made
`for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA
`01923; include the code 1533-3884/13 and $10.00 in correspondence with the CCC.
`
`637
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`Downloaded by Reprints Desk Access on May 24, 2021 | http://arc.aiaa.org | DOI: 10.2514/1.60211
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`638
`
`SAVAGE
`
`Fig. 1 Gimbaled platform schematic.
`
`the key
`(gyros),
`rotation sensors
`implementations. Angular
`instruments in inertial systems, are briefly described and compared
`for the strapdown approaches considered along the way. The
`Honeywell GG1300 RLG is described in more detail, the first RLG to
`meet aircraft strapdown INS accuracy and reliability requirements.
`Most of the paper focuses on the interrelationships and testing of four
`Honeywell strapdown inertial systems developed in the early through
`mid-1970s:
`the Advanced Tactical
`Inertial Guidance System
`(ATIGS), the Laser Inertial Navigation System (LINS), the Ring
`Laser Gyro Navigator (RLGN), and the Laser Inertial Reference
`System (IRS) Prototype using a new, smaller size Honeywell
`GG1342 RLG, the latter known at Honeywell as the 7 × 7 Laser IRS
`Prototype system. Flight testing of the four Honeywell systems is
`described: ATIGS by the U.S. Navy’s Naval Weapons Center
`(NWC), LINS by the U.S. Air Force Central Inertial Guidance Test
`Facility (CIGTF) at Holloman Air Force Base, the RLGN by the U.S.
`Navy’s Naval Air Development Center (NADC), and the 7 × 7 Laser
`IRS Prototype by The Boeing Company. The proposal process for the
`new Boeing 757∕767 commercial airplane strapdown IRS is also
`discussed, which, with 7 × 7 Laser IRS Prototype flight-test results,
`led to the selection of Honeywell for the multiyear 757∕767 IRS
`large-scale procurement contract, the first for both aircraft strapdown
`inertial systems and for RLGs. The paper concludes with an epilogue
`of how the Honeywell and Boeing programs soon led to the general
`conversion from gimbaled to strapdown system technology
`throughout the entire aircraft inertial navigation industry.
`
`Gimbaled Versus Strapdown System
`Implementation Approaches
`Inertial sensors in a gimbaled INS are orthogonally mounted to
`a common base (platform) surrounded by concentric gimbals,
`interconnected to the platform and each other through ball-bearing
`shafts (depicted schematically in Fig. 1, the cylinders representing
`gyros, cubes accelerometers, and input axes are dashed). Nonrotation
`of the Fig. 1 sensor platform is actively controlled by torque motors
`mounted on the gimbal shafts, driven by gyro output measurements
`of platform rotation. Four-gimbal platforms (Fig. 1) were typically
`used in aircraft applications for operation at any vehicle angular
`orientation, while maintaining the three inner gimbal shafts near
`perpendicularity (the optimum orientation for minimum gimbal
`torque-motor size/power
`requirements under dynamic angular
`maneuvering).
`Figure 2 depicts the interfaces between the gimbaled platform and
`the INS navigation computer. Platform accelerometer outputs in
`Fig. 2 (vector components along platform axes) are processed in the
`
`INS computer to calculate velocity and position. Feedbacks from
`the navigation computer bias the platform gyros (when allowable for
`the gyro configuration), commanding the platform to follow pre-
`scribed small angular rotation rates (e.g., to maintain a locally vertical
`platform orientation relative to the Earth in the presence of Earth’s
`rotation rate and aircraft translational motion over the Earth). Figure 3
`depicts the classical strapdown approach to inertial navigation.
