GB1604285A

GB1604285A - Cabin pressure control systems for aircraft

Info

Publication number: GB1604285A
Application number: GB2289578A
Authority: GB
Original assignee: Garrett Corp
Current assignee: Garrett Corp
Priority date: 1977-06-13
Filing date: 1978-05-26
Publication date: 1981-12-09
Also published as: JPS5417298A; FR2394447B1; FR2394447A1

Description

(54) CABIN PRESSURE CONTROL SYSTEMS FOR AIRCRAFT (71) We, THE GARRETT CORPORATION, a Corporation organised under the laws of the State of California, United States of America, of 9851-9951 Sepulveda Boulevard, P.O. Box 92248, Los Angeles, California 90009, United States of America, do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to cabin pressure control systems for aircraft.

According to one aspect of the present invention, a cabin pressure control system for an aircraft comprises two controllers, each of which is capable of controlling the cabin pressure, independently of the other of the two controllers, and each of which has an enabled state and a disabled state, and is arranged, when enabled, to control the cabin pressure, and the system also comprising means arranged to enable and disable the controllers in accordance with predetermined conditions, in such a way that under normal operating conditions only one controller is enabled.

The use of two controllers means that, if the enabled controller should malfunction, control may be transferred to the other controller. Thus, in a preferred embodiment, the system may also include means arranged to monitor the performance of whichever controller is for the time being enabled, hereinafter referred to as the primary controller), and on detection of a malfunction of that controller, to activate transferring means to enable whichever controller was previously disabled (hereinafter referred to as the standby controller), and to disable the primary controller. Because the transferring means will effect the transfer of control between the controllers automatically on detection of a malfunction, the need for a member of the crew of the aircraft to monitor the performance of the pressure control system is largely eliminated, reducing the crew workload.

Preferably, the transferring means is also arranged to change the state of the two controllers for each successive flight of the aircraft. This transfer may be blocked if a malfunction of the primary controller has been detected during the preceding flight.

The system may also be arranged to ensure that, if both controllers should for any reason come into the same state (both enabled or both disabled), one of the controllers will be changed to the other state, to avoid the possibility of both controllers trying simultaneously to control cabin pressure, or of the cabin pressure becoming uncontrolled.

The invention also extends, according to a second aspect, to an aircraft including a cabin pressure control system according to the first aspect.

The invention also provides, according to a third aspect, a method for regulating cabin air pressure in an aircraft containing a pair of cabin pressure controllers, each of which is capable of controlling the cabin air pressure independently of the other of the controllers, the said method comprising the steps of: enabling one of the said pair of cabin pressure controllers; and, in response to predetermined conditions, enabling the other of the said pair of cabin pressure controllers to regulate the cabin air pressure, while disabling the said one of the pair of controllers.

The invention may be carried into practice in various ways, but one specific embodiment will be described below by way of example. This embodiment is also described and claimed in our co-pending applications Nos. 22897/78 (Serial No. 1604287) and 22896/78 (Serial No.

1604286). Any of the features claimed in those applications may be used in conjunction with the features of the present invention. In the drawings: Figure I is a block diagram of a dual automatic cabin pressure control system of this invention; Figure 2 is a block diagram of one of the dual automatic controllers of Figure 1; Figure 3 is a block diagram of AP limit logic for use in the controller of Figure 2; Figure 4 is a graph of rate increment vs. differential pressure for a maximum AP amplifier of Figure 3; Figure 5 is a schematic diagram of a cabin altitude function generator for the controller of Figure 2; Figure 6 is a series of graphs labeled 6A-6C showing voltage/time relationships for the cabin altitude function generator of Figure 5; Figure 7A is a graph of cabin pressure vs. ambient pressure for the controller of Figure 2;; Figure 7B is a graph of cabin pressure vs. ambient pressure for a prior art controller; Figure 8 is a block diagram of automatic transfer circuit of Figure 1; Figure 9 is a logic diagram of a malfunction detection logic of Figure 8; Figure 10 is a logic diagram of a successive flight transfer of Figure 8; Figure 11 is a logic diagram of an on-off control interconnect showing the interconnect logic for the system of Figure 8; and Figure 12 is a plan view of a cabin pressure selector panel suitable for the system of the present invention.

Description of the Preferred Embodiment Figure 1 illustrates a dual automatic cabin pressure control system embodying the present invention. A selector panel 10 (see also Figure 12) is manually controlled to produce output voltages corresponding to the landing aTtitude, set by a selector knob 4, landing field barometric pressure correction, set by a selector knob 5, and a selected limit of the rate of cabin pressure change, set by a selector knob 6. These voltages are transmitted through conductors 12 and 14 to identical automatic cabin pressure controllers 16 and 17, respectively.

The cabin pressure selector 10 may include dual sets of selector potentiometers driven by the selector knobs, 4, 5, and 6, one for cabin pressure controller 16 and one for cabin pressure controller 17. These dual sets of potentiometers would preferably be ganged together so that the command voltages for both controllers are the same. Controllers 16 and 17 are connected through conductors 18 and 19, respectively, to an actuator 20 which controls opening and closing of an outflow valve (not shown).

Each of the controllers 16 and 17 preferably operates a seParate outflow valve motor in the actuator 20 with only the mechanical valve linkage and reduction gears in common.

Each controller has its own power supply, sensor units, electronic logic and control circuits (illustrated in Figure 2) so that they operate completely independently of each other, minimizing the possibility that both controllers could be disabled at the same time.

Controller 16 receives an input voltage proportional to ambient atmospheric pressure from an atmosphere pressure sensor 21 through conductor 22 and an input voltage proportional to cabin pressure from a cabin pressure sensor 23 through conductor 24.

Controller 17 receives corresponding input voltages from an atmosphere pressure sensor 26 through conductor 28 and from cabin pressure sensor 29 through conductor 30. Pressure sensors 21, 23, 26 and 29 may be any suitable pressure transducers which produce a detectable change in output responsive to changes in pressure. An automatic transfer circuit 32 receives inputs from automatic cabin pressure controller 17 through conductor 33, from automatic transfer circuit 34 through conductor 36 and from selector panel 10 through conductor 37. Automatic transfer circuit 32 is also connected to automatic cabin pressure controller 16 through operation of a relay (illustrated schematically as conductor 38) to actuator 20 through conductor 39 and to automatic transfer circuit 34 through conductor 40.

Automatic transfer circuit 34 receives an input from actuator 20 through conductor 41, from automatic cabin pressure controller 16 through conductor 42, and from selector panel 10 through conductor 43 and sends an input to automatic cabin pressure controller 17 through operation of a relay (illustrated schematically as conductor 44).

Each automatic transfer circuit is associated with and controls one of the automatic cabin pressure controllers. Automatic transfer circuit 32 is associated with automatic cabin pressure controller 16 and automatic transfer circuit 34 is associated with automatic cabin pressure controller 17. Automatic cabin pressure controllers 16 and 17 and automatic transfer circuits 32 and 34 also receive inputs from a landing gear switch 46 through conductors 48, 50, 52 and 54, respectively, and from a throttle switch 56 through conductors 58, 60, 62 and 64, respectively, the purposes for which will be subsequently described in detail.

During each flight, on an alternating basis, one controller functions as the primary controller and the other serves as the standby controller. The primary controller is in actual control of the cabin pressure throughout the flight. In the course of the flight, the standby controller monitors the performance of the primary controller. If the primary controller should malfunction, the standby controller will take over control of cabin pressure. If no malfunction occurs, identity of the primary controller is transferred upon landing for the next flight. This automatic transfer of the primary mode between controllers on successive flights enables each controller of the dual automatic control system to automatically check the other on a regular basis to provide maximum reliability for the entire system.Each of the cabin pressure controllers utilize the input voltages from its corresponding atmosphere pressure sensor to compute a desired corresponding cabin pressure. This is compared to the voltage corresponding to cabin pressure from the cabin pressure sensor to generate an error signal utilized by the primary controller to control the position of the actuator 20.

Referring now to Figure 2, the automatic cabin pressure controller 16 and associated components are illustrated in detail, it being understood that controller 17 (Figure 1) has an identical set of components similarly connected.

The conductor 12 connecting the selector panel 10 to the controller 16 comprises a conductor 12a which connects the potentiometer of the "rate limit" selector 6 of the selector panel (Figure 12) to a rate logic circuit 66 of the controller. A conductor 12b connects selectors 4 and 5 for landing field altitude and barometric pressure of the selector panel 10 to a high altitude discriminator circuit 68. A conductor 12c connects the selector panel 10 to a power supply 70 in the controller 16 which provides operating voltage for the selector panel 10. The power supply 70 is also connected by a conductor 42a to a power loss detector circuit which will be subsequently described in connection with Figure 9.

The rate logic circuit 66 receives additional input from a cabin altitude function generator 72 through a conductor 74, from a descent detector circuit 76 through a conductor 78, and from the cabin pressure sensor 23 through a conductor 24a. The cabin pressure sensor 23 also provides its voltage signal through a conductor 24b to a AP limit logic circuit 80. This signal is also conveyed by a conductor 24c to a rate amplifier 90 and by a conductor 42b to a malfunction detection logic circuit which will be described in connection with Figure 9, and to dynamic compensator 100 by conductor 24d.

Connection of the atmosphere pressure sensor 21 to the controller 16 is by a conductor 22a to the cabin altitude function generator 72, via a conductor 22b to the descent detector 76, and by a conductor 22c to the AP limit logic circuit 80. The cabin altitude function generator 72 is also connected by a conductor 84 to the descent detector 76 and by a conductor 86 to the high altitude discriminator 68 which is in turn connected by a conductor 88 to the rate amplifier circuit 90. The AP limit logic circuit 80 is also connected to the rate logic circuit 66 by a conductor 92 and connected by a conductor 42c to the malfunction detection logic circuit of Figure 9.

A ground logic circuit 94 in the controller 16 receives a DC input signal from a source of direct current voltage 96 by parallel paths through the landing gear switch 46 via conductor 48 and through the throttle switch 56 via conductor 58. Output from the ground logic circuit 94 is provided to the descent detector circuit 76 through conductor 98 then to a dynamic compensator circuit 100 through a conductor 102, and to driver 110 through conductor 103.

