INERTIALLY-AIDED AIR DATA COMPUTER ALTITUDE RATE

• US 20180292430A1
• Filed: 04/10/2017
• Published: 10/11/2018
• Est. Priority Date: 04/10/2017
• Status: Active Grant

• Current Assignee: Rosemount Aerospace, Inc. (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Rosemount Aerospace, Inc. (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Ell, Todd Anthony
• Granted Patent: US 10,634,692 B2
• Time in Patent Office: Days
• CPC Class Codes: G01C 21/165 combined with non-inertial navigation instruments
• CPC Class Codes: G01C 5/06 by using barometric means
• CPC Class Codes: G01P 3/62 Devices characterised by the determination or the variation of atmospheric pressure with height to measure the vertical components of speed measuring pressure in general G01L
• CPC Class Codes: G05D 1/0607 specially adapted for aircraft

Abstract

An air data computer configured to be installed on an aircraft includes an inertial sensor assembly having a plurality of accelerometers and a plurality of rate gyroscopes. The air data computer is configured to: determine a pressure altitude of the aircraft based on measured pressure of the airflow about the exterior of the aircraft; determine an estimated attitude of the aircraft based on rotational rate sensed by the plurality of rate gyroscopes; and determine a vertical acceleration of the aircraft based on the estimated attitude of the aircraft and the acceleration sensed by the plurality of accelerometers. The air data computer is further configured to blend the vertical acceleration and the pressure altitude using a complementary filter to produce a blended altitude rate that is output to consuming systems.

Claims

1. An air data computer configured to be installed on an aircraft, the air data computer comprising:

an inertial sensor assembly comprising a plurality of accelerometers and a plurality of rate gyroscopes;

an air data parameter module configured to determine a pressure altitude of the aircraft based on measured pressure of airflow about an exterior of the aircraft;

an attitude determination module configured to determine an estimated attitude of the aircraft based on rotational rate sensed by the plurality of rate gyroscopes; and

an altitude rate blending module configured to;

determine a vertical acceleration of the aircraft based on the estimated attitude of the aircraft and acceleration sensed by the plurality of accelerometers;

blend the vertical acceleration and the pressure altitude using a complementary filter to produce a blended altitude rate; and

output the blended altitude rate.

2. The air data computer of claim 1,

wherein the complementary filter is configured to:

integrate the vertical acceleration to produce an accelerometer-based vertical velocity of the aircraft;

differentiate the pressure altitude to produce a pressure-based altitude rate of the aircraft;

filter the accelerometer-based vertical velocity to remove low frequency signal and noise to produce a high-pass filtered accelerometer-based vertical velocity;

filter the pressure-based altitude rate to remove high frequency signal and noise to produce a low-pass filtered pressure-based altitude rate; and

blend the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate to produce the blended altitude rate of the aircraft.

3. The air data computer of claim 2,

wherein the complementary filter is configured to blend the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate by summing the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate.

4. The air data computer of claim 2,

wherein the complementary filter is a multi-input single-output (MISO) filter that takes the vertical acceleration and the pressure altitude as inputs and produces the blended altitude rate as an output.

5. The air data computer of claim 2,

wherein the complementary filter is configured to filter the accelerometer-based vertical velocity to remove the low frequency signal and noise using a high-pass filter;

and

wherein the complementary filter is configured to filter the pressure-based altitude rate to remove the high frequency signal and noise using a low-pass filter.

6. The air data computer of claim 5,

wherein a cutoff frequency of the high-pass filter matches a cut-off frequency of the low-pass filter.

7. The air data computer of claim 1,

wherein the altitude rate blending module is configured to determine the vertical acceleration of the aircraft based on the estimated attitude of the aircraft and the acceleration sensed by the plurality of accelerometers by multiplying the acceleration sensed by the plurality of accelerometers and a direction-cosine matrix having direction angles corresponding to roll and pitch of the aircraft.

8. The air data computer of claim 1,

wherein the plurality of accelerometers comprises three accelerometers, each configured to sense acceleration along one of three mutually-orthogonal axes;

and

wherein the plurality of rate gyroscopes comprises three rate gyroscopes, each configured to sense rotational rate along one of the three mutually-orthogonal axes.

