The Concorde Artificial Feel System

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Concorde is fitted with an artificial feel system. The purpose of this system is to make the aircraft feel to the pilot like a conventionally controlled aircraft. The system increases the Pilot’s control stiffness through jack springs, as a function of the speed of the aircraft. On a traditionally controlled airliner the controls would be harder to operate when the deflection of the control surfaces is increased at high speeds due to the air resistance.

Concorde’s artificial feel system re-instates this resistance into the controls to make the aircraft handle as a traditional aircraft would. Without the artificial system the flight controls would just move to where they were sent, with very little effort by the pilot, making it very easy to over-control the aircraft.

DESCRIPTION AND OPERATION

 

As the aircraft servo-controls are irreversible, no feed back is transmitted to the pilot regardless of the load on the control surfaces. The feel which is similar to that of a conventional feel, is therefore artificially reproduced by a system interposed in each of the three control axis.

This system comprises essentially of a mechanical differential, a spring rod and two electro-hydraulic jacks.

The Concorde Artificial Feel system has three purposes:

  • To restore loads to the flight controls compatible with precise and stable piloting.
  • To give considerable feel loads to the pilot before reacting flight configurations that are dangerous to the structure.
  • To cut-in on the auto-pilot load limitation.

 

Principle of Operation

 

For the yaw axis, the object is to apply to rudder pedals a load proportional to a function of calibrated airspeed and to the defection limit imposed by structure.

The variable resistance is ensured by two hydraulic jacks, each controlled by a computer. Both jacks are controlled permanently, but only one is operative, the second one remaining at a stop.

The first jack, which acts in priority, is supplied with pressure from the Blue hydraulic system. The control computer is activated when the YAW switch on the ARTIFICIAL FEEL engage switch unit No.1 (on overhead panel in the flight deck compartment) is engaged. “The second jack is supplied with Green hydraulic system pressure and the jack control computer is activated when the YAW switch on the ARTIFICIAL FEEL engage switch unit No.2 (on overhead panel in the flight deck compartment) is engaged.

When both the YAW switches are engaged, the Green jack automatically replaces the Blue jack in the case of a failure of the latter. The stand-by hydraulic pressure (Yellow) is not used.

A spring rod provides the resistance corresponding to tow speed conditions and ensures a minimum safety in case of a double electro-hydraulic failure.

Artificial Feel Mechanical System

(Ref. Fig. 001 )

The Artificial Feet components for the three Flight Control axes are mounted on a single chassis attached to the aircraft structure in the forward compartment (in the area identified as zone 121).

PICTURE 1 (27 22 00)here!

The system comprises of the following:

  • An Integral trim assembly, this enables the loads of the artificial feel system to be cancelled by a mechanical differential.
  • A double action-load limiting spring rod which restores loads proportional to out of trim deflections.
  • A rocker arm which transmits loads which are delivered by the electro-hydraulic jacks.
  • Two electro-hydraulic jacks, these are controlled by computers, ensuring variable resistance.

 

Artificial Feel Electronic System

(Ref. Fig. 002 )HERE!!

There are two identical Artificial Feel systems provided on the aircraft. These two systems operate simultaneously.

The two systems comprise of the following:

  • Two Artificial Feel computers: Computer No.1 (IC235) is located in left-hand electronics rack, on shelf 6-215. Computer No.2 (2C235) in right-hand electronics rack, on shelf -6-216. They develop load laws and monitor the system.
  • Two Artificial Feel engage switch units (IC236 and
  • 2C236) which are located on overhead panel. These allow activation of the system and transmit orders generated by the computer.
  • Two press-to-test push buttons (IC245 and 2C245) located on panel 29-214 at Flight Engineer’s station.  They initiate the computer monitoring channel tests.
  • A Yaw deflection sensor (C246), located in zone 121 and driven by a link actuated by the Yaw torque tube. This sensor comprises of two potentiometers providing the two artificial feel computers with yaw deflection data.

In addition, the two systems are connected to the following components:

  • Air Data Computer, for data concerning Calibrate Air Speed and General ADC failure.
  • Master Warning System which receives the GONG and FEEL warning Light activation order.

 

Assembly Integral Trim

(Ref. Fig. 001 ) HERE 

Picture 2 here!

The assembly comprises of the following:

  • An input lever, which is actuated by the mechanical control from the pilot, the end of this lever is machined as a toothed sector.
  • A worm screw in which the input lever toothed sector is enmeshed. This worm screw which is integral with the assembly flanges is fitted with a double pinion ensuring its drive by the trim reduction gear.
  • The drive is affected by a universal joint in which the rotation axis is the same as that of the input lever.
  • The flanges of the assembly form a right-angle in which the apexes are denoted by the points O, R and M in the illustrations.

This assembly ensures two functions:

  • Cancellation of artificial feel system loads through a differential mechanism.
  • Control of rudders through an irreversible mechanical control. (Flight using trim) independent of the main Flight Control

These two functions are ensured by the toothed sector worm screw mechanism described above.