The output produced by the rate logic circuit 66 is transmitted through conductor 104 to the rate amplifier 90 which produces a corresponding output which is transmitted to the dynamic compensator 100 via conductor 106. The resulting controller output is transmitted from the dynamic compensator 100 through the conductor 108 to a driver circuit 110 of the actuator 20. The driver circuit 110 is connected in a feedback loop comprising conductor 112, contact 114c of switch 114 (see Figure 8), motor 116, conductor 118, tachometer 120, and conductor 122 for controlling operation of the motor 116. Operation of the motor 116 controls operation of the outflow valve (not shown) to which it is coupled through a gearbox 126 and magnetic clutch 128.The clutch 128 is selectively controlled between its engaged and disengaged positions by a direct current voltage which is supplied from the DC source 96 through a manual/automatic selector switch 130 and conductor 132 to the actuating winding 134 of the clutch.

The position of contact 114c determines whether the controller is in actual control of the cabin pressure rate of change. The entire controller is functioning whether it is operating in the primary or standby mode. The difference is that when the controller is primary and no malfunction has been detected, it will be in control of the cabin pressure change rate due to its contact 114c being closed. The standby controller will be functioning but not controlling the cabin pressure because its corresponding contact is open. When a malfunction is detected or the standby controller is selected manually, the controller will control the cabin pressure even though it is operating in the standby mode.

In operation, the automatic cabin pressure controller 16 utilizes manually selected inputs from the selector panel 10 and inputs corresponding to ambient atmospheric pressure and cabin pressure from the sensors 21 and 23, respectively, to produce an output voltage for controlling position of the outflow valve (not shown).

The cabin altitude function generator 72, which will be subsequently described in detail in connection with Figure 5, receives an input voltage from the atmospheric pressure sensor 21 through conductor 22a which is proportional to ambient atmospheric pressure outside the airplane and utilizes this voltage to compute an output voltage corresponding to a calculated value of cabin pressure. The functional relationship established by the cabin altitude function generator 72 automatically provides a corresponding cabin altitude for every possible aircraft altitude so that no manual selection of cruise altitude by the crew is necessary to effect a proper rate of cabin pressure change, it being understood that the maximum permissible cabin altitude (generally 8,000 ft.) is established at the maximum design altitude for the airplane.This output voltage is altered during descent by a positive DC voltage from descent detector 76.

The descent detector 76 receives the ambient pressure responsive voltage from the atmosphere pressure sensor 21 through conductor 2b and utilizes this voltage to determine when the airplane has begun its final descent in preparation for landing. In the preferred embodiment, this is determined to be when the airplane has descended at a rate of at least 1,000 ft. per minute for a period of one minute. When such has been determined, the descent detector 76 transmits an output voltage to the rate logic 66 through conductor 78, which causes the output to be at full value.This output will be maintained at this level regardless of subsequent changes in the airplane's mode of flight until the airplane has landed, whereupon the ground logic circuit 94 will transmit a resetting signal through conductor 98 to the descent detector 76 in response to closing of the landing gear switch 46.

The positive output voltage of cabin altitude function generator 72 is transmitted through conductor 74 to the rate logic 66 wherein it is added to a negative input voltage proportional to actual cabin pressure, which voltage is received from cabin pressure sensor 23 through conductor 24a. The result of this addition indicates whether the aircraft is ascending or whether it is in the dwell function. If the difference between the actual cabin pressure and the commanded cabin pressure is great, the aircraft is ascending; if the difference is small the aircraft is in the dwell function.Accordingly, if the output of the cabin altitude function generator 72 is sufficiently greater than the output of the cabin pressure sensor 23 so as to indicate that the airplane is in a scheduled climb, the rate logic will produce an output through the conductor 104 which is at a first predetermined level, preferably 100% of the input voltage from conductor 12a. If less than this predetermined difference exists, the rate - limit will be scaled to lesser value, preferably about 50% of the rate limit input. The rate logic 66 is able to determine whether or not the aircraft is descending through a signal from descent detector 76 through conductor 78 which is true when the aircraft is descending and false otherwise.

From the selector panel 10, the rate logic 66 receives an input voltage through conductor 12a which corresponds to a manually selected limit of rate of cabin pressure change or "rate limit". In the preferred embodiment rate logic 66 has two possible outputs, either 100% of the input it receives from selector panel 10 or 50% of the input it receives from selector panel 10. If on the basis of the information that it receives from the sum of the actual cabin pressure and the commanded cabin pressure or from descent detector it determines that the aircraft is either ascending or descending it will pass 100% of the input value of rate limit. If rate logic 66 determines that the aircraft is in the dwell function, it will only pass 50% of the rate limit input from selector panel 10.

The output of rate logic 66 is passed to rate amplifier 90 through conductor 104 to determine at what level rate amplifier 90 will saturate. At a summing node of rate amplifier 90 the output of high altitude discriminator 68 through conductor 88 and the output of cabin pressure sensor 23 through conductor 24c are combined to supply the amplification input.

The selector panel 10 provides an output voltage through conductor 12b to the high altitude discriminator 68 which corresponds to the preselected landing field altitude, corrected for barometric pressure. Also fed to the high altitude discriminator is the output of the cabin altitude function generator 72 which is proportional to the calculated cabin altitude, i.e. the altitude to which the controller 16 should cause the aircraft cabin to reach.

The high altitude discriminator 68 blocks the voltage signal corresponding to the lower of these altitudes and transmits the signal corresponding to the higher altitude to the rate amplifier 90 through conductor 88. The rate amplifier 90 compares this to the signal corresponding to actual cabin pressure received from the cabin pressure sensor 23 through conductor 24c. If comparison of these signals indicates that actual cabin pressure corresponds to an altitude lower than that transmitted by the high altitude discriminator 68, the rate amplifier 90 will not be permitted to produce a command signal calling for descent in cabin altitude.

This feature of the rate amplifier 90 is designed to operate during descent of the aircraft to a landing field having an altitude signficantly greater than sea level, such as the airport at Denver or Mexico City. During descent, actual cabin pressure will be lower than the pressure indicated by the cabin altitude function generator 72, indicating that the cabin is at a higher altitude than that required. As aircraft altitude decreases, cabin altitude will decrease correspondingly. Because the maximum cabin altitude is only about 8,000 ft, it will reach the level of the landing field much before the aircraft reaches that altitude. If descent of the cabin were permitted to continue, it would then be necessary to provide an extended period of depressurization of the aircraft after it had landed.Accordingly, after the desired cabin altitude as indicated by the output of the cabin altitude function generator 72 has reached an altitude equivalent to that selected for the landing field, the high altitude discriminator 68 will prevent further descent of the cabin, supplying the landing field altitude as the command altitude. When actual cabin pressure has reached a barometrically corrected value for this altitude, descent of the aircraft cabin terminates.

An additional input to the rate logic 66 is provided by AP limit logic circuit 80. The circuit receives inputs from the pressure sensors 21 and 23 corresponding to atmospheric pressure and cabin pressure and determines therefrom the pressure differential between inside cabin and outside ambient pressures in a manner which will be described in detail in connection with Figure 3. If this pressure differential is greater than a predetermined value so as to potentially endanger the structural integrity of the plane, an output is produced by the AP limit logic circuit 80 to the rate logic 66 through conductor 92 which increments the output signal of the rate limiter by an amount proportional to the excess pressure differential between atmosphere and cabin pressure.This causes the output of the rate logic to be modified so as to change the saturation level of rate amplifier 90, regardless of the relationship of the commanded altitude to the actual altitude.

The input to rate amplifier 90 from high altitude discriminator 68 through conductor 88 is always positive. High altitude discriminator 68 receives two inputs, one from the selector panel 10 which is the landing field altitude with barometric correction and one from the cabin altitude function generator 72 through conductor 86. Both of these inputs are positive and the discriminator 68 will compare the two to determine which is the higher altitude or lowest pressure. The lowest pressure is the signal that is passed to rate amplifier 90. Cabin pressure sensor 23 always has a negative output thus a negative voltage is passed to rate amplifier 90 through conductor 24c. The summation of the positive high altitude discriminator output and the cabin pressure sensor output determines whether the aircraft is ascending or descending.If the aircraft is ascending, the output voltage of high altitude discriminator 68 which will be the commanded cabin pressure will always be lower than the actual pressure from cabin pressure sensor 23. Therefore, for the ascending function, the input to rate amplifier 90 will be negative voltage. When the aircraft is descending, the commanded cabin pressure from high altitude discriminator will be greater than the actual cabin pressure from cabin pressure sensor 23. Thus, the input voltage to rate amplifier 90 will be positive.

Whether the input voltage is positive or negative determines the ultimate saturation point for rate amplifier 90. In the preferred embodiment the saturation point for negative inputs corresponds to the input from rate logic 66, whereas for positive inputs the saturation point corresponds to 3/7th of the input from rate logic 66. The output of rate amplifier 90 is the rate command which is input to dynamic compensator 100 through conductor 106 where it is combined with the first derivative of the output of cabin pressure sensor 23.

The output of rate amplifier 90 has the reverse sign of the input. Thus, when the aircraft is ascending and the input is negative, the output signal to open the valves (which is the output of the rate amplifier 90) will be positive. When the plane is descending so that an output signal which will tend to close the valves for pressurizing the aircraft cabin is required, the output of rate amplifier 90 will be negative.