9. The air data computer of claim 1, wherein the air data parameter module is further configured to determine a true airspeed of the aircraft based on the measured pressure of airflow about the exterior of the aircraft, the air data computer further comprising:

an inertial compensation and correction module; and

a Kalman estimator module configured to;

determine a change in integrated acceleration of the air data computer over a time interval based on acceleration sensed by the plurality of accelerometers and rotational rate sensed by the plurality of rate gyroscopes;

determine a set of error correction values based on a difference between the change in the integrated acceleration of the air data computer and a change in the true airspeed of the air data computer; and

provide the set of error correction values to the inertial sensor compensation and correction module;

wherein the inertial sensor compensation and correction module is further configured to;

apply the set of error correction values to each of the acceleration sensed by the plurality of accelerometers and the rotational rate sensed by the plurality of rate gyroscopes to produce an error-corrected acceleration and error-corrected rotational rate;

provide the error-corrected rotational rate to the attitude determination module for determination of the estimated attitude of the aircraft; and

provide the error-corrected acceleration to the altitude rate blending module for determination of the vertical acceleration of the aircraft.

10. The air data computer of claim 9,

wherein the Kalman estimator module is configured to determine the set of error correction values via an extended Kalman filter that utilizes the difference between the change in the integrated acceleration of the air data computer and the change in the true airspeed of the air data computer as input and produces the set of error correction values as output.

11. A method comprising:

sensing acceleration of an air data computer configured to be installed on an aircraft along a plurality of axes via a plurality of accelerometers of an inertial sensor assembly of the air data computer;

sensing rotational rate of the air data computer along the plurality of axes via a plurality of rate gyroscopes of the inertial sensor assembly;

determining a pressure altitude of the aircraft based on measured pressure of airflow about an exterior of the aircraft;

determining an estimated attitude of the aircraft based on the rotational rate sensed by the plurality of rate gyroscopes;

determining a vertical acceleration of the aircraft based on the estimated attitude of the aircraft and the acceleration sensed by the plurality of accelerometers;

blending the vertical acceleration and the pressure altitude using a complementary filter to produce a blended altitude rate; and

outputting the blended altitude rate.

12. The method of claim 11,

wherein blending the vertical acceleration and the pressure altitude using the complementary filter comprises:

integrating the vertical acceleration to produce an accelerometer-based vertical velocity of the aircraft;

differentiating the pressure altitude to produce a pressure-based altitude rate of the aircraft;

filtering the accelerometer-based vertical velocity to remove low frequency signal and noise to produce a high-pass filtered accelerometer-based vertical velocity;

filtering the pressure-based altitude rate to remove high frequency signal and noise to produce a low-pass filtered pressure-based altitude rate; and

blending the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate to produce the blended altitude rate of the aircraft.

13. The method of claim 12,

wherein blending the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate comprises summing the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate.

14. The method of claim 12,

wherein integrating the vertical acceleration, differentiating the pressure altitude, filtering the accelerometer-based vertical velocity, filtering the pressure-based altitude rate, and blending the high-pass filtered accelerometer-based vertical velocity and the low-pass filtered pressure-based altitude rate are performed using a multi-input single-output (MISO) filter that takes the vertical acceleration and the pressure altitude as inputs and produces the blended altitude rate as an output.

15. The method of claim 12,

wherein filtering the accelerometer-based vertical velocity to remove the low frequency signal and noise comprises filtering the accelerometer-based vertical velocity using a high-pass filter;

and

wherein filtering the pressure-based altitude rate to remove the high frequency signal and noise comprises filtering the pressure-based altitude rate using a low-pass filter.

16. The method of claim 15,

wherein a cutoff frequency of the high-pass filter matches a cut-off frequency of the low-pass filter.

17. The method of claim 11,

wherein determining the vertical acceleration of the aircraft based on the estimated attitude of the aircraft and the acceleration sensed by the plurality of accelerometers comprises multiplying the acceleration sensed by the plurality of accelerometers and a direction-cosine matrix having direction angles corresponding to roll and pitch of the aircraft.

18. The method of claim 11,

wherein sensing the acceleration along the plurality of axes via the plurality of accelerometers comprises sensing the acceleration along three mutually-orthogonal axes via three accelerometers;

and

wherein sensing the rotational rate along the plurality of axes via the plurality of rate gyroscopes comprises sensing the rotational rate along the three mutually-orthogonal axes via three rate gyroscopes.

19. The method of claim 11, further comprising:

determining a true airspeed of the aircraft based on the measured pressure of airflow about the exterior of the aircraft;

determining a change in integrated acceleration of the air data computer over a time interval based on the acceleration sensed by the plurality of accelerometers and the rotational rate sensed by the plurality of rate gyroscopes;

determining a set of error correction values based on a difference between the change in the integrated acceleration of the air data computer and a change in the true airspeed of the air data computer; and

applying the set of error correction values to each of the acceleration sensed by the plurality of accelerometers and the rotational rate sensed by the plurality of rate gyroscopes.

20. The method of claim 19,

wherein determining the set of error correction values comprises determining the set of error correction values via an extended Kalman filter that utilizes the difference between the change in the integrated acceleration of the air data computer and the change in the true airspeed of the air data computer as input and produces the set of error correction values as output.