GB2271422A

GB2271422A - Compensated inertial guidance system

Info

Publication number: GB2271422A
Application number: GB9320517A
Authority: GB
Inventors: Norman Frederick Watson
Original assignee: GEC Ferranti Defence Systems Ltd; GEC Marconi Avionics Holdings Ltd
Current assignee: Leonardo MW Ltd
Priority date: 1992-10-08
Filing date: 1993-10-05
Publication date: 1994-04-13
Anticipated expiration: 2013-10-05
Also published as: GB2271422B; GB9320517D0; GB9221165D0

Abstract

An inertial guidance system comprising an inertial sensor assembly (I.S.A.) (1) the I.S.A. comprising a casing (1a) having a plurality of motion sensors (1b to 1g) housed therein. The I.S.A. (1) is flexibly connected to a vehicle (3b) by means of anti-vibration mounts (4) and sensing means (5a to 7b) is provided externally of the I.S.A. (1) to detect relative movement between the I.S.A. (1) and the vehicle (3b) so as to provide compensation of the guidance data. <IMAGE>

Description

COMPENSATED INERTIAL GUIDANCE SYSTEM The present invention relates to an inertial guidance system and particularly, though not exclusively, to a flight control system utilising an inertial guidance system.

Previously proposed inertial guidance systems comprise a plurality of motion sensors, typically accelerometers and gyroscopes, housed in a casing which is secured to a vehicle such as an aircraft. The sensors measure acceleration and angular movement relative to a fixed datum point relative to the casing, usually within it. As the casing is secured to the vehicle, the motion of the vehicle can be calculated. Such systems are known as strap down systems.

However, in the case of a vehicle such as an aircraft, thermal and mechanical shocks are encountered which are not due to course changes, or changes in the position of the aircraft, but originate within the aircraft, such as for example vibration due to engine operation. To overcome this problem it has been proposed to isolate the guidance sensors from shock by flexibly mounting the casing on anti-vibration mountings which can absorb some of the shock. Some compromise however must be reached between isolating the casing from shock and transmitting the effects of motion changes which are to be measured. For example if the mountings are too soft, the casing will be allowed too much freedom at low frequency and will not be responsive to the higher frequency motion which must be measured for flight control.

If on the other hand the mountings are too hard, the vibration encountered in the aircraft in flight would be detected by the sensors and may damage these inertial sensors.

The present invention has arisen in the attempt to provide a system which obviates or mitigates these problems.

In accordance with the present invention, there is provided an inertial guidance system comprising an inertial sensor assembly (I.S.A.) the l.S.A. comprising a casing having a plurality of motion sensors housed therein, the l.S.A. being flexibly connected to a vehicle, characterised in that sensing means are provided externally of the l.S.A.

to detect relative movement between the l.S.A. and the vehicle for providing compensation of the inertial guidance system outputs.

The sensing means can be mounted either on the casing of the l.S.A. or fixed relative to the vehicle. For accuracy, it is preferred that sufficient sensing means are provided for determining relative movement for each degree of freedom of the casing.

Preferably, six sensors are provided for determining relative movement in each of the X, Y, and Z directions. It is preferred that the sensors are non-contact sensors in order that no significant influence should be exerted on the motion of the I.S.A. It is also preferred that the sensors detect higher frequency displacement than the motion sensors within the casing.

Such a system is particularly applicable for a flight control system for an aircraft.

The present invention will now be described, by way of example only, with reference to the accompanying drawings in which: Figure 1 is a schematic representation of one embodiment of the invention and Figure 2 is a representation of the compensation circuitry of the invention.

Referring to Figure 1 there is shown an inertial sensor assembly (I.S.A.) generally denoted as a wire frame 1 in the Figure. I.S.A.'s are well known in the art as described in a book by K.R. Britting (1971) entitled 'Avionics Navigation Systems' and generally comprise a casing 1 a on which motion sensors, typically accelerometers and gyroscopes, are mounted. In the present example three ring laser gyroscopes (R.L.G.'s) Ib to id and three accelerometers le to lg are mounted on the casing la in mutually orthogonal directions. The R.L.G.'s detect angular motion about the axes x, y and z and the inertial accelerometer's acceleration along the axes x, y and z of datum 2 as illustrated.

The l.S.A. 1 is mounted on a housing 3 by means of anti-vibration mountings 4 which allow a limited degree of relative motion. Anti-vibration mountings are well known in the art as described in the book by William T. Thomson (1981) entitled 'Theory of vibration'. The housing 3 is securely fastened, by a suitable fixing means 3a, to the floor 3b of the aircraft in which the inertial guidance system is to operate.

Three pairs of sensors 5a and 5b, 6a and 6b and 7a and 7b are mounted on the housing 3 to detect displacement of the l.S.A. 1 relative to the housing 3. The sensor pairs 5a/5b, 6a/6b and 7a/7b are arranged so that displacement of the l.S.A.

in each of the six possible degrees of freedom can be determined. The six degrees of freedom are shown as arrows on the datum 2 and comprise linear displacement in the x, y or z directions and angular displacement about these axes. In the embodiment shown the sum of the outputs of sensor pair 5a, 5b corresponds to the linear displacement of the I.S.A. along the x axis whilst the difference between the outputs corresponds to the angular displacement of the l.S.A. about the z axis.

Similarly sensor pair 6a and 6b will be responsive to linear displacement in the y axis direction and angular displacement about x and so forth.