The first derivative of the cabin pressure, which is calculated by the dynamic compensator 100, will be negative when the plane is ascending since this will be a decrease in cabin pressure and it will be positive when the plane is descending since this is an increase in cabin pressure. Also input into the summing node of dynamic compensator 100 is the output of ground logic 94 through conductor 102. The output of ground logic 94 which becomes the input to dynamic compensator 100 is one of three states, it is either a positive DC, a negative DC or zero. When the aircraft is required to be prepressurized for takeoff, the voltage input to dynamic compensator 100 is positive DC to override the summation of the first derivative of cabin pressure and the output of rate amplifier 90.When the aircraft has landed and it is required that the aircraft be depressurized, the output of the ground logic 94 to compensator 100 is a negative voltage which again will override the summation of the first derivative of cabin pressure and the output of rate amplifier 90. While the aircraft is in flight the output of ground logic 94 is zero so that the summation of derivative of cabin pressure and the output of rate amplifier 90 are the determining factors for opening or closing the actuator valve. If the aircraft is ascending the output of rate amplifier 90 will be positive and the first derivative of the cabin pressure will be negative. The rate command from rate amplifier 90 is modified by the value of the first derivative to lessen it so that the rate command does not tell the valve to open as fast as if it were unmodified.If the aircraft is descending the output of rate amplifier 90 will be negative and the first derivative of cabin pressure will be positive and again the summation is merely the modification of the rate command from rate amplifier 90.

Dynamic compensator 100 takes the summation of the first derivative of cabin pressure and the rate command of rate amplifier 90, whether modified by inputs from ground logic 94 or not, and passes this signal to driver 110 through conductor 108.

Thus a cabin rate change command is transmitted from the rate amplifier 90 through conductor 106 to the dynamic compensator 100 which conditions the signal to be suitable for transmission to the driver 110. In addition, a signal is transmitted to the dynamic compensator 100 from the ground logic circuit 94 when the landing gear switch is closed indicating that the aircraft has landed to command that the outflow valve (not shown) be fully opened.

The rate command passed to the driver 110 causes this circuit to apply a voltage to the motor 116, assuming contact 114c is closed, suitable to operate the motor in a desired direction and rate of speed. Contact 1 14c is closed when the automatic cabin pressure controller is in control of the cabin pressure or "on". When the switch is open, the controller is deemed to be "off". Operation of the motor causes the tachometer 120 to produce a feedback signal permitting proper control of the motor by the driver.

The electromagnetic clutch 128 is engaged whenever the manual/automatic selector switch 130 is in the automatic position. When this switch is set at manual, actuating winding 134 is de-energized so that the clutch 128 disengages and the valve is no longer controlled by the motor 116. This permits the aircraft crew to manually control the outflow valve by conventional means not shown).

Figure 3 is a block diagram of AP limit logic 80 of Figure 2. Summing node 136 receives a constant voltage input proportional to the maximum allowable differential between ambient atmosphere pressure and cabin pressure through conductor 138. A voltage proportional to the cabin pressure is received from cabin pressure sensor 23 through conductor 24b, and a voltage proportional to ambient atmospheric pressure is transmitted from atmosphere pressure sensor 21 to summing node 136 through conductor 22c. The output of summing node 136 is applied to a maximum AP amplifier 140 through conductor 142. The output of maximum AP amplifier 140 goes through a conductor 144 and bifurcates to conductor 92 which provides the input into rate logic 66 and to conductor 146 which provides input into an inhibit generator 148.Inhibit generator 148 then sends a signal to the automatic transfer circuit through conductor 42c which acts as a AP inhibit signal.

The inputs to summing node 136 from maximum AP reference through conductor 138 and the atmosphere pressure from the atmosphere pressure sensor 21 through conductor 22c are both negative. The input of the cabin pressure from cabin pressure sensor 23 through conductor 24b is positive.

AP limit logic 80 compares these two signals and if the cabin to atmosphere differential exceeds a preselected value for any reason, AP limit logic will increase the preselected rate limit to compensate for the increased climb rate of the aircraft.

In operation, the atmospheric pressure and cabin pressure are constantly monitored by the application to summing node 136 of voltage signals which are proportional to pressure.

By subtracting the atmospheric pressure voltage from the cabin pressure voltage, it obtains a voltage proportional to the differential pressure. From this differential pressure is subtracted a voltage proportional to the allowable AP so that only when the differential pressure exceeds this value will the limit logic supply an increment. The excess differential pressure is then amplified to produce an incremental increase in cabin pressure change rate.

This is applied to the rate logic 66 to increment the rate as previously described.

While AP limit logic 80 is overriding the selected rate limit, it also prevents automatic transfer to the standby controller by issuing a "true" logic signal to the automatic transfer logic of the other controller by conductor 42c, the effect of which will be later explained.

It will be noted that in the preferred embodiment maximum AP amplifier 140 is reverse biased when a negative voltage is applied to the input through conductor 142. Since the ambient atmosphere pressure Pa will always be less than or equal to cabin pressure Pcs a positive voltage will appear at conductor 142 and forward bias the amplifier. However, an output at conductor 144 from maximum AP amplifier 140 is not desired until the differential between the cabin pressure and the atmosphere pressure exceeds a predetermined level. By the introduction of a negative maximum AP voltage through conductor 138 into summing node 136, a positive voltage will not appear at conductor 142 until Pc exceeds the summation of the negative maximum AP reference voltage and the atmosphere pressure voltage.

The result can be seen on Figure 4. The output remains at zero until the input reaches the value designated as A, at which point, Pc is greater than maximum AP reference plus Pa and a positive voltage appears at conductor 142 and forward biases maximum AP amplifier 140.

Maximum AP amplifier 140 is preferably a linear amplifier and produces an output through conductor 144 that increases linearly until it reaches a value designated as C. The output of the amplifier 140 is applied to inhibit generator 148 which informs the automatic transfer circuit that the detected difference from scheduled cabin pressure is not due to controller malfunction. This inhibits transfer. The signal is also applied to the rate logic 66 of controller 16 which increments the rate signal by an amount proportional to the output of the maximum AP amplifier 140.

If the input, for example, is a voltage equal to that shown as D on the graph, the output will correspond to the value noted as E. The output of rate logic 66 will be incremented by the voltage represented by E. C on the output axis is the value where maximum AP amplifier 140 saturates and the output corresponds to an input of B from summing node 136. As the sum of the cabin pressure voltage, the atmospheric pressure voltage and the maximum AP reference voltage exceeds value B, amplification will no longer take place.

In the preferred embodiment, C on the graph would correspond to a voltage proportional to a rate increment of 800 ft. per minute. Therefore the rate, although incremented above the selected rate limit, will never exceed the selected rate limit plus 800 feet per minute.

Referring now to Figure 5, cabin altitude function generator 72 consists primarily of three amplifiers Ul, U7 and U3 and two complementary pairs of transistors, transistor Q1 connected with transistor Q2 and transistor Q3 connected with transistor Q4.

The voltage at the non-inverting input of amplifier U1 is voltage V6 which is controlled by V2 through resistor R1 and the voltage feedback from the output of amplifier U1 through resistor R2. The voltage at the inverting input of amplifier U1 is voltage V7 which is the voltage produced by a feedback loop including the transistors Q1 and Q2.

The output of amplifier U1 is voltage V4. This voltage is applied through resistor R5 to the bases of transistors Q3 and Q4 and through resistor R3 to the bases of transistors Ql and Q2. The emitters of transistors Q1 and Q2 are connected and through resistor R4 feed amplifier U1 and capacitor C1. Capacitor C1 is connected to ground. The collector connection of transistor Q1 is at voltage V3 and the collector connection of transistor Q2 is ground. Voltage V3 is the output voltage of amplifier U3.The amplifier U3 has a resistive negative feedback network R9, R10, R1 1 and R13, which receives a negative bias voltage at conductor 42a. which is the inverse of positive voltage V2 at conductor 42at. The voltage at the inverting input of amplifier U3 is composed of voltage V1 through conductor 22a2 which passes through resistor R12, and the output voltage of resistive network R9, R10, R11 and R13. The non-inverting input of amplifier U3 is grounded.

Voltage V4 also feeds the two complementary transistors Q3 and Q4 through resistor R5 which bifurcates and feeds the bases of Q3 and Q4. The collector of Q4 is connected to ground and the collector of Q3 is connected to conductor 22al which supplies voltage Ví which is a voltage proportional to the atmospheric pressure from the atmosphere pressure sensor 21. The emitters of transistors Q3 and Q4 are connected together and supply voltage V8 which passes through resistor R6 and over capacitor C2 to supply the non-inverting input of amplifier U2. The voltage at the inverting input of the amplifier U2 is determined by the feedback loop from the output of U2 through potential divider R7 and R8. The output of amplifier U2 is V0.

V0 is the output at conductor 74 to rate logic 66 and to high altitude discriminator 68 through conductor 86. V1 is the input that comes in through conductor 22al from atmosphere pressure sensor 21. V1 is a voltage proportional to the ambient atmospheric pressure. V2 is a constant positive DC voltage used as a biasing voltage which comes from the power supply 70. Vl is input at conductor 22al and -V1 is input at conductor 22a2. V2 is input at conductor 42a1 and -V2 is input at conductor 42a When V6 is more positive than V7, the output voltage V4 of amplifier U1 will be a positive voltage which will travel through resistor R3 to the bases of the transistors Q1 and Q2.

When V4 iS positive, NPN transistor Q1 will be on and PNP transistor Q2 will be off. When transistor Otis on, voltage V3 will pass through the transistor and appear as voltage Vi at the emitter connection to resistor R4. R4 is part of the RC charging network with capacitor C1 and voltage V7 at the inverting input terminal of amplifier U1 will increase exponentially towards V3.

When V, rises to a value slightly greater in value than voltage V6, the output V4 of amplifier U1 will become negative. When V4 is negative it will feed back through resistor R2 to voltage V6 and cause voltage V6 also to become slightly less positive. V4 will pass through resistor R3 to the bases of the Q1, Q2 complementary pair and will turn transistor Q1 off and will turn transistor Q2 on. When transistor Q2 is on, capacitor C1 will discharge towards zero through resistor R4 since transistor Q2 has its collector connected to ground.

When V7 discharges to a value slightly less than the new value of V6, the output V4 of amplifier U1 will again swing positive. When V4 goes positive it will make V6 slightly more positive again through resistor R2; in other words, the amplifier U1 exhibits backlash or Schmitt action. V4 will also pass through resistor R3 again to the complementary pair Q1, Q2 and turn transistor Q1 on and transistor Q2 off and repeat the cycle as previously described.