Spring Rod

 (Ref. Fig.001 and 003)

The spring rod is made up of:

  • A body (stationary) anchored to the chassis
  • A mobile rod attached to point R on the integral trim assembly flanges.
  • Springs providing the load limit for returning the rudder pedals to neutral position and the variable resistance corresponding to low speed conditions.
  • Two sliders ensuring junction between the springs and the mobile rod. Whatever the direction of movement, the reaction of the spring rod tends to align points O, M and A.

 

Jack – Artificial Feel

 (Ref. Fiq. 004 )

One section of the artificial feel jack is hydraulic and the other section is electrical. Hydraulic supply to the cylinder is via a spool valve controlled by an electro-valve.

PICTURE 3 HERE!

Pressure control in the jack is regulated by a servo valve, the electro-valve is controlled by the electrical monitoring channel and the servo-valve is controlled by the electrical control mode.

The piston rod is fitted with a load detector which compares the actual load on the jack with the control order. If the electro-valve does not receive a signal, there is no hydraulic pressure on the servo-valve and the jack chambers are connected to tank return.

If the electro-valve is energized, hydraulic pressure is admitted to the servo-valve.

The servo-valve regulates the pressure from signals received from the control channel. Admitted to the front chamber of the jack, this pressure maintains a load corresponding to the control order. The rear chamber is connected to tank return.

Computers – Artificial Feel

(Ref. Fig.’ 005 ) .

Each computer is located inside a housing, and this housing comprises of the following:

  • On the front panel, a P23 connector for test and maintenance purposes, an hour meter and a handling grip.
  •  On the rear panel a double connector (DP X2) is provided for connection to the aircraft electrical network. 
  • A computer controls the hydraulic pressure of a jack.

For each jack the electronic assembly consists of the following:

  • One control channel
  • One monitoring channel
  • The supply of these channel load detectors
  • The supply of these channels
  • The circuit necessary for testing of monitoring systems.

The control channel achieves the following:

  • The development of the control electrical order from the various control signals with which the channel is provided.
  • The comparison of this order with the return signals from the Load detector.
  • The development of the servo-valve control signal.

PICTURE FIVE HERE!

The monitoring channel supervises the control channel.

  • It receives identical (and independent) control signals and develops a second electrical order.
  • It compares this second order to the return signal from the load detector second circuit.
  • It switches off the electro-valve electrical supply if the error signal resulting from the comparison exceeds a fixed threshold.

Computer No.1 receives information from Air Data Computer No.1 and controls the Blue Jacks – Computer No.2 receives information from the Air Data Computer No.2 and controls the Green Jacks.

Hydro-mechanical Operation

PICTURE 6 HERE!

Actuated by the flight control rod, the input lever acts through the integral trim assembly:

On the spring rod about point R, on the rocker arm about point M. The ends of the rocker arm are connected to two hydraulic jacks at B and C. The upper (blue) jack is supplied by the blue hydraulic system and the lower (green) jack is supplied by the green hydraulic system.

When both jacks are supplied they apply two equal forces at points B and C. The lever AB of the rocker arm, being longer than AC gives priority to the blue jack. The green jack abuts the stop at point E.

Control actuation compresses the spring rod and displaces the rocker arm against the action of the hydraulic jack. To overcome spring rod resistance and jack action, loads proportional to control displacement and resistance in the system must be applied to the control.

Blue jack failure

‘(Ref. Fig. 007 ) .

If a fault occurs and the monitoring channel closes the electro-valve:

  • There is no pressure at the servo-valve.
  • Both chambers of the jack are connected to the tank.
  • The rocker arm actuated by the green jack tilts.
  • The blue jack abuts the stop.
  • The system operates on the green jack.

If both jacks fail only the spring rod remains active, and a speed reduction is then necessary.

PICTURE 7 HERE!

Electronic Operation

 

Achieved Load Law

(Ref. Fig. 008 )

The load exerted by the jack is a linear Law:

  • Proportional to a function of calibrated air speed developed by the ADC.
  • Reduced by a constant value.

The Load exerted by the jack can be expressed as follows:

Fj =kv-h

Where:

Fj is the load exerted by the jack.

V is the command voltage, function of calibrated air speed developed by the ADC.

k and h are constants.

A sudden increase of the load is achieved in order to prevent rudder deflections which could be dangerous to the structure at certain speeds.

When the deflection (detected at the Yaw torque tube by a potentiometer) becomes greater than the deflection limit which is a second function of calibrated air speed (developed by the ADC), the load law is as follows:

  • Multiplied by a 1.1 factor
  • Increased by a constant value

The Load exerted by the jack can be then expressed as follows:

Fj = 1.1 kv- h + c

Where:

c is a constant.

NOTE:

This limiting function provides intentionally produced hysteresis to avoid repeated switching in the region of the deflection limit.

Control Channel

(Ref. Fig. 009 )

This channel is made up of two cards:

  • A functional amplifier
  • A functional output amplifier

(1) In the functional amplifier, the return signal from the load detector is compared to the sum of the signals representing the load law.

