GB1579339A - Aircraft control systems - Google Patents

Description

(54) IMPROVEMENTS IN OR RELATING TO AIRCRAFT CONTROL SYSTEMS (71) I, SECRETARY OF STATE FOR DEFENCE, LONDON, do hereby declare the invention, for which I pray that a patent may be granted to me, and the method by which it is to be performed, to be particularly described in and by the following statement: The present invention relates to aircraft fly-by-wire control systems.

In an aircraft fly-by-wire control system signals from the pilot's controls are passed electrically via a computer to actuators which operate the aerodynamic control surfaces. For convenience, the term pilot's controls will, in this specification, mean only those controls (conventionally control column and rudder) used to move the aerodynamic control surfaces.

FlyWby-wire control systems are replacing conventional control systems in many modern aircraft.

The use of a computer in fly-by-wire systems, due to the computer's speed of response, allows the inherent aircraft stability, which is due to its geometry, to be extensively modified.

Indeed (although this has no bearing on the present invention) the use of fly-by-wire systems allows aircraft to be flown whose inherent stability is such that using conventional control systems they would be uncontrollable. The response characteristics (that is the way in which an aircraft responds to movements of the pilot's controls) of a fly-by-wire controlled aircraft are determined by programming suitable control laws (governing the relationship between movement of the pilot's controls and movement of the aerodynamic control surfaces) into the computer. However, whilst this use of a computer allows the response characteristics.

of a fly-by-wire aircraft to be modified much more easily than the response characteristics of a conventionally controlled aircraft, where modifications can only be made by changes in aircraft geometry, it does not solve all the designer's problems regarding response characteristics. The problem is that optimisation of the response characteristics for one flight condition can sometimes only be achieved at the expense of the response characteristics at other flight conditions. For example, in a fighter aircraft a compromise may have to be made between obtaining adequate normal acceleration transient response when manoeuvring to lock on to a target and acceptable attitude response when tracking the target.In recent experiments with a fly-by-wire controlled aircraft having acceptable normal acceleration respdnse for manoeuvring it was found that when tracking a target in turbulence there was a tendency for the pilot to over control in pitch, causing a pilot induced oscillation' (PIO). This degraded aiming accuracy and increased pilot workload.

It has been suggested that the computer should have programmed within it a family of control laws, each law being optimised to a particular piloting task, there being a manual mode selector to allow the pilot to select the appropriate control law for each task during flight. However piloting tasks overlap one another and manual mode selections of this type is frequently impracticable.

According to the present invention an aircraft fiy-by-wire control system having a computer which receives input signals which are a function of a pilot's control demands, includes programmed into the computer a set of criteria with which the direction and amplitude of the pilot's control demands are compared, and means for modifying the aircraft control laws according to the comparison to optimise the aircraft control laws for the existing flight conditions.

The invention is most advantageously used in conjunction with the pilot's pitching mode control, with the criteria selected to determine whether the requirement is for rapid alteration of normal acceleration level or for maintenance of a current attitude or flight path angle.

One means of modifying the aircraft control laws involves the use of one or more phase advance filters, one or more phase lag filters, or both, the or each filter being switched into or out of a relevant control circuit according to the comparison.

The problem, and the benefits of the present invention, may be better understood from the following description of one embodiment of the invention, which is described by way of example only, with reference to the accompanying drawings accompanying the Provisional Specification of which: Figures 1A and 1B show two pitch rate response characteristics, Figure 2 shows a flyby-wire control system according to the invention, Figures 3A to 3D show a typical flyby-wire response characteristic, and the way in which this can be modified by the use of filters, and Figures 4A to 4D show the response characteristics which can be obtained by use of the invention.

Figure 1 shows the effect on an aircraft of movement of the pitch control lever followed by release of the lever, allowing the lever to return to neutral, for two different aircraft response characteristics. In Figure 1A the response characteristics are designed for manoeuvrability. With this type of response when the control lever is released there is an overshoot in the pitch rate, which results in the aircraft's final pitch attitude differing from the planned attitude by an amount which depends on the pitch rate overshoot. In attempting to regain and maintain the required pitch attitude it is easy for a pilot, especially in turbulence, to set up an oscillatory motion.

In Figure 1B with response characteristics designed for stability, the pitch rate returns to zero without overshoot when the control lever is released. This results in the pitch attitude overshooting the desired attitude, but the pilot is able to regain and maintain the desired attitude with little danger of inducing oscillation. This is achieved, however, only at the expense of a loss of manoeuvrability, as can be seen by comparing the normal accelera tion graph of Figure 1B with that of Figure 1zA.