`Unlike Fig. 2, the inertial sensor mount in Fig. 3 is directly connected
`(usually through silicone elastomeric isolators) to the vehicle struc-
`ture ( “strapdown”), thereby eliminating the Fig. 2 intermediate
`gimbal/torque-motor assembly. With minor differences, the naviga-
`tion computations in Fig. 3 are the same as for the gimbaled INS
`configuration of Fig. 2. The basic difference is that the specific
`force acceleration components, provided directly from platform
`accelerometers in Fig. 2 to the navigation computations block, are
`calculated in Fig. 3 with a vector-transformation operation performed
`in the system computer. This analytically converts the strapdown
`accelerometer outputs to the values that would be measured from
`accelerometers mounted on a Fig. 2 gimbaled platform. The second
`input to the vector transformation block in Fig. 3 is the angular
`orientation (attitude) of the gyro/accelerometer strapdown mount
`relative to the equivalent Fig. 2 gyro-stabilized platform. The attitude
`data are calculated by high-speed digital integration operations on
`strapdown rate-gyro inputs. The feedback gyro biasing operation in
`Fig. 1 for commanding platform rotation rates is also present in Fig. 3,
`
`"PLATFORM" SPECIFIC FORCE
`ACCELERATION VECTOR
`
`SYSTEM COMPUTER
`
`ACCELEROMETERS
`
`GYROS
`
`STABLE
`PLATFORM
`
`GIMBALS
`WITH
`TORQUE
`MOTORS
`
`VELOCITY AND
` POSITION
`
`NAVIGATION
`COMPUTATION
`
`GYRO BIAS RATES TO
`COMMAND PLATFORM
`INERTIAL ROTATION
`
`ELASTOMERIC
`ISOLATORS
`
`VEHICLE ATTACHMENT
`Fig. 2 Gimbaled inertial navigation system.
`
`Downloaded by Reprints Desk Access on May 24, 2021 | http://arc.aiaa.org | DOI: 10.2514/1.60211
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`SAVAGE
`
`639
`
`STRAPDOWN
`SPECIFIC FORCE
`ACCELERATION
`VECTOR
`
`STRAPDOWN
`ACCELEROMETERS
`
`STRAPDOWN
`RATE GYROS
`
`VEHICLE ATTACHMENT
`
`STRAPDOWN
`INERTIAL ANGULAR
`RATES
`
`SYSTEM COMPUTER
`
`PLATFORM
`SPECIFIC FORCE
`ACCELERATION
`VECTOR
`
`VECTOR
`TRANSFORMATION
`
`STRAPDOWN
`TO
`PLATFORM
`ATTITUDE
`
`NAVIGATION
`COMPUTATIONS
`
`VELOCITY AND
` POSITION
`
`ATTITUDE
`INTEGRATION
`OPERATIONS
`
`PLATFORM
`INERTIAL
`ANGULAR
`RATES
`
`Fig. 3 Rate-gyro-based strapdown inertial navigation system.
`
`ELASTOMERIC
`ISOLATORS
`
`but as part of the attitude computational process, so that computed
`attitude outputs become referenced to Fig. 2 stable “platform axes.”
`Prior to engaging the inertial navigation function, the initial
`angular orientation of the Fig. 2 platform (or Fig. 3 computed
`attitude) must be established during “initial alignment” operations.
`Under quasi-stationary conditions, this is a self-alignment function in
`which the physical gimbaled platform (or strapdown computed
`attitude) is controlled to a locally level orientation, based on acceler-
`ometer measurements in Fig. 2 (or transformed acceleration measure-
`ments in Fig. 3). Simultaneously, accelerometer heading relative to
`true north is ascertained from gyro controlled platform accelerometer
`rates in Fig. 2 (or transformed acceleration rates in Fig. 3) in response
`to earth’s rotation rate, based on the fundamental characteristics
`that the horizontal component of earth rate points north.
`An alternative to the Fig. 3 rate-gyro strapdown approach is to use
`attitude gyros whose output represents the angular orientation
`relative to nonrotating inertial space, of the gyro case (hence, the
`sensor mount to which the gyros are attached). The concept is
`depicted in Fig. 4. Comparing Fig. 4 with Fig. 3 shows the principal
`advantage afforded by use of attitude gyros: eliminating Fig. 3 high-
`speed computation requirements for integrating strapdown angular
`rates into attitude. For 1960 and early 1970 computer technology
`limitations, this was an important advantage. The penalty was the
`attitude-gyro requirement for accurate readout capability at any
`angular orientation of the strapdown sensor mount (fixed to the
`vehicle). Other operations in Fig. 4 parallel those in Fig. 3, but with
`the integration of platform inertial rotation rates executed at a low rate
`in a separate computation block, then combined with gyro data to
`form the Fig. 3 strapdown-to-platform attitude data for acceleration
`transformation. In the 1960–1970 time frame, only the ESG (with
`modification for wide-angle readout) had the accuracy potential for
`aircraft strapdown INS application.
