Strapped down attitude and heading reference system for aircraft employing skewed axis two-degree-of-freedom rate gyros

• US 4,212,443 A
• Filed: 05/18/1978
• Issued: 07/15/1980
• Est. Priority Date: 05/18/1978
• Status: Expired due to Term

• Current Assignee: Honeywell Incorporated (Honeywell International Inc.)
• Original Assignee: Sperry Corporation
• Inventors: Duncan, Damon H., Klemes, Martin S.
• Primary Examiner(s): Eisenzopf, Reinhard J.
• Application Number:
  US05/907,228
• Time in Patent Office: 789 Days
• Field of Search: 33/321, 33/322, 33/323, 33/366, 73/178 R, 73/504, 73/505, 73/516 LM, 74/5 F, 74/5.6 D, 74/5.34, 74/5.37, 74/5.47, 244/3.2, 244/3.21, 244/79, 244/165, 244/175-177, 318/648, 318/649, 318/651, 364/434, 364/453, 364/459
• US Class Current: 244/177
• CPC Class Codes: G01C 19/42 for indicating rate of turn; for integrating rate of turn
• CPC Class Codes: G01C 21/166 Mechanical, construction or arrangement details of inertial navigation systems
• CPC Class Codes: G01C 21/183 Compensation of inertial measurements, e.g. for temperature effects

Abstract

A set of two two-degrees-of-freedom rate gyroscopes and three linear accelerometers are assembled in a single module adapted to be mounted within a single aircraft electronics control unit, the unit comprising a strapped-down attitude and heading reference system. The module base provides a common keyed support for the two pre-calibrated gyros and three accelerometers in intimate mechanical and thermal association. The two gyros are oriented in the base and the base oriented in the aircraft so that the spin axis of one gyro is oriented parallel to the aircraft Z axis and that of the other gyro parallel to the aircraft Y axis while the gyro pickoffs and torquers (input and output axes, respectively) are rotated or skewed forty-five degrees about the spin axes to positions such that the input and output axes lie along the slant heights of a forty-five degree half angle right circular cone, the axis of which lies along the aircraft X axis. Thus a single module provides greatly simplified rate sensing about the aircraft primary axes with correspondingly simplified fault detection of the rate sensors.

Claims

1. In a unitary strapped down attitude and heading reference system for an aircraft having first, second and third primary rotation axes and including a system enclosure, said system comprising,

(a) an inertial reference assembly module including

(i) a base support member having at least first and second mutually orthogonal reference surfaces and adapted to be fixed within said enclosure with said surfaces disposed in first and second orthogonal planes normal respectively to first and second primary aircraft axes,

(ii) first and second two-degree-of-freedom rate gyroscopes, each gyroscope having a housing including a rotor and a motor for spinning said rotor about a spin axis and pick-off and torquer means cooperative with said rotor and defining two mutually orthogonal input axes normal to said spin axis and means responsive to said pick-off means for supplying two output signals proportional respectively to aircraft rotation rates about said two input axes,

(iii) means for aligning and securing said first and second gyroscope housing relative to said first and second base support reference surfaces respectively such that their spin axes are respectively normal thereto and parallel to said first and second aircraft axes and their input axes are respectively skewed at an angle on the order of forty-five degrees relative to the third primary aircraft axis, said gyroscope input axes thereby lying parallel to the slant height of a cone having a half angle on the order of forty-five degrees, the axis of said cone being parallel to said third aircraft axis, the difference between the output signals of the first of said gyroscopes corresponding to components of aircraft rotation rates about one of said first and second axes and the difference between the output signals of said second of said gyroscopes corresponding to components of aircraft rotation rates about the other of said first and second axes, and the output signal components of both gyroscopes corresponding to aircraft rotation rates about said third axis, said last-mentioned signal components being normally equal,

(b) computer means responsive to said output signals of said inertial reference assembly module including

(i) means responsive to the signal outputs of said first gyroscope for providing a signal corresponding to the difference therebetween and scaling means for modifying the amplitude thereof in accordance with said skew angle whereby to provide a resultant signal proportional to aircraft rotation rate about said first aircraft axis,

(ii) means responsive to the signal outputs of said second gyroscope for providing a signal corresponding to the difference therebetween and scaling means responsive to said difference signal for modifying the amplitude thereof in accordance with said skew angle whereby to provide a resultant signal proportional to aircraft rotation rate about said second aircraft axis, and

(iii) means responsive to the signal outputs of said first and second gyroscopes respectively for providing first and second signals corresponding respectively to the sum thereof and scaling means responsive to each of said sum signals for modifying the amplitudes thereof in accordance with said skew angle whereby to provide two resultant signals proportional to aircraft rotation rates about said third aircraft axis, and means for averaging the values of said two resultant signals.

2. The system as set forth in claim 1 wherein said first, second, and third aircraft axes correspond to the aircraft pitch, yaw and roll axes respectively.

3. The system as set forth in claim 1 wherein said first, second, and third aircraft axes correspond to the aircraft pitch roll and yaw axes respectively.

4. The system as set forth in claim 1 wherein said first, second, and third aircraft axes correspond to the aircraft roll, yaw and pitch axes, respectively.

5. The system as set forth in claim 1 further including

(a) means responsive to said two sum signals for providing a further signal proportional to the difference therebetween, and

(b) monitoring means including filtering means responsive to said difference signal for providing a malfunction signal when the value thereof exceeds a predetermined value.

6. The inertial reference assembly as set forth in claim 1 further including:

an electronics module connected with each of said gyroscopes, each electronics module including means for defining the calibration characteristics unique to its respective gyroscope, said electronics module and gyroscope constituting a removable and replaceable integral subassembly, and

wire harness and connector means on said base member for connecting each said subassembly with said computer means, whereby said gyroscopes may be removed and replaced without recalibration of the entire inertial reference assembly module.