NOTE:

In the case of limit of deflection, this load law is modified by a resistor network controlled by an analog gate which itself receives the Load multiplication Logic order from a circuit on the monitoring channel comparison amplifier card.

After demodulation, the resulting signal forms the control error signal (c). A bias is superimposed on the error signal (difference between order to be carried out and measured force) and the sum of these two signals forms the functional channel control signal.

A demodulator filters parasitic components in the control signal by converting the alternating w components of the latter to a DC signal. The parasitic components are then eliminated by the corrective network.

(2) The functional output amplifier comprises of the following:

  • A phase lead corrective network
  • A current amplifier
  • A circuit limiting the servo valve current

At the channel output, the servo valve control current I is divided in two parts:

I = Io + AI

 Where:

Io= constant current resulting from the bias signal amplification.

AI = variable current, positive or negative, resulting from the error signal amplification.

This method permits the control of the servo-valve the zero of which is offset (zero hydraulic flow for a Io control current) in order to connect the jack to the tank in the event of accidental suppression of the servo-valve current.

To reduce the inertia of the jack servo-valve flapper a 400 Hz signal is superimposed on the error signal at the “Dither” input of the output amplifier.

Monitoring Channel

 

This channel is made up of three cards:

  • A monitoring amplifier
  • A comparator demodulator
  • A monitoring output amplifier

(1) Monitoring amplifier

  • The monitoring channel is provided with the following inputs: Force order and feedback inputs identical with those of the functional channel.
  • A TEST input which causes the triggering of the monitoring channel.

In addition, this card comprises a circuit which compares the rudder pedal deflection with a function of calibrated air speed. When the deflection becomes greater by 400 mV than the speed function, a logic command is generated which controls’ the Load multiplication for the two controls and monitoring channels.

NOTE:

There is no bias input on the monitoring amplifier.

(2) Comparator demodulator

The AC variation (E) voltage, developed by the monitoring amplifier, is summed separately with two AC voltages proportional to the desired triggering thresholds (k threshold). The result of each summing is demodulated polarity of voltage from either channel of the demodulator is negative if variation (&) is greater than the triggering threshold during a period of time greater than the monitoring timing.

(3) Monitoring output amplifier

If a variation greater than the triggering threshold has been detected, or if a failure of the associated ADC occurs, the supply of the electro-valve and the engage switch holding coil is cut off.

  • To avoid disconnection in the event of a temporary 115 VAC power supply loss.
  • The comparator is inhibited if the 115 VAC loss occurs on the three axes (pitch, yaw roll) of the computer considered.
  • The magnetically held engage switch is supplied with a stand by voltage from the 115 volt monitoring card. Inhibit duration is 1.25 second. D. 115 Volt Monitoring Card

This card is common to the three axes (roll, yaw, pitch). It comprises of the following:

  • A regulated DC power supply, which, from the 28 volt validity of the ADC, enables the monitoring output amplifiers and the magnetically held engage switch to be supplied in the event of loss of 115 VAC.
  • A circuit which generates a signal inhibiting the monitoring comparators in the event of loss of the 15 volt supplies on the three axes. Inhibit duration (1.25 second) is controlled by a mono-stable circuit.

In addition, this card comprises a circuit designed to supply the trim deflection sensor on functional side

Description and Operation

 

General Power Supply

A transformer supplies the following outputs from the 115 V 400 Hz power supply.

  • Supply of the load detector functional measuring circuit (10 V RMS).
  • Supply of the load detector monitoring measuring circuit (10 V RMS).
  • Supply of the various potentiometers generating force orders (symmetrical windings 2 x 26 volts RMS: Ref + and Ref -)
  • Supply of electrical channels: two symmetrical windings supply a diode bridge followed by two LC filters. These circuits deliver the two + 16 V and – 16 V voltages which are the stabilized power supplies.

Controls and Indicating

(Ref. Fig. 009 )

Controls and Indicating of the two Artificial Feel Systems

Each artificial feel system is activated by the engagement of the magnetically held YAW switch, integral with each artificial feel engage switch unit No.1 and No.2, which are located on the overhead panel. This switch remains engaged if the monitoring channel does not detect a fault of the control channel.

When detecting a fault:

  • The supply to the switch engagement holding coil is cut-out.
  • The supply to the electro-valve is cut-out.
  • The supply to the servo-valve is cut-out.
  • The engagement switch disengages and indicates OFF.
  • The gong sounds and the FEEL warning light illuminates on the master warning panel. However, these two warnings are only activated when, both systems 1 and 2 being engaged, both monitoring channel’s detect a fault.

Test

At Flight Engineer’s station, the two ARTIFICIAL FEEL push buttons, TEST 1 and TEST 2 enable the indicating system of system No.1 and No.2 to be checked.

When the engagement switches of a system are engaged, action on the corresponding test push button causes the switch to disengage.

Electrical Supply

SERVICE BUSBAR C/B PANEL

Computer No.1 (IC236) No.2 ESSENTIAL 115 VAC 6X 2-2213

YAW stage power supply.

Computer No.2 (2C236) B AVIONICS 115 VAC 11X 13-216

YAW stage power supply

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