`
`Compared with the traditional gimbaled INS approach, strapdown
`technology promised future cost reduction and improved reliability
`through elimination of mechanical parts, gimbal shaft-angle
`transducers, slip-ringed electrical connections to/from platform
`components, and high-power gimbal-motor-drive electronics [2–4].
`Another strapdown advantage was touted in redundant applications
`(particularly for the Fig. 3 rate-gyro approach) by using skew-aligned
`sensor input axis geometries (mounting the inertial sensors in a
`nonorthogonal arrangement, having no three input axes coplanar)
`[3,5–8]. Figure 3 orthogonal axis angular rates/acceleration com-
`ponents are then obtained from any set of three or more skewed gyros/
`accelerometers by analytical conversion operations in the system
`computer. As a result, only one additional gyro/accelerometer is
`required for each level of system redundancy (in contrast with
`classical
`redundant gimbaled INS configurations requiring a
`duplicate INS for each redundancy level). For the Fig. 3 rate-gyro
`implementation, strapdown sensor outputs can also be used for
`other vehicle functions (e.g., aircraft axis angular rate/acceleration
`for flight control/stability augmentation),
`thereby reducing the
`multiplicity of dedicated aircraft inertial sensors normally required
`for non-INS related functions. (Note,
`to achieve a consistent
`redundancy level throughout a system, skewed redundant sensors
`must also be interfaced with equivalent redundancy level computers,
`power supplies, and cabling designed to also block single failure
`propagation between redundant channels.)
`Before strapdown technology could be considered viable, two
`major technological advances were required to achieve accuracy
`levels already attained in gimbaled INS high-volume production:
`1) computer advances
`to handle new rate-gyro strapdown
`INS throughput requirements and, most
`important, 2) never-yet-
`achieved gyro accuracies in a nongimbaled strapdown dynamic rate
`environment.
`
`STRAPDOWN
`SPECIFIC FORCE
`ACCELERATION
`VECTOR
`
`STRAPDOWN
`ACCELEROMETERS
`
`STRAPDOWN
`ATTITUDE GYROS
`
`VEHICLE ATTACHMENT BY
`ELASTOMERIC ISOLTORS
`
`STRAPDOWN INERTIAL
`ANGULAR ATTITUDE
`
`SYSTEM COMPUTER
`
`PLATFORM
`SPECIFIC FORCE
`ACCELERATION
`VECTOR
`
`VECTOR
`TRANSFORMATION
`
`STRAPDOWN
`TO
`PLATFORM
`ATTITUDE
`
`PLATFORM
`INERTIAL
`ATTITUDE
`OFFSET
`
`PLATFORM
`INERTIAL
`ATTITUDE
`INTEGRATION
`OPERATIONS
`
`VELOCITY AND
` POSITION
`
`NAVIGATION
`COMPUTATIONS
`
`PLATFORM
`INERTIAL
`ROTATION
`RATES
`
`Fig. 4 Attitude-gyro-based strapdown inertial navigation system.
`
`Downloaded by Reprints Desk Access on May 24, 2021 | http://arc.aiaa.org | DOI: 10.2514/1.60211
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`640
`
`SAVAGE
`
`The gimbaled platform in an INS exists for two basic reasons:
`to reduce gyro error (as noted previously) induced by high-input
`angular rates and to create a stable mount for the accelerometers at
`a known angular orientation relative to the Earth (for Earth-based
`position/velocity determination) [9]. By eliminating the stabilized
`platform, performance requirements for strapdown gyros dramati-
`cally increase in scale-factor accuracy and sensor-to-sensor
`alignment for reduced error buildup under attitude changes. A more
`subtle requirement for aircraft systems is the need for long-term
`stability of critical inertial sensor performance parameters due to the
`lack of built-in rotation test equipment (used to measure and
`compensate inertial sensor performance parameters in a test facility).