Thus voltage V4 will appear as a square wave. Voltage V4 also passes through resistor R5 and feeds the base connections of the NPN transistor Q3 and PNP transistor Q4. When V4 is negative, transistor Q4 will be conducting and the output of transistor Q4 will be zero since its collector is grounded. When voltage V4 is positive, transistor Q3 will be conducting and will pass voltage V1 through transistor Q3 to resistor R6. Thus, voltage V8 at resistor R6 will have the same relationship to voltage V1 as voltage V5 has to voltage V3. Resistor R6 and capacitor C2 filter voltage Vs which is the non-inverting input to amplifier U2.

Amplifier U2 scales voltage V8 so that the proper voltage V0 is output at conductor 74 which is applied to the rate logic 66 and conductor 86 which is applied to high altitude discriminator 68.

Voltage V3 is the output of amplifier U3 and will remain as a positive output as long as the input to amplifier U3 at its inverting input terminal is less than zero. The voltage at this terminal is determined by the voltage -Vl, which is the inverse of the output of atmosphere pressure sensor 21, and is connected to the terminal through resistor R12, and by the voltages applied to the resistive feedback network R9, R10, R11, and R13. In addition to the bias voltage -V2 and the output voltage V3, the feedback network is also influenced by the voltage on the line 84, which is the output of the descent detector 76.When this voltage is negative, indicating that the aircraft is ascending or in level flight, a diode D1 connected between the junction of resistors R9 and R10 and the line 84 is forward biased, and the voltage of this junction point is therefore clamped. A diode D2 which connects the junction point to the inverting input of the amplifier U3 is therefore reverse biased, so that resistors R9 and R10 have no effect on the response of the amplifier U3, whose gain and bias are now controlled entirely by the resistors R11, R12 and R13.

If the aircraft now descends, the voltage on the line 84 from the descent detector will become positive, reverse biasing the diode D1 and unclamping the voltage at the junction between the resistors R9 and R10. The relative values of the resistors R9, R10, R11 and R12 are such that the diode D2 will now be forward biased, so that the bias and gain of the amplifier U3 are now controlled by all the resistors in the feedback network, as well as R12.

Under these conditions, the bias and gain are both reduced.

Figure 6 is a series of graphs labeled 6A-6C which show the various voltage relationships associated with cabin altitude function generator 72. Graph 6A shows the relationship of voltages V3, V6, V2 and V7. As can be seen, voltage V6 swings around voltage V2, because of the effect of positive feedback resistor R2. Voltage V7 will exponentially charge towards voltage V3 to a value slightly greater than V6 whenever transistor Q1 is on. When Q1 is turned off and Q2 is turned on, voltage V7 will exponentially discharge towards zero to a value slightly less than voltage V6 which has shifted to the negative side of V2 due to the feedback through resistor R2. When V7 becomes less positive than V6, the output of amplifier U1 goes positive, turns Q2 off and Q1 on thus repeating the cycle.Graph 6B shows voltage V4 and its respective values when Q1 is on and Q2 is off and also when Q1 is off and Q2 is on. Graph 6C shows the relationship of V3 and V5. V5 is equal to V3 when Q1 is on and Q2 is off. When Q2 is on V5 equals the collector voltage of Q2 which is at ground.

It will be noted that if the circuit is scaled so that V7 is nearly linear, then the average value of V7 is equal to the mean of the two values alternately adopted by V6; this mean will be referred to as V6,. The average value of V, is equal to the average value of V5 or V, (ave) = V5 (ave) Thus the following relationships are established:: V6, = V7 (ave) The average value of V5 is equal to V3 multiplied by the time period that Q1 is on (t1) divided by the time period Q1 is on plus the time period when Q2 is on (t2) or t1 V5 (ave) = V3 t + t2 since V7 (ave) = V6' and V7 = V5 (ave) then V6, = V5 (ave) and V6' = V3 ti + t2 or V6, = t1 V3 t1 + t2 The time period that Q1 is on divided by the time period Q1 is on plus the period Q2 is on is the duty cycle of V4 or t1 the duty cycle of V4 = t1 = V6' V3 Since transistor Q3 and Q4 are driven by V4 in the same manner as Q1 and Q2, V8 bears the same relationship to V1 as V5 bears to V3

v5 (ave) Vl ( t3t3+ t ) since the duty cycle is determined by V4, t3 = t1 for similar transistors.

t 3 + t4 t1 + t2 Therefore,

V8 (ave) = Vl(V6' ) since V0 = R1V8 (ave) then

since vl QPa and V3 - K2 V1 + x3 eK2 Pa + Kc and V6 is a constant therefore

which gives an output voltage dependent on the sensed atmospheric pressure.

Figure 7A is a graph of the function generated by the preferred embodiment of cabin altitude function generator 72. Curve AB is the function used during climb and cruise, when the resistors R9 and R10 are prevented from affecting the amplifier U3 by the clamping action of the diode D1. Curve AC is the function used during descent when diode D1 is reverse biased. In the preferred embodiment curve AB is generated on the basis of keeping the cabin pressure rates at a minimum when the aircraft is ascending at its maximum rate without exceeding the maximum AP for which the aircraft is designed. For curve AB, Pc as generated by the function is always greater than Pa as long as Pa is less than or equal to X.

For Pa greater than X, Pc becomes less than Pa. Under these conditions it can be seen that when attempting to land at an altitude which has a corresponding pressure greater than X, the cabin pressure would try to be less than the ambient pressure. This would be the equivalent of commanding the outflow valve to open so that the inside cabin pressure could go below the outside ambient pressure which is an impossible condition. This would result in the outflow valve being fully opened with no cabin pressure rate control. In order to avoid this, resistors R9 and R10 are introduced on descent influencing the denominator of the cabin pressure function generated in the preferred embodiment by making the negative voltage input to amplifier U3 less negative thus lowering the value of voltage V3.The function of the preferred embodiment cabin altitude function generator is Pa Pc = K2Pa + K3 .

Since the value of K2 Pa + K3 is proportional to V3, lowering of V3 will lower the denominator and thus give a curve as depicted by curve AC. The situation where Pa is greater than point X or the altitude of the aircraft is lower than the altitude corresponding to point X, does not present a problem on takeoffs due to the fact that the aircraft ambient pressure Pa will be decreasing faster than the cabin pressure and operation on the curve AB will never occur and Pa will always be less than P Figure 7B shows the relationship between the cabin pressure (Pc) and the ambient aircraft pressure (Pa) of previous automatic cabin pressure controllers. The ideal relationship as shown in Figure 7A by curve A-B was attempted through straight line approximations DE, EF and FG.Although previous attempts consist of many straight line segments to approximate the ideal curve, three straight line segments are depicted for demonstration. As can be seen, two deficiencies occur in this method of constructing a cabin pressure versus ambient pressure curve. The first is that the straight line approximations do not quite achieve the curve as desired; the second is that through the use of straight line approximations, points of singularity occur at every joining of the straight line segments. The points of singularity do not appear to be far reaching in effect on the curve Pc versus Pa. However, when the rate of change of cabin pressure (the first derivative of Pc) is calculated, points of singularity appear as spikes on what should appear as a straight line for a constant rate of change of cabin pressure.

Referring now to Figure 8, the automatic transfer circuit 34 and associated components are illustrated in detail, it being understood that automatic transfer circuit 32 has an identical set of components similarly connected.

Malfunction detection logic 150 is connected to power supply 70 of the automatic controller 16 through conductors 42a and 42b and is connected to cabin pressure selector 10 through conductors 43a and 43b. Malfunction detection logic 150 is also connected to a valve switch 124 through conductor 41, AP limit logic 80 through conductor 42c, throttle switch 56 through conductor 64a, landing gear switch 46 through conductor 54a, automatic/manual transfer reset 154 through conductor 156, automatic transfer inhibit 158 through both conductors 160 and 162, to primary/standby status memory 164 through conductor 166, to standby off logic 168 through conductor 170, to OR gate 172 through conductor 174, and to automatic transfer circuit 32 through conductor 36a.

The connections in automatic transfer circuit 32 from conductors originating in automatic transfer circuit 34 are shown as dotted lines adjacent to the conductor originating in automatic transfer circuit 34 and included within a single circle. With this in mind it can be seen that malfunction detection logic 150 of automatic transfer circuit 34 is connected to automatic transfer inhibit 158 and OR gate 176, both of automatic transfer circuit 32 through conductor 36a.

OR gate 176 of automatic transfer circuit 34 is connected with successive flight transfer 206 through conductor 177, to standby off logic 168 through conductor 178, and to interconnect logic 180 through conductor 182 on its input side.

OR gate 176 is connected to on-off control 184 through conductor 186 on its output side.

On-off control 184 is connected to automatic cabin pressure controller 17 through conductors 44a and 44b, to interconnect logic 180 through conductors 188 and 190, to standby off logic 168 through inverter 192 via conductor 194, to primary/standby status memory 164 through conductor 196, to primary on latch 198 through conductor 200, to successive flight transfer 206 through conductor 208, to AND gate 210 through conductor 212, and to interconnect logic 180 of automatic transfer circuit 32 through conductor 36d.

The output side of OR gate 172 is connected to on-off control 184 through conductor 214.

In addition to being connected to malfunction detection logic 150 through conductor 174, the input side of OR gate 172 is also connected to successive flight transfer 206 through conductor 216, primary on latch 198 through conductor 218, and interconnect logic 180 through conductor 220.

Interconnect logic 180 is connected to on-off control 184 of automatic transfer circuit 32 through conductor 40d. Primary on latch 198 is connected to standby off logic 168 of automatic transfer circuit 32 through conductor 40b, so standby off logic 168 of automatic transfer circuit 34 is connected to primary on-latch 198 of automatic transfer circuit 32.

Standby off logic 168 also is connected to primary/standby status memory 164 through conductor 222 and to selector panel 10 through inverter 223 via conductor 43c.

Primary/standby status memory 164 has dual connections to successive flight transfer 206 through conductors 224 and 226. Primary/standby status memory 164 is connected to AND gate 210 through conductor 228, to primary lamp 230 through conductor 232, to exclusive OR gate 234 of automatic transfer circuit 34 through conductor 236, and to exclusive OR gate 234 of automatic transfer circuit 32 through conductor 36c.