In an aircraft fly-wire control system accord ing to the invention (Figure 2) signals repres entative of movement of a pitch control lever 10 are passed to a computer 11. The com puter 11 processes the signals and compares them with predetermined criteria. Depending on the comparison the computer then selects either a phase advance filter 12 or a phase lag filter 13, through which signals pass to an actuator 14 which moves the pitch control surface 15. In general the signals will be mixed with inputs from other sensors such as, for example, pitch rate and normal acceleration sensors which will themselves be filtered in a mixing filter 16 in such a way as to ensure adequate stability in turbulence. Figure 3 shows the effects of filters on the response characteristics of a known fly-by-wire controlled aircraft.The response without filters is shown in Figure 3A. Computed responses with a phase advance filter (1 + 0.5s) (1 + 0.2s) (henceforth referred to as an acceleration filter) and with a phase lag filter (1 + 0.2s) (;1 + 0.5s) (henceforth referred to as an attitude filter) are shown in Figures 3B and 3C respectively.

Figure 3B shows the computed response when the phase advance filter is in circuit while the control lever 10 is positioned away from the neutral position and the phase lag filter is switched in upon release of the control lever 10. The response characteristics are thus such as will allow quick attainment of a desired pitch attitude followed by ease of maintenance of that attitude.

In practice the computer must have programmed into it various criteria against which control lever 10 movements can be compared, to allow the most advantageous selection of filter to be made. Examples of criteria are: a. large amplitude step type commands, where the pilot desired to rapidly attain a steady normal acceleration. This requires the acceleration filter in the circuit; b. control lever 10 release to zero, when the pilot desires to maintain the current attitude or flight path angle, and would be distracted by attitude overshoot. This requires the attitude filter in circuit.

c. control lever 10 reversal through zero, where the pilot is seeking normal acceleration reversal. This requires the acceleration filter in circuit.

d. small control lever 10 movement about zero, where the pilot is primarily concerned with making small flight path angle changes.

The attitude filter is required in circuit; e. small control lever 10 movements about a mean offset position, where the pilot is seeking small flight path angle changes superimposed on an established normal acceleration.

The attitude filter is required in circuit.

Criteria (a), (b) and (c) can be mechanised from the sign and sense of the control lever 10 movement. For (d) and (e) threshold levels must be defined within which conditions compatible with criteria (a) for reselecting the acceleration filter are ignored.

The effect of using the above criteria, and filters as described earlier, for aircraft whose response characteristics are illustrated in Figure 3A are illustrated in Figure 4. A threshold level of 10% of control lever 10 movement was selected above which the acceleration filter was selected. For criteria (d) the decision whether to switch was taken by comparing the current control lever 10 demand with the average demand over the preceding two seconds. This average value was updated every 0.1 s. For each criteria the switching decision was taken every 0.01 s. Small movements of the control lever 10 following an initial large displacement causes the attitude filter to be selected, but not before a delay of approximately 14 s during which the acceleration filter is retained to allow steady state normal acceleration to be attained.In Figure 4A normal acceleration is commanded and then small stick movements are made during the manoeuvre. Initially the acceleration filter is switched in, followed by the attitude filter. In Figure 4B the pilot is shown as demanding a level of normal acceleration which he then increases. The acceleration filter remains in operation during this procedure; selections of the attitude filter following the initial demand are inhibited by the 14 second time delay, so ensuring that the commanded acceleration response is attained. The attitude filter is switched in on release of the control lever 10.

The case where the pilot alternately selects positive and negative normal acceleration is illustrated in Figure 4C. During these manoeuvres the acceleration filter is retained, apart from momentary engagements of the attitude filter on each stick release. Illustrated in Figure 4D is a demand for a normal acceleration followed by a return to zero acceleration and small control lever 10 movements near to its neutral position. Initially the acceleration filter is selected, followed by the attitude filter.

It will be realised that the criteria mentioned above were selected with a view to a particular aircraft, and that for a different aircraft or different piloting tasks different criteria may be preferable. The invention has been demonstrated with two filters selected according to pre-defined criteria. The concept can be extended to a range of three or more filters, some of which may appear within the basic stabilisation loop of the aircraft (thereby altering, for example, the turbulence attenuation properties of the stabilisation loop). Further, the selection process may involve a gradual transition between filters rather than the discrete changes illustrated above. The number of filters, and the filter characteristics, will depend on the type of aircraft to which the invention is being applied and on the tasks which it is required to undertake.

Whilst the invention has been described as being used with the pitch control circuit only, there may be cases where it can advantageously be used in the roll control circuit, the yaw control circuit, or both.

The invention may be implemented using digital, analogue, or hybrid analogue-digital technology.