`In a gimbaled INS, rotation calibration can be provided by the gimbal
`assembly during a special test mode. (For the Delco Carousel
`gimbaled system, the stable platform is used as a base for mounting
`a synchronously controlled rotating “table,” which houses the hori-
`zontal sensors. By continuously rotating the table relative to the
`stabilized platform at 1 rpm, horizontal accelerometer and gyro
`biases become averaged, effectively canceling their impact on
`position/velocity error buildup.) Additionally, normal self-alignment
`operations before navigation mode engagement implicitly compen-
`sate critical sensor errors in a gimbaled INS. During self-alignment,
`all inertial navigation systems develop platform tilt and heading
`errors that cancel horizontal accelerometer and east gyro bias. For
`gimbaled systems, the cancellation remains after navigation mode
`entry because the sensor platform remains at its self-alignment
`orientation under subsequent vehicle maneuvering. In contrast,
`strapdown sensors rotate with the vehicle during navigation, altering
`their orientation from the alignment attitude (the worst case being a
`180 deg heading rotation following initial alignment, effectively
`doubling the impact of the sensor errors).
`To meet the strapdown performance challenge, major design
`changes had to be incorporated in conventional angular-momentum-
`based gyros, and a new angular rate sensor was introduced, the RLG,
`based on the relativistic properties of light [10]. (The term “gyro” is
`now commonly used for all angular rate sensing inertial instruments.
`It is derived from “gyroscope,” the term originally used for angular
`rate sensing based on the gyroscopic angular-momentum properties
`of rotating mass.)
`
`Conventional Strapdown Gyro Development
`The distinguishing characteristic between angular-momentum-
`based gyro configurations is the method used to contain the spinning
`rotor without inducing error from spurious torques on the rotor
`assembly. Angular-momentum gyros considered during the 1960s
`for strapdown application were the single-degree-of-freedom floated
`rate integrating gyro (RIG), the two-axis dynamically compensated
`dry TRG, and the ESG ([11–13], Chaps. 7–9 in [14], Chap. 4 in [15]).
`The RIG (Fig. 5) supports the rotor assembly by the buoyancy of
`surrounding viscous fluid. The TRG (Fig. 6) supports the rotor by
`flexure pivots connected through an intermediate gimbal to the
`spin-motor shaft. Pivot-flexure spring torques on the rotor, developed
`under off-null operation, are thereby compensated by dynamic
`motion of the spinning gimbal. The two-axis ESG (Fig. 7) supports
`
`PICKOFF ANGLE
`PROPORTIONAL TO INTEGRAL
`OF INPUT AXIS RATE MINUS
`TORQUER BIAS RATE
`
`OUTPUT
`AXIS
`
`SPIN REFERENCE AXIS
`(CASE FIXED)
`
`SPIN AXIS
`
`ROTOR (SPIN MOTOR)
`
`TORQUE
`GENERATOR
`
`INPUT AXIS
`(CASE FIXED)
`
`HERMETICALLY SEALED
`CYLINDRICAL ROTOR
`ASSEMBLY (OR FLOAT)
`
`PIVOT FOR JEWEL RING
`OR MAGNETIC CENTERING
`OF ROTOR ASSEMBLY
`
`VISCOUS FLUID FULLY SUPPORTS
`CYLINDRICAL ROTOR ASSEMBLY AT
`NEUTRAL BUOYANCY AND PROVIDES
`DAMPING ABOUT OUTPUT AXIS
`Fig. 5 Single-degree-of-freedom floated RIG.