Since the circuits of automatic transfer 32 and 34 are identical, exclusive OR gate 234 of automatic transfer circuit 34 receives an input from primary/standby status memory 164 of automatic transfer circuit 32 through conductor 40c. On its output side, exclusive OR gate 234 is connected to auto/manual transfer reset 154 through conductor 238. Auto/manual transfer reset 154 is connected to cabin pressure selector 10 through conductor 240 and to auto transfer inhibit 158 through conductor 242. Auto transfer inhibit 158 is connected to successive flight transfer 206 through conductor 244 and to cabin pressure selector 10 through conductor 246. Successive flight transfer 206 is connected to throttle switch 56 through conductor 64b and to landing gear switch 46 through conductor 54b.

During each flight one controller is operated in the primary mode and one controller is operated in the standby mode. When the aircraft lands, successive flight transfer 206 will switch modes of the controllers for the next flight. The specific details of connections and operation of successive flight transfer 206 will be discussed in connection with Figure 11.

However, its output is of immediate interest. If automatic cabin pressure controller 17 is "on", successive flight transfer 206 of automatic transfer circuit 34 will issue a true signal to primary/standby status memory 164 through conductor 226 and to OR gate 176 through conductor 177 upon closing of the landing gear switch 46 at aircraft touchdown. Any true input at OR gate 176 is passed through conductor 186 to on-off control 184. On-off control 184 then energizes coil 114b of control switch 114 which opens contact 114c between driver 110 and motor 116 (see Figure 2). If automatic cabin pressure controller 17 is off, successive flight transfer 206 will issue a true signal to primary/standby status memory 164 through conductor 224 and to OR gate 172 through conductor 216 upon closure of the landing gear switch 46.Any true input at OR gate 172 is passed through conductor 214 to on-off control 184. On-off control 184 then energizes coil 114a of control switch 114 which closes contact 114c between driver 110 and motor 116 (see Figure 2).

Successive flight transfer 206 receives an input through conductor 208 from on-off control 184 to indicate whether or not the controller is on. Successive flight transfer 206 also programs the primary/standby status memory 164 and instructs it as to which mode automatic cabin pressure controller 17 will operate for the next flight. If a true signal is issued to primary/standby status memory 164, through conductor 226, it will indicate that automatic cabin pressure controller 17 is going to be operating in the standby mode. If a true signal is issued through conductor 24 that will indicate that automatic cabin pressure controller 17 will be operating in the primary mode for the upcoming flight.

When automatic cabin pressure controller 17 is operating in the standby mode, automatic transfer circuit 34 is monitoring the performance of cabin pressure of automatic cabin pressure controller 16. If cabin pressure controller 16 malfunctions, automatic transfer circuit 34 will switch control to automatic cabin pressure controller 17. When automatic cabin pressure controller 17 is operating in the primary mode, automatic transfer circuit 34 is disabled and automatic transfer circuit 32 monitors the performance of cabin pressure controller 17. Automatic transfer circuit 32 will switch control to automatic cabin pressure controller 16 if automatic cabin pressure controller 17 malfunctions.

Malfunction detection logic 150 receives inputs from controller 16, cabin pressure selector panel 10, throttle switch 56, landing gear switch 46 and the valve switch 124; the specific details of which will be discussed in connection with Figure 9. On the basis of the inputs which malfunction detection logic 150 receives, it is able to determine such information as whether the aircraft is in flight, whether it is ascending or descending, whether the cabin pressure change rate exceeds the selected rate, whether the AP limit logic 80 is controlling the cabin pressure rate of change, whether a flow problem exists in the aircraft, and whether the primary automatic cabin pressure controller is adequately powered.

The normal output of malfunction detection logic 150 at conductor 174 is a false or zero output. Malfunction detection logic 150 of the preferred embodiment is set or its output at conductor 174 goes true if, while the aircraft is in flight, either the primary automatic cabin pressure controller loses power, or the cabin pressure is ascending at a rate in excess of the selected rate and neither the AP limit logic 80 is controlling nor does a flow problem exist, or the aircraft cabin altitude is descending at a rate greater than the selected rate. A true output from the malfunction detection logic 150 at conductor 174 simultaneously turns automatic cabin pressure controller 17 on and automatic cabin pressure controller 16 off.

Automatic cabin pressure controller 17 is turned on by the true signal passing to OR gate 172 through conductor 174 which in turn activates on-off control 184 as previously described. Automatic cabin pressure controller 16 is turned off by passing the true signal to automatic transfer circuit 32 through conductor 36a. The connections in automatic transfer circuit 32 can be seen by dotted line 40a which shows the identical connections between the malfunction detection logic of automatic transfer circuit 32 and automatic transfer circuit 34. The true signal is received by OR gate 176 of automatic transfer circuit 32 through conductor 40a. It is then passed to its on-off control 184 through conductor 186 which turns its controller off as previously described.

Not only does malfunction detection logic 150 turn one controller off and one controller on, it also sets automatic transfer inhibit 158 of automatic transfer circuit 34 and an identical automatic transfer inhibit in automatic transfer 32. Again, the connection in automatic transfer circuit 32 can be seen in dotted conductor 40a which comes from automatic transfer circuit 32. By setting automatic transfer inhibit, automatic transfer back to the malfunctioning controller is blocked, the malfunction detection logic 150 is disabled and the successive flight transfer circuit 206 is inhibited so that the status of the controllers will remain the same as when the malfunction occurred.

When automatic transfer inhibit 158 is set, it will relay a signal to the cabin pressure selector panel through conductor 246 which will illuminate a light 3 (see Figure 12) indicating to the flight crew that automatic transfer function is locked out. Additionally, the fact that the standby is in control of the cabin pressure rate of change is annunciated to the crew through AND gate 210, which receives a true signal from the primary/standby memory 164 through conductor 228 when the controller is in the standby mode. In addition to the signal that it receives through conductor 228, AND gate 210 also receives a signal through conductor 212 which comes from on-off control 184 which is true when the controller 17 is on. Whenever the controller 17 is standby and on, the output of the AND gate 210 will go to cabin pressure selector panel 10 and illuminate a light 2 (Figure 12) which indicates that the standby is in control of the cabin pressure.

Provision is made on cabin pressure selector panel 10 for the manual selection of either the primary or the standby controller. When toggle switch 1 (see Figure 12) is moved to the standby position, it will issue a true signal through conductor 43b which will set malfunction detection logic 150 and malfunction detection logic will issue a true signal through conductor 174. This true will also issue through conductor 43c to inverter 223 which is connected to standby off logic168. The inverter will change the true signal to a false and will have no effect on standby off logic.

When the toggle switch is moved to the primary position a false signal is issued through conductors 43b and 43c. This false signal will arrive at malfunction detection logic 150 and will not change its output. However, a false signal through conductor 43c will go to inverter 223, causing a true signal to be issued to standby off logic 168, setting it so that its output will be a true. When standby off logic 168 is set and is issuing a true signal as its output, the true signal will reset malfunction detection logic 150 and its output at conductor 174 will be a false. The true signal will then pass to OR gate 176 through conductor 178 which turns the automatic cabin pressure controller 17 off as previously described.

This true signal is also applied to automatic transfer circuit 32 through conductor 36b.

This connection can be seen by dotted conductor 40b which shows the connection between standby off logic of automatic transfer circuit 32 to primary on latch 198 of automatic transfer circuit 34.

When automatic cabin pressure controller 16 is turned off by on-off control 184, the absence of a true signal is communicated through conductor 194 to inverter 192. Inverter 192 then issues a true output to standby off logic 168 which resets standby off logic.

Resetting standby off logic 168 changes its output to zero or a false signal thus completing the loop and returning standby off logic to its normal state. Primary on latch 198 is set by a true signal from standby off logic of automatic transfer circuit 32 and will issue a true signal to OR gate 172 which will turn automatic cabin pressure controller 17 on as has been previously described. When automatic cabin pressure controller system 17 is turned on by on-off control 184, a true signal is issued which will return to primary on latch 198 through conductor 200 and reset the primary on latch again to its normal state which is no output.

If automatic transfer has taken place and the flight crew wishes to return control of the cabin pressure rate to a primary controller, this can also be done by pushing the reset switch in cabin pressure selector panel 10. Depressing the reset button will issue a signal to automatic/manual transfer reset 154 which will then pass to malfunction detection logic 150 through conductor 156 and automatic transfer inhibit 158 through conductor 242. Resetting the malfunction detection logic 150 will cause the output at conductor 174 to go to zero or false. Resetting the automatic transfer inhibit 158 will remove the previously established inhibit and will extinguish the illumination of transfer lock out in the cabin pressure selector panel 10.Resetting automatic transfer inhibit 158 will also remove the inhibit from successive flight transfer 206 through conductor 244 and from malfunction detection logic 150 through conductor 160. Thus while use of toggle switch 1 (Figure 12) alone will restore control to the primary controller, a transfer lockout switch, which may be incorporated into the transfer lockout light 3, may be activated to reset the automatic monitoring and transfer features associated with the standby controller.

When primary/standby status memory 164 is instructed by successive flight transfer 206 as to whether automatic cabin pressure controller 17 will be primary or standby, it will issue a signal to exclusive OR gate 234 through conductor 236. A false input into exclusive OR gate 234 indicates that automatic cabin pressure controller 17 is in a standby mode. A true input indicates primary mode. This true signal is also passed to automatic transfer circuit 32 through conductor 36c. The connections within automatic transfer circuit 32 can be seen by the identical connections within automatic transfer circuit 34 as shown by dotted conductor 40c whih is input into exclusive OR gate 234. Exclusive OR gate 234 receives the output coming from primary/standby status memory of automatic transfer circuit 32 through conductor 40c.

If both inputs to exclusive OR gate 234 are either true or false indicating that both controllers are either in the primary or standby mode, exclusive OR gate 234 will issue a false signal to the automatic/manual transfer reset 154 through conductor 238. Automatic/ manual transfer reset 154 will then reset the malfunction detection logic 150 and automatic transfer inhibit 158. By resetting malfunction detection logic 150 it will indicate that, whatever its output is, it should be reset to zero. By this, the controllers are prevented from simultaneously operating in either the primary or in the standby mode.