WHAT I CLAIM IS: 1. An aircraft fly-by-wire control system having a computer which receives input signals which are a function of a pilots control demands, including programmed into the computer a set of criteria with which the direction and amplitude of the pilots control demands are compared, and means for modifying the aircraft control laws according to the com parisons to optimise the aircraft response characteristics for the existing flight conditions.

2. A control system as claimed in Claim 1 wherein the means for modifying the aircraft control laws includes one or more phase advance filters, one or more phase lag filters, or both, the or each filter being switched into or out of a relevant control circuit according to the comparison.

3. A control system as claimed in Claim 1 or in Claim 2 and wherein the aircraft response characteristics are modifiable in the pitching plane.

4. A control system as claimed in Claim 3 wherein the criteria are chosen to indicate whether the requirement is for rapid alteration of normal acceleration level or for maintenance of a current attitude or flight path angle.

5. A control system as claimed in any one of Claims 1 to 4 wherein the aircraft response characteristics are modifiable in the rolling plane.

6. A control system as claimed in any one of Claims 1 to 5 wherein the aircraft response characteristics are modifiable in the yawing plane.

7. An aircraft control system substantially as herein described with reference to the drawings accompanying the provisional specification.

8. An aircraft having a control system as claimed in any one of Claims 1 to 7.

**WARNING** end of DESC field may overlap start of CLMS **.

Claims (8)

**WARNING** start of CLMS field may overlap end of DESC **. whether to switch was taken by comparing the current control lever 10 demand with the average demand over the preceding two seconds. This average value was updated every 0.1 s. For each criteria the switching decision was taken every 0.01 s. Small movements of the control lever 10 following an initial large displacement causes the attitude filter to be selected, but not before a delay of approximately 14 s during which the acceleration filter is retained to allow steady state normal acceleration to be attained. In Figure 4A normal acceleration is commanded and then small stick movements are made during the manoeuvre. Initially the acceleration filter is switched in, followed by the attitude filter.In Figure 4B the pilot is shown as demanding a level of normal acceleration which he then increases. The acceleration filter remains in operation during this procedure; selections of the attitude filter following the initial demand are inhibited by the 14 second time delay, so ensuring that the commanded acceleration response is attained. The attitude filter is switched in on release of the control lever 10. The case where the pilot alternately selects positive and negative normal acceleration is illustrated in Figure 4C. During these manoeuvres the acceleration filter is retained, apart from momentary engagements of the attitude filter on each stick release. Illustrated in Figure 4D is a demand for a normal acceleration followed by a return to zero acceleration and small control lever 10 movements near to its neutral position. Initially the acceleration filter is selected, followed by the attitude filter. It will be realised that the criteria mentioned above were selected with a view to a particular aircraft, and that for a different aircraft or different piloting tasks different criteria may be preferable. The invention has been demonstrated with two filters selected according to pre-defined criteria. The concept can be extended to a range of three or more filters, some of which may appear within the basic stabilisation loop of the aircraft (thereby altering, for example, the turbulence attenuation properties of the stabilisation loop). Further, the selection process may involve a gradual transition between filters rather than the discrete changes illustrated above. The number of filters, and the filter characteristics, will depend on the type of aircraft to which the invention is being applied and on the tasks which it is required to undertake. Whilst the invention has been described as being used with the pitch control circuit only, there may be cases where it can advantageously be used in the roll control circuit, the yaw control circuit, or both. The invention may be implemented using digital, analogue, or hybrid analogue-digital technology. WHAT I CLAIM IS:

1. An aircraft fly-by-wire control system having a computer which receives input signals which are a function of a pilots control demands, including programmed into the computer a set of criteria with which the direction and amplitude of the pilots control demands are compared, and means for modifying the aircraft control laws according to the com parisons to optimise the aircraft response characteristics for the existing flight conditions.

2. A control system as claimed in Claim 1 wherein the means for modifying the aircraft control laws includes one or more phase advance filters, one or more phase lag filters, or both, the or each filter being switched into or out of a relevant control circuit according to the comparison.

3. A control system as claimed in Claim 1 or in Claim 2 and wherein the aircraft response characteristics are modifiable in the pitching plane.

4. A control system as claimed in Claim 3 wherein the criteria are chosen to indicate whether the requirement is for rapid alteration of normal acceleration level or for maintenance of a current attitude or flight path angle.

5. A control system as claimed in any one of Claims 1 to 4 wherein the aircraft response characteristics are modifiable in the rolling plane.

6. A control system as claimed in any one of Claims 1 to 5 wherein the aircraft response characteristics are modifiable in the yawing plane.

7. An aircraft control system substantially as herein described with reference to the drawings accompanying the provisional specification.

8. An aircraft having a control system as claimed in any one of Claims 1 to 7.