`
`TORSIONALLY
`FLEXIBLE COUPLING
`
`ROTOR
`
`DYNAMIC GIMBAL MOTION
`UNDER OFF-NULL OPERATION
`CANCELS FLEXURE SPRING
`TORQUE ON ROTOR
`
`DUAL PICKOFFS MEASURE
`ANGLES BETWEEN SPIN
`AXIS AND CASE
`
`TORQUER
`MAGNETS FOR
`BIASING
`ATTACHED TO
`ROTOR
`
`NEAR VACUUM
`BETWEEN ROTOR
`AND CASE REDUCES
`WINDAGE TORQUE
`
`SPIN-MOTOR SHAFT
`(ALIGNED WITH GYRO CASE)
`
`Fig. 6 Two-axis dry TRG.
`
`the free rotor with an electrostatic field applied by case-mounted
`electrodes.
`The RIG (Fig. 5) measures the integrated difference between input
`axis angular rate and applied bias rate, the latter provided by an
`electrical torquer. The TRG (Fig. 6) measures the two-axis angular
`orientation between the gyro case and rotor. The TRG also contains a
`torquer assembly for intentionally precessing the spinning rotor
`axis relative to inertial space. To meet the strapdown gyro perfor-
`mance challenge, closed-loop TRG and RIG configurations had to be
`developed, in which electrical input to the torquers are provided in
`feedback fashion to maintain pickoff output null. The torquer input
`command thereby becomes proportional to gyro case angular rate and
`the strapdown rate-gyro output in Fig. 3.
`The ESG (Fig. 7) measures the two-axis angular orientation of the
`case relative to the rotor which, having no accurate torquing means,
`maintains a fixed angular orientation relative to inertial space. The
`angular orientation of the case relative to the rotor is determined with
`the Fig. 7 Honeywell hollow-shell ESG rotor approach by measuring
`the time interval between pickoff sensed light reflections from a
`scribe pattern etching on the rotor. Rockwell Autonetics Division
`(Anaheim, California), a latecomer to ESG technology, used a small
`solid mass-unbalanced rotor to generate a detectable modulation
`signature in the suspension electrodes for readout: mass unbalance
`modulation (MUM). ([11,13], Chap. 8 in [14]). In gimbaled applica-
`tions, ESG case-to-rotor angular orientation is maintained at the
`pickoff null by gimbal command rotation. For strapdown applica-
`tions, ESG pickoffs have to be accurate at any case orientation
`relative to the rotor (wide-angle readout) for Fig. 4 attitude-gyro
`measurements.
`Both RIGs and TRGs required enlarged high-rate torquer
`assemblies and associated precision electronics to precess the
`spinning rotors at high angular rates [11,16]. Strapdown RIG pivots,
`
`SPIN AXIS
`
`OPTICAL PICKOFF MEASURES
`ANGULAR ORIENTATION OF
`CASE RELATIVE TO SPIN AXIS
`
`CASE MOUNTED ELECTRODES
`GENERATE ELECTROSTATIC
`FIELD FOR FREE-ROTOR
`SUSPENSION
`
`SPIN COIL
`
`ROTOR
`
`VACUUM
`MAINTAINED
`BETWEEN FREE
`ROTOR AND CASE
`
`ION GETTER PUMP
`
`SCRIBE PATTERN
`ETCHING REFLECTS
`LIGHT GENERATED BY
`OPTICAL PICKOFF
`
`Fig. 7 Honeywell hollow-shell rotor ESG.
`
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`SAVAGE
`
`641
`
`originally used for delicate unstressed centering of the floatation
`supported rotor assembly (Fig. 5), now had to withstand severe lateral
`bearing loads under output axis rotation, thereby distorting output
`axis integration response. Gas bearings for improved RIG spin-motor
`reliability required stiffening for high dynamic angular rate, increas-
`ing startup stiction under successive on–off cycles. The inevitable
`result was added bias and scale-factor error from mechanical stresses
`under dynamic angular rates, high-rate torquer heating, and on–off/
`cooldown, with long-term bias stability remaining a major limita-
`tion in aircraft strapdown INS applications. ESG modification for
`strapdown operation introduced new error mechanisms as a func-
`tion of vehicle attitude; precision wide-angle readouts generated
`scale-factor error, and increased bias error was induced from
`suspension-field forces acting at different case-mounted-electrode/
`rotor orientations, both effects requiring complicated calibration
`procedures and increased production cost. With these changes,
`aircraft strapdown INS requirements also dictated two orders-of-
`magnitude improvement
`in RIG/TRG torquer-loop scale-factor
`accuracy (compared with gimbaled system requirements), and
`2 arcseconds ESG readout accuracy for arbitrary rotor/case attitude.