Primary/standby status memory 164 will also issue a true signal to primary lamp 230 through conductor 232 if automatic cabin pressure controller 17 is primary for this flight.

This lamp indicates to the flight crew which cabin pressure controller is operating in the standby mode so that if a malfunction is noted by the flight crew, the maintenance crew will know which automatic cabin pressure controller malfunctioned. Operation of automatic transfer inhibit 158 upon malfunction prevents the change of primary controller by successive flight transfer 206 upon landing and maintains illumination of primary lamp 230 for repair identification.

In order to prevent a situation where both automatic cabin pressure controllers are either on or off, an interconnect logic 180 is provided, the detailed operation of which will be discussed in connection with Figure 11. Interconnect logic 180 receives inputs from on/off control 184 of automatic transfer circuit 34 and on-off control of automatic transfer circuit 32. If automatic cabin pressure controller 17 is on, interconnect logic 180 will receive a true input through conductor 190, if it is off it will receive a true input through conductor 188.

Whether automatic cabin pressure controller 17 is on is also issued as a true signal to the interconnect logic of automatic transfer circuit 32. The connections in automatic transfer circuit 32 can best be seen through dotted conductor 40d which connects interconnect logic 180 with on-off control of automatic transfer circuit 32.

Interconnect logic 180 will evaluate these inputs and determine whether both controllers are on or both controllers are off. If both controllers are on, interconnect logic 180 will issue a true signal to OR gate 176 through conductor 182 which, in turn, will turn the automatic cabin pressure controller off as has been previously described. If both controllers are off, interconnect logic 180 will issue a true signal to OR gate 172 through conductor 220 which, in turn, will turn automatic cabin pressure controller 17 on as has been previously described.

Referring now to Figure 9, the basic malfunction monitoring and switchover circuitry of the malfunction detection logic 150 of automatic transfer circuit 34 is shown in detail, it being understood that malfunction detection logic of automatic transfer circuit 32 has an identical set of components similarly connected. Additional inputs described in connection with Figure 8 for switching between primary and standby controllers are not duplicated herein.

Power loss detector 248 is connected to the power supply of automatic cabin pressure controller 16 through conductor 42a and the input side of OR gate 250 through conductor 252. The input side of OR gate 250 is also connected to rate monitor circuit 254 through conductor 256 and to AND gate 258 through conductor 260. Rate monitor circuit 254 is connected to automatic cabin pressure controller 16 through conductor 42b and to cabin pressure selector panel 10 through conductor 43a.

AND gate 258 is connected to valve switch 124 through conductors 262 and 41 (Figure 8) and to the inhibit generator of AP limit logic of automatic cabin pressure controller 16 (Figure 1) through inverter 264 through conductor 42c and to rate monitor circuit 254 through conductor 265. The output side of OR gate 250 is connected to the input side of AND gate 266 through conductor 268. The input side of the AND gate 266 is also connected to the output side of OR gate 270 through conductor 272. The input side of OR gate 270 is connected to throttle switch 56 through conductor 64a and to inverter 274 through conductor 276. Inverter 274 is connected to landing gear switch 46 through conductor 54a.

The output side of AND gate 266 is connected to malfunction switch over control 278 through conductor 280. Malfunction switch over control 278 is connected to OR gate 172 (Figure 8) through conductor 174.

The input side of AND gate 282 is connected to rate monitor circuit 254 through conductor 284 and to inverter 286 through conductor 288. Inverter 286 is connected to valve switch 124 through conductors 290 and 262. The output side of AND gate 282 is connected to flow light 7 (Figure 12) through conductor 294.

Malfunction switch over control 278 will issue a true signal to OR gate 172 (see Figure 8) through conductor 174 which will turn off automatic cabin pressure controller 16 and turn on automatic cabin pressure controller 17 when it receives a true signal from AND gate 266 through conductor 280. AND gate 266 will issue a true output only when the input from OR gate 250 through conductor 268 and the input from OR gate 270 through conductor 272 are both true.

The output from OR gate 270 indicates whether or not the aircraft is in flight. If the aircraft is in flight the output will be true, if the aircraft is on the ground the output will be false. Flight is indicated when either one of two inputs to OR gate 270 is true. The first input from throttle switch 56 will be true when the throttle is in the advanced position. The second input through conductor 276 from inverter 274 will be true when the signal from landing gear switch 46 through conductor 54a to inerter 274 is false. The landing gear switch 46 issues a false signal when the landing gear switch is open. When the aircraft is on the ground, landing gear switch will be closed and will be issuing a true signal. This true signal will arrive at inverter 274 and a false signal will be issued to OR gate 270 through conductor 276. When the aircraft is in the air the landing gear switch will be open and a false signal will be issued to inverter 274. When a false signal is received by inverter 274 a true signal is passed to OR gate 270 through conductor 276 thus, when the aircraft is in the air a true signal will be issued by OR gate 270 to AND gate 266 through conductor 272. Input from the throttle switch provides a flight signal when the aircraft is in the process of taking off.

Whether the aircraft is in flight acts as an inhibit since a true input will be present when the aircraft is in flight and a true input is necessary before AND gate 266 will issue a true output to malfunction switch over control 278.

OR gate 250 will issue a true output to AND gate 266 through conductor 268 when either of its three inputs is true. One input to OR gate 250 is from power loss detector 248. Power loss detector 248 receives a power signal from automatic cabin pressure controller 16. As long as power from the power supply is within a specified range in automatic cabin pressure controller 16 the output of power loss detector 248 is false. When power deviates from the specified range in automatic cabin pressure controller 16, power loss detector 248 issues a true signal to OR gate 250 through conductor 252. This true signal will be passed to AND gate 266 through conductor 268 which, in turn, will issue a true signal to malfunction switch over control 278 if a true signal has been received from OR gate 270 indicating that the aircraft is in flight.

A second input to OR gate 250 comes from rate monitor circuit 254. Rate monitor circuit 254 receives an input of the sensed cabin pressure climb rate from automatic cabin pressure controller 16 through conductor 42b and the selected rate limit from cabin pressure selector panel 10 through conductor 43a. Rate monitor 254 compares these two inputs and determines whether the cabin pressure change rate exceeds the selected rate limit. If the aircraft is descending and the sensed rate exceeds the selected rate limit it will issue a true signal to OR gate 250 through conductor 256. This true signal will be passed by OR gate 250 to AND gate 266 through conductor 268 as has been previously described.

The third input to OR gate 250 of the preferred embodiment comes from AND gate 258.

AND gate 258 receives three inputs, all of which must be true before it will issue a true signal to OR gate 250. The first input is from rate monitor circuit 254. When rate monitor circuit 254 determines that the sensed cabin pressure change rate exceeds the selected rate limit and the aircraft is ascending, it will issue a true signal to AND gate 258 through conductor 265.

The second input to AND gate 258 is from valve switch 124. When the valve switch is closed it will issue a false signal. A true signal will be issued by the valve switch 124 when the outflow valve is open.

The third input to AND gate 258 is from AP limit logic of automatic cabin pressure controller 16. Under normal conditions the inhibit generator of AP limit logic will be issuing a false signal which will be input to inverter 264 through conductor 42c. The output of inverter 264 under normal conditions will be a true signal and will not block the output of the AND gate 258. When the AP limit logic is incrementing the cabin pressure rate of change it will also issue a true signal to inverter 264 through conductor 42c. The true signal arriving at inverter 264 will be passed to AND gate 258 as a false signal thus inhibiting the output of AND gate 258.The output of AND gate 258, when true, indicates that the cabin is ascending at a rate greater than the preselected rate limit and the output flow valves are not completely closed and that the AP limit logic is not controlling the cabin pressure rate of change.

Malfunction switchover control 278 will issue a true signal indicating that automatic cabin pressure controller 17 should be turned on only when the plane is in flight and either the automatic cabin pressure controller 16 has lost power, or the cabin is descending at a rate greater than the selected rate, or the cabin is ascending at a rate greater than the selected rate limit while the AP limit logic is not in control and the outflow valve is not completely closed.

Malfunction detection logic 150 performs an annunciation function in addition to detecting a malfunction in the primary controller. The output of rate monitor circuit 254 bifurcates and one output goes to AND gate 258 and the other to AND gate 282 through conductor 284. AND gate 282 also receives an output from inverter 286 through conductor 288. The input to inverter 286 comes from valve switch 124. As was said earlier, valve switch 124 issues a true signal when the outflow valve is open and a false signal when the valve is closed. When the outflow valve is closed, the false signal is input to inverter 286 which will pass a true signal to AND gate 282 through conductor 288.AND gate 282 will issue an output through conductor 294 to cabin pressure selector panel 10 illuminating flow light 7 when it receives a signal from rate monitor circuit 254 that the altitude of the cabin is ascending at a rate greater than the selected rate limit and the outflow valve closed. Thus, illumination of flow light 7 indicates to the flight crew that the cabin pressure is decreasing while the outflow valves are fully closed. This means that there is a leak on the aircraft of a magnitude greater than the cabm inflow and that the failure to properly pressurize is not due to a faulty automatic cabin pressure controller. The crew is then aware that it must either plug the leak or increase inflow to permit proper control of cabin pressure.

Referring now to Figure 10, the detailed connections of successive flight transfer circuit 206 are illustrated. The throttle switch 56 is connected via conductor 64b2 to an AND gate 296, to AND gate 298 through inverter 300 via conductor 64by, and to a latch 302 through conductor 304. The landing gear switch 46 is connected via conductor 54b2 to AND gate 296, to AND gate 298 through conductor 54by, to AND gate 306 through conductor 308, and to inverter 310 through conductor 312. Inverter 310 is connected to a latch 314 through conductor 316.

