Total energy based flight control system

• US 6,062,513 A
• Filed: 09/14/1998
• Issued: 05/16/2000
• Est. Priority Date: 09/14/1998
• Status: Expired due to Term

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Lambregts, Antonius A.
• Primary Examiner(s): Eldred, J. Woodrow
• Application Number:
  US09/152,732
• Time in Patent Office: 610 Days
• Field of Search: 244/76 R, 244/175, 244/182, 244/191, 244/181, 244/194, 244/195, 701/3, 701/4
• US Class Current: 244/175
• CPC Class Codes: G05D 1/0005 with arrangements to save energy
• CPC Class Codes: G05D 1/0638 by combined action on the pitch and on the motors

Abstract

In a method of flight control in which a thrust command is computed based on the total aircraft energy error relative to flight path and speed control commands, and an elevator command is computed based on the energy distribution error relative to the same flight path and speed control commands, an improvement is provided including an elevator control command in response to a column control input by the pilot. In the short term, the computing establishes a change in flight path angle beyond the sustainable flight path angle at the trim speed for the prevailing thrust condition. In the long term, the computing establishes a change in speed relative to a set reference speed, the speed change being proportional to the column control input. In the long term, the computing establishes a flight path angle equal to a sustainable value for the prevailing thrust condition and the altered speed condition.

Claims

1. In a method of flight control in which a thrust command is computed based on a total aircraft energy error relative to flight path and speed control commands, and an elevator command is computed based on an energy distribution error relative to the same flight path and speed control commands, the improvement comprising:

computing an elevator control command in response to a control column input by the pilot;

the computing including;

(a) establishing in the short term a change in flight path angle beyond the sustainable flight path angle at the trim speed for the prevailing constant thrust condition; and

(b) establishing in the long term a change in speed relative to a set reference speed, the speed change being proportional to the column control input, and a flight path angle equal to a sustainable value for the prevailing constant thrust condition and the altered speed condition.

2. The improvement according to claim 1, wherein computing includes calculating a flight path angle command signal;

the flight path angle command being used only during a condition when the computed thrust command is at a preset upper or lower thrust limit or when a throttle servo is disengaged;

the flight path angle command being developed by pure integration of a signal representative of the pilot'"'"'s control column input when the computed thrust command is within the range between the preset upper and lower thrust limits.

3. The improvement according to claim 1, wherein the total aircraft energy error is computed from a signal representative of the total energy rate error E_Sε

, the total energy rate error being formed as the sum of a signal representative of the flight path angle error γ

.sub.ε

relative to the flight path angle command and a signal representative of an acceleration error V.sub.ε

relative to a longitudinal acceleration command.

4. The improvement according to claim 1, wherein the total energy distribution error is computed from a signal representative of the total energy distribution rate error D.sub.ε

, the total energy distribution rate error being formed as the difference between a signal representative of a flight path angle error γ

.sub.ε

relative to the flight path angle command and a signal representative of an acceleration error V.sub.ε

relative to a longitudinal acceleration command.

5. The improvement according to claim 1, wherein the elevator control command is a function of a signal representative of a flight path angle error signal γ

.sub.ε

;

wherein the flight path angle error signal is formed as a difference between the current flight path angle γ and

a manual commanded flight path angle γ

_C.sbsb.MAN, the γ

_C.sbsb.MAN signal being formed at the output of an integrator which includes an input of a flight path angle rate command signal γ

_C which is proportional to the control column input.

6. The improvement according to claim 5, wherein the flight path angle rate command signal γ

_C is a function of a signal δ

_C representative of a control column displacement multiplied by an incremental load factor Δ

n_Z for a positive control column input.

7. The improvement according to claim 6, wherein the incremental load factor Δ

n_Z is a function of airspeed V and airspeed at a 1 g stall V_STALL in the relation ##EQU30##

8. The improvement according to claim 7, wherein the incremental load factor Δ

n_Z is limited on the low side to a value greater than or equal to zero, and is limited on the high side to a pre-selected value representative of the maximum allowable positive incremental load factor.

9. The improvement according to claim 5, wherein the flight path angle rate command signal γ

_C is a function of a signal δ

_C representative of a control column displacement multiplied by the ratio of gravity g to groundspeed V_G.