`By the end of the 1960s, strapdown conventional gyros were
`incapable of meeting general aircraft strapdown INS requirements
`without performance specification relief or limiting usage to appli-
`cation areas where reduced INS accuracy was acceptable (e.g.,
`operation with inertial aids to mitigate navigational error buildup
`or limiting strapdown technology to lower angular rate applications).
`Additionally, as with gimbaled system conventional gyros, strap-
`down versions required active temperature control (with heaters) to
`stabilize thermally sensitive performance parameters at compensa-
`tion-calibrated values. The associated warm-up requirement
`precluded a desirable faster INS reaction time, the time from system
`turn-on to entry into the navigation mode (including platform initial
`north alignment determination). Ironically, accelerometer thermally
`sensitive bias trending during heading alignment was actually the
`warm-up time determining factor. For conventional gyros and
`accelerometers mounted in close proximity, gyro heating becomes a
`dominant accelerometer thermal input driver, requiring acceler-
`ometer
`temperature control
`(and warm up)
`for heated gyro
`compatibility.
`
`RLG Development
`To directly meet the Fig. 3 strapdown rate-gyro challenge from
`a different perspective, the RLG was introduced in 1963. Unlike
`traditional angular-momentum gyros whose operation is based on
`the Newtonian inertial properties of rotating mass, the operating
`principle for the RLG is based on the relativistic properties of optical
`standing waves generated by oppositely directed laser beams
`contained in a closed optical path ([11,12], Chap. 13 in [14], Chap. 8
`in [15], [17,18]).
`Figure 8 depicts the basic operating elements in an RLG: a closed
`optical cavity containing two independent beams of light, both of
`the same single frequency. The beams travel continuously between
`the reflecting surfaces of the cavity in a closed optical path, one in the
`clockwise direction and the other counterclockwise, each occupying
`the same physical space. The light beams are sustained by the lasing
`action of a helium-neon gas discharge within the optical cavity. The
`reflecting surfaces are dielectric mirrors designed to selectively
`reflect the frequency associated with the particular helium-neon
`transition being used. The counter-rotating beams combine into a
`standing wave of light that remains inertially fixed as the gyro cavity
`rotates. A small fraction of each beam escapes the cavity, one
`reflected through a corner prism, both recombined on photodiode
`readout detectors. The corner prism is designed to produce a small
`angle between the recombining beams, thereby creating an optical
`interference fringe pattern on the photodiodes, each fringe equivalent
`to a magnified image of the standing wave within the gyro. As
`the cavity rotates,
`the fringes traverse the diodes, each fringe
`corresponding to half the laser wavelength λ for the equivalent
`angular rotation of the gyro (λ∕D rad), where D is the average gyro
`width (e.g., 2 arcseconds of gyro rotation per fringe passage).
`
`READOUT
`PHOTODIODES
`
`CORNER
`PRISM
`
`DIELECTRIC
`MIRROR
`SUBSTRATE
`
`CLOCKWISE AND
`COUNTERCLOCKWISE
`LASER BEAMS
`
`DIELECTRIC
`MIRROR
`SUBSTRATE
`
`POWER
`DETECTOR
`PHOTODIODE
`
`DIELECTRIC
`MIRROR
`SUBSTRATE
`
`HE/ NE GAS
`DISCHARGE
`
`PATH
`LENGTH CONTROL
`PIEZOELECTRIC
`TRANSDUCER
`
`MIRROR
`COATINGS
`
`Fig. 8 RLG operating elements.
`
`Photodiode readout logic generates digital output pulses for each
`fringe quarter-wave passage. Two diodes are used, separated from
`each other by one-quarter of a photodiode-sensed fringe, so that
`resulting diode sinusoidal outputs are 90 deg phase separated.