The input side of AND gate 306 is also connected to latch 314 through conductor 318 and the inverter 320 of AND gate 306 is connected to automatic transfer inhibit 158 through conductor 244. The output side of AND gate 306 is connected to landing gear switch control 322 through conductor 324. Landing gear switch control 322 is connected to primaryl standby status memory 164 through conductors 224 and 226, to OR gate 172 through conductor 216, and to OR gate 176 through conductor 177 (see Figure 8). Landing gear switch control 322 is also connected to on-off control 184 through conductor 208. Latch 314 is connected to OR gate 326 through conductor 328. OR gate 326 is connected to latch 302 through conductor 330 and to sixty second delay 332 through conductor 334. Latch 302 is connected to twenty second delay 336 through conductor 338.Twenty second delay 336 is connected to the output side of the AND gate 298 through conductor 340. Sixty second delay 332 is connected to the output side of AND gate 296 through conductor 342.

AND gate 298 and AND gate 296 both receive two inputs. One input is from the landing gear switch 46 and the other input is from the throttle switch 56. Landing gear switch 46 will issue a true signal when it is closed, indicating that the aircraft is on the ground. Throttle switch 56 will issue a true signal when it is advanced, indicating that the aircraft is taking off or is in flight. The input to AND gate 298 from throttle switch 56 must pass through inverter 300 before it becomes an input to AND gate 298 thus when the aircraft is on the ground, the input to AND gate 298 is true. When the aircraft is in flight or making its take-off run, the input to AND gate 296 is true.Since neither the input to AND gate 298 through conductor 54b, nor the input to AND gate 296 through conductor 54b2 from landing gear switch 46 passes through an inverter, both AND gates 296 and 298 receive true inputs only when the aircraft is on the ground. If landing gear switch 46 closes, it issues a true signal to AND gate 298 through conductor 54b, and to AND gate 296 through conductor 54b2.If the throttle switch 56 is not in the advanced position, indicating that the aircraft has landed, it issues a false signal to AND gate 296 through conductor 64b2 and to inverter 300 through conductor 64bl. AND gate 296 will have a false output while AND gate 298 will have a true output since the false input to inverter 300 results in a true input to AND gate 298. This condition indicates that the aircraft has landed and the controller should switch modes.When AND gate 298 has a true output, it is passed to twenty second delay 336 through conductor 340.

The twenty second delay is to eliminate transfers due to bounce on landing. After the twenty second delay a true signal is passed to latch 302 through conductor 338. Latch 302 is set and will issue a true output to OR gate 326 through conductor 330. OR gate 326 will then issue a true signal to set latch 314 through conductor 328. Latch 314 then issues a true signal to AND gate 306 through conductor 318 which passes to landing gear switch control 322 through conductor 324. Landing gear switch control 322 then switches modes between the controllers as described in connection with Figure 8.

During take off, throttle switch 56 is advanced issuing a true signal to AND gate 296 and, through inverter 300, a false signal to AND gate 298. Thus the output of AND gate 298 becomes a false and the output of AND gate 296 becomes a true. The output of AND gate 296 passes to sixty second delay 332 through conductor 342. The true signal from throttle switch 56, when advanced, will reset latch 302 through conductor 304 and the output of latch 302 will become a false to OR gate 326 through conductor 330. Latch 302 is reset so that the transfer can again take place the next time the aircraft lands. If landing gear switch 46 does not open within 60 seconds after advanced throttle, sixty second delay 332 will issue a true to OR gate 326 through conductor 334 which will be passed to set latch 314 through conductor 328.

Assuming the normal take off, landing gear switch 46 will open thus issuing a false input to AND gate 296 through conductor 54b2 and the AND gate 298 through conductor 54bl.

However, the false signal issued by landing gear switch 46 passes to inverter 310 through conductor 312 which in turn issues a true signal to reset latch 314 through conductor 316 and causes the output of latch 314 to AND gate 306 through conductor 318 to return a false. The output of landing gear switch is received by AND gate 306. If the aircraft is in flight a false signal is received and successive flight transfer is blocked. Automatic transfer inhibit 158 will issue a true signal to inverter 320 through conductor 244 when the automatic transfer inhibit has been activated. A true signal to inverter 320 will be passed to AND gate 306 as a false signal, preventing a true output by AND gate 306. Thus while automatic transfer inhibit is issuing a true signal no landing gear switch transfer will take place.

Referring now to Figure 11, the interconnect logic 180 of automatic transfer circuit 34 and automatic transfer circuit 32 are shown in detail along with their interconnections. Since the internal connections of interconnect logic 180 of automatic transfer circuit 34 and the interconnect logic of automatic transfer circuit 32 are identical, the connections of interconnect logic 180 of automatic transfer circuit 34 only will be described.

The input side of AND gate 344 is connected to on-off control 184 through conductor 190 which is in turn connected to terminal 346 through conductor 36d; it is also connected to terminal 348 and terminal 350 through conductor 352, and to terminal 354 through conductor 356. The output side of AND gate 344 is connected to OR gate 176 through conductor 182.

The input side of AND gate 358 is connected to on-off control 184 through conductor 188, to terminal 348 and terminal 350 through conductor 360 and through inverter 362 to terminal 354. Terminal 354 of automatic transfer circuit 34 is connected to terminal 346 of automatic transfer circuit 32. The remaining external terminals of the interconnect logic 180 of automatic transfer circuit 34 are unconnected. Terminal 348 and terminal 364 (ground connection) of the interconnect logic 180 of automatic transfer circuit 32 are connected to each other.

If automatic cabin pressure controller 16 and automatic cabin pressure controller 17 are either both on or both off, the interconnect logic 180 will operate on automatic cabin pressure controller 17 through interconnect logic 180 of automatic transfer circuit 34 to turn controller 17 on if both controllers were off or to turn controller 17 off if both controllers were on. Change in status of controller 16 is prevented by the disabling connection between terminal 348 and terminal 364 of automatic transfer circuit 32. By connecting terminal 348 to terminal 364, a false signal is input to AND gate 344 and AND gate 358 of inerconnect logic of automatic transfer circuit 32.Since each of the AND gates of interconnect logic of automatic transfer circuit 32 will constantly have a false input, they will never have a true or operative output and further discussion of their operation is unnecessary.

When automatic cabin pressure controller 17 is on, on-off control 184 issues a true signal which passes to AND gate 344 through conductor 190. If automatic cabin pressure controller 16 is simultaneously on, on-off control of automatic transfer circuit 32 will issue a true signal which will pass through terminal 346 of the interconnect logic of automatic transfer circuit 32 to terminal 354 of interconnect logic 180 of automatic transfer circuit 34.

This true signal will pass to inverter 362 and be input into AND gate 358 as a false and to AND gate 344 through conductor 356 as a true signal. The third input to AND gate 344 will be from terminal 350 or the power supply which will be a constant true input. Three true inputs at AND gate 344 will result in a true output to OR gate 176 through conductor 182 and will turn automatic cabin pressure controller 17 off as described in connection with Figure 9.

If automatic cabin pressure controller 17 is off, on-off control 184 will issue a true signal to AND gate 358 through conductor 88. If automatic cabin pressure controller 16 is off, on-off control of automatic transfer circuit 32 will be issuing a false signal to terminal 346.

This false signal will then pass to terminal 354 of interconnect logic 180 of automatic transfer circuit 34 and pass to inverter 362 which in turn will issue a true signal to AND gate 358. The false signal at terminal 354 is also passed to AND gate 344 through conductor 356 which will keep the output at AND gate 344 false. The third input to AND gate 358 is from terminal 350 which is a constant true signal through conductor 360. The three true inputs to AND gate 358 will result in a true output to OR gate 172 through conductor 220. A true signal at the input to OR gate 172 will turn automatic cabin pressure controller 17 on as described in connection with Figure 8.

The purpose of having identical interconnect logics in automatic transfer circuits 32 and 34 is to retain the complete interchangeability of the two units. Providing external terminals allows the disabling of one inerconnect logic while the other remains completely functional.

Specific embodiments of an aircraft cabin pressure control system have been shown, illustrating and describing the use of dual automatic controllers, a method for alternating their use in an aircraft, a method for achieving a linear rate of change of cabin pressure, a way of identifying the defective controller when it malfunctions, a method for determining whether an unexpected result was due to a controller malfunction, methods for preventing both controllers from being in the same state, a method for detecting undesirable rate changes in cabin pressure, and a method for preventing disasters due to the differential between cabin pressure and outside ambient pressure. It is to be understood that the foregoing embodiments are presented by way of example only and that the invention is not to be construed as being limited thereto, but only by the scope of the following claims.

WHAT WE CLAIM IS: 1. A cabin pressure control system for an aircraft, the system comprising two controllers, each of which is capable of controlling the cabin pressure, independently of the other of the two controllers, and each of which has an enabled state and a disabled state, and is arranged, when enabled, to control the cabin pressure, and the system also comprising means arranged to enable and disable the controllers in accordance with predetermined conditions, in such a way that under normal operating conditions only one controller is enabled.