10. The improvement according to claim 5, wherein a second input to the integrator is a signal γ

_C.sbsb.MOD provided when the thrust command is at an upper or lower limit;

when the thrust command is at the upper limit, the signal γ

_C.sbsb.MOD being a function of the flight path angle error signal γ

.sub.ε

, current longitudinal acceleration, and a ratio of thrust margin relative to the upper thrust limit over the weight of the aircraft,

when the thrust command is at the lower limit, the signal γ

_C.sbsb.MOD being a function of the flight path angle error signal γ

.sub.ε

, current longitudinal acceleration, and a ratio of thrust margin relative to the lower thrust limit over the weight of the aircraft,

and for all other conditions the signal γ

_C.sbsb.MOD being zero.

11. The improvement according to claim 1, wherein the elevator control command is computed as a function of a signal representative of a flight path angle error signal γ

.sub.ε

;

the flight path angle error signal γ

.sub.ε

being formed as a difference between the current flight path angle γ and

a manual commanded flight path angle γ

_C.sbsb.MAN ;

wherein the γ

_C.sbsb.MAN signal is formed at the output of an integrator which has two inputs, a first input being a flight path angle rate command signal γ

_C which is proportional to the control column input and a second input being a short term adjustment signal γ

_C.sbsb.MOD available when the thrust command is at an upper or lower limit.

12. The improvement according to claim 11, wherein the manual flight path angle command signal γ

_C.sbsb.MAN is an airmass referenced value, a corresponding inertially referenced value being used for display to the pilot on a cockpit instrument display.

13. The improvement according to claim 1, wherein during manual flight control the total aircraft energy error is computed from a signal representative of the total energy rate error E_Sε

, the total energy rate error being formed as a combination of a signal representative of the flight path angle error γ

.sub.ε

, a signal representative of an acceleration error V.sub.ε

, and the flight path angle rate command signal γ

_C multiplied by a gain KGET; and

wherein during manual control flight the total energy distribution error is computed from a signal representative of the total energy distribution rate error D.sub.ε

, the total energy distribution rate error being formed as a combination of a signal representative of a flight path angle error γ

.sub.ε

, a signal representative of an acceleration error V.sub.ε

, the flight path angle rate command signal γ

_C multiplied by a gain KGEE, and the flight path angle rate command signal γ

_C multiplied by a gain KGDP and differentiated.

14. The improvement according to claim 13, wherein the gains KGET and KGEE are determined to yield a pre-selected steady state response lag of flight path angle relative to said flight path angle command for a pilot'"'"'s stick or column (δ

_C) input, for a condition where thrust command remains within the range of the minimum and maximum thrust limit.

15. The improvement according to claim 14, wherein the gain factor KGEP is determined such that an overall throughput gain of the flight path angle rate command signal to the elevator in the signal path including KGEP is approximately equal to a pitch rate feed feedback gain to the elevator command, so that the signal path including KGEP effectively serves as the pitch rate command signal.

16. In a method of flight control in which a thrust command is computed based on a total aircraft energy rate error E_Sε

and an elevator command is computed based on an energy distribution rate error D.sub.ε

, the improvement comprising computing an elevator control command in response to a control column input by the pilot;

the computing including;

(a) computing the energy distribution rate error as a function of an acceleration error, a current flight path angle γ

, and a manual commanded flight path angle γ

_C.sbsb.MAN ;

the γ

_C.sbsb.MAN signal initially being in excess of the sustainable flight path angle established for the prevailing thrust condition by an amount proportional to the column input; and

(b) processing the γ

_C.sbsb.MAN signal so that the amount of γ

_C.sbsb.MAN in excess of the sustainable flight path angle for the prevailing thrust condition and an altered speed condition is eventually eliminated, with the aircraft establishing a change in speed relative to a set reference speed, the speed change being proportional to the column input.

17. The improvement according to claim 16, wherein the set reference speed can be set and displayed via a cockpit mode control panel.

18. The improvement according to claim 16, wherein the pilot is able to change the set reference speed of the airplane via a pitch trim switch on the control column.