`Comparison between diode outputs determines the direction of
`rotation, positive or negative, depending on whether one diode output
`is leading or lagging the other.
`Laser stands for light amplification by the stimulated emission of
`radiation. In an RLG, the emission process is provided by the helium/
`neon gas discharge that generates light waves at a discrete atomic
`transition wavelength when impacted by photons of the same
`wavelength. For lasing to occur, the RLG mirrors must reflect the
`emitted light around the closed beam path, so that it returns in phase
`with itself. The beam intensity will then be amplified into resonance
`until the light emitted (“gain”) balances all cavity losses, which then
`also maximizes beam power. The gain is set by the current magnitude
`applied to the gas discharge. To satisfy the return-in-phase condition
`for lasing, the beam path length must be controlled to an integral
`multiple of discrete lengths, corresponding to the optical wavelength
`of the helium/neon discharge. This is achieved implicitly in the RLG
`by mirror adjustment to a position for peak beam power. Beam power
`is measured by a photodiode power detector attached to one of
`the mirror substrates (Fig. 8). Piezoelectric transducers attached to
`the outer mirror substrates (Fig. 8) provide the means for actively
`controlling mirror position, enabling minute adjustments by an
`electrically applied voltage. The control voltage is generated in
`closed-loop fashion to sustain maximum output from the power
`detector. In addition to enabling lasing, the path-length control
`process also produces two very important angular rate sensing
`operational benefits: 1) stabilization of RLG performance parameters
`and 2) elimination of path-length changes from gyro block thermal
`expansion.
`The RLG concept bypassed many of the conventional gyro
`design issues. Piezoelectric transducer path-length control elimi-
`nated the principle source of thermal error sensitivity without
`requiring direct temperature control. This important characteristic
`eliminated the warm-up time penalty experienced with conventional
`gimbaled inertial systems. Stable high scale-factor accuracy, the
`critical performance parameter for strapdown applications, was an
`inherent quality, independent of bias-producing mechanisms. Simple
`wide-angle readouts could be used with minimum impact on
`accuracy (because 360 deg laser signal detector scaling corresponds
`to arcseconds actual gyro input axis rotation). However, the RLG has
`a unique “lock-in” error mechanism of its own that cannot be elimi-
`nated and had to be circumvented before RLGs could be considered
`for high-accuracy INS application [19].
`RLG lock-in is caused by imperfections in the laser reflecting
`mirrors and optical elements placed in the beam path that introduce
`clockwise/counterclockwise beam energy coupling. Beam coupling
`
`Downloaded by Reprints Desk Access on May 24, 2021 | http://arc.aiaa.org | DOI: 10.2514/1.60211
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`Niantic's Exhibit No. 1018
`Page 005
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`SAVAGE
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`tends to force the standing wave to follow the cavity rotation (losing
`its nonrotating inertial property) and produces a “dead zone” in gyro
`output at low input rate (e.g., 0.03 deg ∕s, depending on mirror
`quality). The general solution is to artificially bias the gyro input to
`avoid dwelling in the lock-in region. RLG design activities in the
`1960s centered on competing methods for overcoming lock-in, the
`principle groups involved being Sperry Gyroscope Division of
`Sperry Rand, Great Neck, New York; Hamilton Standard Division of
`United Technologies, Windsor Locks, Connecticut; and Honeywell
`Aerospace and Defense Group, Minneapolis, Minnesota.
`Sperry used a modular RLG design approach in which the helium/
`neon laser sustaining plasma was contained in gain tubes inserted in
`the optical beam path, with reflecting mirrors for beam circulation
`mounted externally (Chap. 13 in [14], [20–22]). Separating the
`plasma from direct mirror contact eliminated the potential for long-
`term mirror dielectric coating degradation. Biasing was achieved by
`applying a cyclic saturating magnetic field to a ferromagnetic coating
`deposited on the reflecting mirror substrates, enabling magnetically
`controlled path-length shifts (the equivalent of input bias) through
`the transverse magneto-optical Kerr effect (“magnetic