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

Claims

**WARNING** start of CLMS field may overlap end of DESC **. The input side of AND gate 344 is connected to on-off control 184 through conductor 190 which is in turn connected to terminal 346 through conductor 36d; it is also connected to terminal 348 and terminal 350 through conductor 352, and to terminal 354 through conductor 356. The output side of AND gate 344 is connected to OR gate 176 through conductor 182. The input side of AND gate 358 is connected to on-off control 184 through conductor 188, to terminal 348 and terminal 350 through conductor 360 and through inverter 362 to terminal 354. Terminal 354 of automatic transfer circuit 34 is connected to terminal 346 of automatic transfer circuit 32. The remaining external terminals of the interconnect logic 180 of automatic transfer circuit 34 are unconnected. Terminal 348 and terminal 364 (ground connection) of the interconnect logic 180 of automatic transfer circuit 32 are connected to each other. If automatic cabin pressure controller 16 and automatic cabin pressure controller 17 are either both on or both off, the interconnect logic 180 will operate on automatic cabin pressure controller 17 through interconnect logic 180 of automatic transfer circuit 34 to turn controller 17 on if both controllers were off or to turn controller 17 off if both controllers were on. Change in status of controller 16 is prevented by the disabling connection between terminal 348 and terminal 364 of automatic transfer circuit 32. By connecting terminal 348 to terminal 364, a false signal is input to AND gate 344 and AND gate 358 of inerconnect logic of automatic transfer circuit 32.Since each of the AND gates of interconnect logic of automatic transfer circuit 32 will constantly have a false input, they will never have a true or operative output and further discussion of their operation is unnecessary. When automatic cabin pressure controller 17 is on, on-off control 184 issues a true signal which passes to AND gate 344 through conductor 190. If automatic cabin pressure controller 16 is simultaneously on, on-off control of automatic transfer circuit 32 will issue a true signal which will pass through terminal 346 of the interconnect logic of automatic transfer circuit 32 to terminal 354 of interconnect logic 180 of automatic transfer circuit 34. This true signal will pass to inverter 362 and be input into AND gate 358 as a false and to AND gate 344 through conductor 356 as a true signal. The third input to AND gate 344 will be from terminal 350 or the power supply which will be a constant true input. Three true inputs at AND gate 344 will result in a true output to OR gate 176 through conductor 182 and will turn automatic cabin pressure controller 17 off as described in connection with Figure 9. If automatic cabin pressure controller 17 is off, on-off control 184 will issue a true signal to AND gate 358 through conductor 88. If automatic cabin pressure controller 16 is off, on-off control of automatic transfer circuit 32 will be issuing a false signal to terminal 346. This false signal will then pass to terminal 354 of interconnect logic 180 of automatic transfer circuit 34 and pass to inverter 362 which in turn will issue a true signal to AND gate 358. The false signal at terminal 354 is also passed to AND gate 344 through conductor 356 which will keep the output at AND gate 344 false. The third input to AND gate 358 is from terminal 350 which is a constant true signal through conductor 360. The three true inputs to AND gate 358 will result in a true output to OR gate 172 through conductor 220. A true signal at the input to OR gate 172 will turn automatic cabin pressure controller 17 on as described in connection with Figure 8. The purpose of having identical interconnect logics in automatic transfer circuits 32 and 34 is to retain the complete interchangeability of the two units. Providing external terminals allows the disabling of one inerconnect logic while the other remains completely functional. Specific embodiments of an aircraft cabin pressure control system have been shown, illustrating and describing the use of dual automatic controllers, a method for alternating their use in an aircraft, a method for achieving a linear rate of change of cabin pressure, a way of identifying the defective controller when it malfunctions, a method for determining whether an unexpected result was due to a controller malfunction, methods for preventing both controllers from being in the same state, a method for detecting undesirable rate changes in cabin pressure, and a method for preventing disasters due to the differential between cabin pressure and outside ambient pressure. It is to be understood that the foregoing embodiments are presented by way of example only and that the invention is not to be construed as being limited thereto, but only by the scope of the following claims. WHAT WE CLAIM IS:

1. A cabin pressure control system for an aircraft, the system comprising two controllers, each of which is capable of controlling the cabin pressure, independently of the other of the two controllers, and each of which has an enabled state and a disabled state, and is arranged, when enabled, to control the cabin pressure, and the system also comprising means arranged to enable and disable the controllers in accordance with predetermined conditions, in such a way that under normal operating conditions only one controller is enabled.

2. A system as claimed in Claim 1, which includes means arranged to indicate which

controller is for the time being enabled.

3. A system as claimed in Claim 2, in which the indicating means comprises two separate indicator devices, one associated with each of the controllers, and arranged to indicate whether the associated controller is enabled or disabled.

4. A system as claimed in Claim 1 or Claim 2 or Claim 3, in which the enabling and disabling means includes transferring means arranged, under certain conditions, to enable whichever controller was previously disabled, and to disable whichever controller was previously enabled.

5. A system as claimed in Claim 4, which includes means arranged to monitor the performance of whichever controller is for the time being enabled, (hereinafter referred to as the primary controller), and on detection of a malfunction of that controller, to activate the transferring means to enable whichever controller was previously disabled (hereinafter referred to as the standby controller), and to disable the primary controller.

6. A system as claimed in Claim 5, in which the means monitoring the performance of whichever controller is for the time being enabled forms a part of the other controller.

7. A system as claimed in Claim 5 or Claim 6, which is so arranged that the identity of the primary and standby controllers remains unchanged on detection by the monitoring means of a malfunction of the primary controller.

8. A system as claimed in Claim 7, which includes means arranged to indicate which controller is for the time being the primary controller.

9. A system as claimed in any of Claims 5 to 8, which also includes resetting means arranged, on receipt of a resetting signal after detection by the monitoring means of a malfunction of the primary controller, to enable the primary controller and disable the standby controller.

10. A system as claimed in any of the preceding claims, in which the enabling and disabling means includes detecting means arranged to respond to the condition of both controllers being enabled by disabling one of the controllers, and to respond to the condition of both controllers being disabled by enabling one of the controllers.

11. A system as claimed in Claim 10 in which the detecting means is so arranged that the controller which is disabled in response to the first-mentioned condition is the same as the controller which is enabled in response to the second-mentioned condition.

12. A system as claimed in Claim 11, in which the detecting means includes, for each controller, a circuit which may be selectively enabled to respond to the first and second-mentioned conditions by disabling and enabling the associated controller, and the system also including means enabling one of the said circuits and disabling the other.

13. A system as claimed in Claim 12 in which the two controllers are substantially identical, and the two circuits associated with the controllers are also substantially structurally identical.

14. A system as claimed in Claim 12 or Claim 13, in which each of the said circuits comprises a first logic element arranged to receive a first input indicating that the associated controller is enabled, a second input indicating that the other controller is enabled, and a third input indicating that the said circuit is enabled, the output of the first logic element being connected to the associated controller to disable the controller when the first, second and third inputs are all present, and a second logic element arranged to receive a fourth input indicating that the associated controller is disabled, a fifth input indicating that the other controller is disabled, and a sixth input indicating that the said circuit is enabled, the output of the second logic element being connected to the associated controller to enable the controller when the fourth, fifth and sixth inputs are all present.

15. A system as claimed in Claim 14 in which the third and sixth inputs of the said circuit associated with one of the controllers are connected to a source of a signal of a constant logic condition, and the corresponding inputs of the other circuit are connected to a source of a signal of the opposite logic condition.

16. A system as claimed in any of the preceding claims, in which each controller is arranged, when enabled, to tend to maintain the cabin pressure at a set-value of pressure.

17. A system as claimed in Claim 16, in which each controller also includes means arranged to sense the ambient atmospheric pressure, and is so arranged that the said set-value of pressure is a function of the sensed ambient atmospheric pressure.

18. A system as claimed in Claim 16 or Claim 17, in which each controller also includes means arranged to sense the excess of cabin pressure over ambient atmospheric pressure, and to modify the behaviour of the controller in such a way as to minimise the said excess, in response to the said excess exceeding a predetermined value.

19. A system as claimed in any of the preceding claims, in which each controller is arranged to control cabin pressure by acting on a flow control valve.

20. A system as claimed in Claim 19, which comprises a single flow control valve arranged to be controlled by whichever controller is enabled for the time being.

21. A system as claimed in Claim 20, in which the flow control valve includes a drive motor, and switching means arranged to connect the drive motor to be driven by an output of whichever controller is enabled.

22. A system as claimed in Claim 16, or any of Claims 17 to 21 when appendant to Claim 16, which also includes means arranged to provide an output signal whenever the enabled controller is unable to maintain the cabin pressure at the set-value of pressure.

23. A system as claimed in Claim 22, when appendant to Claim 19 or Claim 20 or Claim 21, in which the flow control valve is an outflow valve, and the said means is arranged to provide the said output signal if the outflow valve reaches a closed position and the cabin pressure is nonetheless decreasing at more than a predetermined rate.

24. A system as claimed in Claim 22 or Claim 23, which also includes means arranged to indicate that the enabled controller is unable to maintain the cabin pressure at the set-value of pressure.

25. A system as claimed in Claim 22 or Claim 23 or Claim 24, when appendant to Claim 5, in which the monitoring means is disabled by the presence of the output signal from enabling the standby controller and disabling the primary controller.

26. An aircraft incorporating a cabin pressure control system as claimed in any of the preceding claims.

27. An aircraft as claimed in Claim 26, when appendant to Claim 4, in which the transferring means is arranged to change the state of the two controllers for each successive flight of the aircraft.

28. An aircraft as claimed in Claim 27, in which the transferring means includes one or more signal sources arranged to signal the end of one flight of the aircraft and/or the commencement of a subsequent flight of the aircraft.

29. An aircraft as claimed in Claim 28 in which the signal sources include a landing gear switch.

30. An aircraft as claimed in Claim 28 or Claim 29 in which the signal sources include a throttle switch.

31. An aircraft as claimed in Claims 29 and 30, in which the transferring means includes a bistable element arranged, when brought to a predetermined one of its two states, to change the state of the two controllers, the bistable element being arranged to be set in response to a signal from the landing gear switch indicating that the aircraft has landed, and to be reset in response to a signal from the throttle switch indicating a demand for increased power from the aircraft's engine or engines.

32. An aircraft as claimed in any of Claims 26 to 31, which also includes a source arranged to supply air to the cabin, and in which the cabin pressure control system includes an outflow control valve arranged to be controlled by the enabled controller to regulate cabin pressure.

33. A method for regulating cabin air pressure in an aircraft containing a pair of cabin pressure controllers, each of which is capable of controlling the cabin air pressure independently of the other of the controllers, the said method comprising the steps of: enabling one of the said pair of cabin pressure controllers; and, in response to predetermined conditions, enabling the other of the said pair of cabin pressure controllers to regulate the cabin air pressure, while disabling the said one of the pair of controllers.

34. A method as claimed in Claim 33, in which one controller of the pair is enabled during one flight, and the other controller of the pair is enabled during the following flight.

35. A method for regulating cabin air pressure in an aircraft according to Claim 34, which includes the steps of: providing a landing gear switch; and actuating the landing gear switch to enable one of the said pair of cabin pressure controllers and to disable the other of said pair of controllers.

36. A method as claimed in Claim 33 or Claim 34 or Claim 35, in which the said other controller of the pair monitors the performance of the said one controller, while the said one controller is enabled.