The sensors are non-contact sensors such that they do not affect the motion of the l.S.A. in any significant way and may utilise magnetic or electric fields or light beams to monitor the position of the l.S.A. In the present example Eddy current sensors are used, though Linear Variable Differential Transformers (L.V.D.T.'s) or any other suitable sensor such as those described in the book by S.T. Smith and D.G.

Chetwynd (1992) entitled 'Foundations of Ultraprecision Mechanism Design' may be utilised.

In order to separate the various components from the outputs of the sensor pairs the output from respective pairs of sensors are connected to sum and difference amplifying circuit means 8, 9 and 10. The sum and difference amplifying circuit means 8, 9 and 10 each have two respective outputs a and b in which the first, a, corresponds to the linear displacement in a given direction and the second, b, to angular displacement. Within the present embodiment output 8a represents the linear displacement of the I.S.A. in the x direction and output 8b the angular displacement in the z direction.

Referring to Figure 2 the analogue outputs 8a to 1 Ob are converted to digital signals by an analogue to digital (A/D) converter 11. The output signals from the R.L.G.'s ib to Id and inertial accelerometers le to lg are also converted to digital signals by the ND converter 11. The digital data from the A/D converter is then supplied to a digital computer 12, which uses the data from the sensors 5a to 7b to compensate for induced errors in the output of the R.L.G.'s and accelerometers of the I.S.A.. Examples of induced errors that the system is particularly suited to compensate for are dither reaction of the R.L.G., dither induced coning and sculling motion and changes in the damping characteristics of the anti-vibration mountings. Each of these examples will now be described further.

In inertial guidance systems it is often desirable to utilise at least one R.L.G.

These devices generally include lock-in compensation which operates by periodically rotating or "dithering" the R.L.G. This in turn causes periodic forces to be exerted on the casing 1 a on which the R.L.G. is mounted causing movement of the casing known as "dither reaction". R.L.G.'s are typically dithered at frequencies around 400Hz a frequency at which the R.L.G. may not measure accurately the motion of the casing la. In addition this frequency may also lie outside the normal operating bandwidth of the inertial accelerometers 1e to 19. By use of the sensors 5a to 7b, which are able to detect these dither frequency motions, movement of the l.S.A. relative to the housing 3 can be combined with the R.L.G. and accelerometer outputs and dither reaction compensated for.

Strap down navigation systems using dithered R.L.G.'s may be susceptible to long term drift errors caused by inadequate integration accuracy in the presence of dither-induced coning and sculling motion. Coning motion compensation algorithms are often used to correct for these drift errors and depend on accurate measurements from the R.L.G.'s of the l.S.A. motion at dither frequency.

The removal of the error resulting from the relative motion due to dither of the R.L.G. block with respect to the l.S.A. is difficult to achieve. Notch filtering the output from the R.L.G. is an inadequate solution since it also removes the true I.S.A. motion at dither frequency. Other techniques such as relative displacement sensing or "optical compensation" are both costly and complex.

The additional sensors 5a to 7b referred to in this invention provide a simple relatively inexpensive means of providing compensation for dither induced coning and sculling motion. The outputs from the R.L.G. and the inertial accelerometers can be low pass filtered to remove the component at the dither frequency and then combined with data from the sensors 5a to 7b which has been high pass filtered. Use of sensors able to detect movement at the dither frequency of the R.L.G., such frequencies possibly being outside the normal operating bandwidth of the inertial accelerometers, allows the use of less accurately dither motion compensated R.L.G.'s and lower bandwidth accelerometers, offering a saving in cost.

Over periods of time, anti-vibration mountings can change shape or stiffness and so the position of the l.S.A. within the aircraft might change somewhat. Position changes such as these can be detected by the external sensors 5a to 7b and allowed for in calculations such that it is not necessary for the mountings to remain in a stable condition throughout their life.

The additional sensors 5a to 7b may be used to detect the relative motion of the l.S.A. with respect to the vehicle mounting frame. This data may be used to compensate for the attenuation and phase shifts caused by the vibration isolators, thus allowing wide bandwidth data to be produced with low phase shifts (important for flight control systems) while still allowing relatively low frequency vibration isolators to be used.

By providing for this compensation, an inertial guidance system can be used in flight control systems with greater accuracy. The invention is not limited to the embodiment shown. For example the displacement sensors 5a to 7b could be mounted on the casing 1 a or incorporated into the anti-vibration mountings 4.

Claims

1. An inertial guidance system comprising an inertial sensor assembly (I.S.A.) the l.S.A. comprising a casing having a plurality of motion sensors housed therein, the l.S.A. being flexibly connected to a vehicle, characterised in that sensing means are provided externally of the l.S.A. to detect relative movement between the l.S.A. and the vehicle for providing compensation of the inertial guidance system output.

2. A system as claimed in claim 1, wherein the sensing means are mounted on the casing.

3. A system as claimed in claim 1, wherein the sensing means are fixed relative to the vehicle.

4. A system as claimed in any preceding claim wherein sufficient sensing means are provided for determining relative movement for each degree of freedom of the l.S.A.

5. A system as claimed in claim 4, comprising six sensors for determining relative movement in each of the X, Y and Z directions.

6. A system as claimed in any preceding claim, wherein the sensors detect higher frequency displacement than the motion sensors within the casing.

7. A system as claimed in claim 4, 5 or 6, wherein the sensors are non-contact sensors which exert no significant influence on the motion of the I.S.A.

8. A flight control system for an aircraft comprising a system as claimed in any preceding claim.

9. A system which is substantially as herein described with reference to the accompanying drawings.