Fail safe augmentation system

• US 4,095,763 A
• Filed: 02/25/1977
• Issued: 06/20/1978
• Est. Priority Date: 02/25/1977
• Status: Expired due to Term

• Current Assignee: Bell Helicopter Textron Incorporated (Textron Inc)
• Original Assignee: Textron Inc
• Inventors: Builta, Kenneth E.
• Primary Examiner(s): Blix, Trygve M.
• Assistant Examiner(s): Eisenzopf, Reinhard J.
• Application Number:
  US05/771,898
• Time in Patent Office: 480 Days
• Field of Search: 244/78, 244/83 G, 244/194, 318/563, 318/564, 318/565, 318/566, 91/363 A
• US Class Current: 244/194
• CPC Class Codes: G05D 1/0077 using redundant signals or controls

Abstract

A fail safe aircraft control system wherein duplicate control loops are employed to produce identical control signals for the system actuator. Output signals from the two loops are compared to detect when the outputs differ by a preset limit as an indication of a component failure. When such a failure is detected the system actuator is deactivated thus preventing the failed augmentation system from affecting control of the aircraft.

Claims

1. An aircraft control system, comprising:

input means for introducing a manual commnd to position a control surface,

stability control augmentation means comprising;

(a) an actuator connected through linkage to the control surface,

(b) at least two drive circuits for said acutator, each comprising;

(i) a position sensor responsive to said input means and generating a position signal, and

(ii) a drive amplifier responsive to the position signal and providing an output to drive said actuator,

(c) means responsive to the outputs of the drive amplifiers to produce a difference signal, and

(d) means responsive to the difference signal to disable said actuator.

2. An aircraft control system as recited in claim 1 which includes means for detecting when said difference signal exceeds a predetermined threshold.

3. An aircraft control system as recited in claim 2, which includes means for disabling said actuator when said difference signal exceeds said predetermined threshold.

4. An aircraft control system as recited in claim 1 wherein each of said drive circuits includes:

(a) an actuator position sensor, and

(b) a summing unit for summing the outputs of said position sensor and said actuator position sensor.

5. An aircraft control system as recited in claim 4 which includes means for adding said difference signal to one of said summing units.

6. An aircraft control system as recited in claim 1 which includes means for locking the position of said actuator after said actuator has been disabled.

7. An aircraft control system, comprising:

input means for introducing a manual command to position a control surface,

force feel means comprising;

(a) an actuator connected through linkage to said input means,

(b) at least two drive circuits for said actuator each comprising;

(i) a force sensor responsive to said input means and generating a force signal, and

(ii) a drive amplifier responsive to the force signal and providing an output to drive said actuator,

(c) means responsive to the outputs of the drive amplifiers to produce a difference signal, and

(d) means responsive to the difference signal to disable said actuator.

8. An aircraft control system as recited in claim 7 which includes means for detecting when said difference signal exceeds a predetermined threshold.

9. An aircraft control system as recited in claim 8 which includes means for disabling said actuator when said difference signal exceeds said predetermined threshold.

10. An aircraft control system as recited in claim 7 wherein each of said drive circuits includes:

(a) an actuator position sensor, and

(b) a summing unit for summing the outputs of said force sensor and said actuator position sensor.

11. An aircraft control system as recited in claim 10 which includes means for adding said difference signal to one of said summing units.