2. Description of the Related Art Recent changes in aircraft flight regulations for flight through the North Atlantic corridor have reduced the minimum allowable vertical distance between two aircraft flying through that airspace. Referred to as Reduced Vertical Minimum Separation, or RVSM, these regulations also require that an aircraft flying in that airspace be capable of monitoring its geo- potential height, i. e. its pressure-based height, and taking corrective action to avoid intrusion into the airspace of another aircraft where the intrusion violates the new RVSM distances. The aircraft must be capable of monitoring the altitude and airspeed of the aircraft and providing adequate warning to the pilot when the aircraft deviates from its preprogrammed flight parameters.

Aircraft typically employ electromechanical instruments to display altitude and airspeed data. Such instruments include pressure transducers for sensing air pressure, with the transducer output being converted into altitude or airspeed data, depending on the particular instrument. Due to the generally non-linear relationship between pressure and airspeed, it is necessary to linearize airspeed data derived from pressure measurements before such data can be displayed by the aircraft instruments. Pressure transducers are therefore specified for linear operation in the

pressure range for a particular aircraft. For example, transducers for commercial aircraft are typically specified to operate at pressures between 0 and 8.38 inHg ; for military aircraft, the range is between 0 and 33.6 inHg. However, the cost of the transducers increases with the operable pressure range. In addition, the resolution of the transducer is a function of the full-scale operating range of the device. Accordingly, as the operating range increases and the available output from the transducer remains constant, the resolution of the device decreases.

Transducers are also subject to drift over time which is also a function of the full-scale operating range of the device.

As a result, a transducer specified to operate over a broader pressure range will experience a greater amount of drift and will therefore require more frequent calibration.

SUMMARY OF THE INVENTION The altitude control system of the present invention is a triple-redundant system which includes a digital air data computer, pilot and copilot altimeters and airspeed indicators, and an altitude alerter. Redundancy is provided in that the computer, altimeters, and airspeed indicators are each capable of simultaneous, independent pressure measurement. Accordingly, the altitude control system of the present invention provides three independent sources of altitude and airspeed data, all connected to the aircraft pressure system.

In the inventive arrangement, the pilot, copilot and digital air data computer systems are each configured to independently measure static and Pitot pressure and display airspeed and altitude data. Although the altimeter and airspeed indicator normally operate as repeaters for the computer, they are also able to compare the data received from the computer with data sensed by their integral transducers to determine if the computer

and self-sensed data are valid. The altitude control system automatically switches from a NORMAL operating mode to a STANDBY operating mode if an error is detected in the data from any of the three pressure sensing systems. The present invention therefore provides a triplex redundant system capable of satisfying the new RVSM measurement and redundancy requirements.

The digital air data computer, airspeed indicators and altimeters each include pressure transducers connected to the aircraft pressure system for independent measurement and calculation of airspeed and altitude data. The airspeed indicator in particular includes an integral pressure transducer as part of a normalizing circuit to provide linear performance at airspeeds ranging from between approximately 0 and 400 knots (pressures between 0 and 8.38 inHG) for commercial aircraft, and from between approximately 0 and 720 knots (pressures between 0 and 33.6 inHg) for military aircraft. In contrast to the prior art, in which transducers are specified for linear performance over the entire aircraft operating range, the pressure transducer included in the airspeed indicator of the present invention is configured for linear operation over only a portion of the aircraft operating range; in particular, the transducer output will be linear at airspeeds between approximately 0 and 285 knots, corresponding to pressures between approximately 0 and 4.08 inHg. Accordingly, the transducer output will be non-linear at airspeeds above approximately 285 knots.

By subjecting the transducer to pressures exceeding its normal or specified operating limits, the present invention overstresses the transducer causing it to operate non-linearly; specifically, the transducer is overstressed to approximately 21% of its maximum specified operating pressure for commercial applications and to approximately 83k for military applications.

Due to the non-linear relationship between pressure and airspeed, the pressure change required to detect a 1 knot change

in airspeed at 25 knots is approximately 1/25 the pressure change required to detect a 1 knot change in airspeed at 500 knots.

While a 1 knot variation in airspeed at 500 knots is relatively insignificant, the same change in airspeed at 25 knots is extremely important, especially recognizing that take-off and landing occur at the lower airspeeds. It is clearly desirable for an airspeed indicator to be capable of detecting pressure changes, i. e. airspeed changes, with the same degree of accuracy over the full operating range of the aircraft. To accomplish this, a bridge circuit in the transducer is connected to a differential amplifier having a feedback loop which causes the sensitivity of the bridge to decrease as pressure increases. By coupling the bridge circuit with the differential amplifier, the performance of the transducer is substantially linearized at pressures outside its specified operating range, i. e. above 4.08 inHg, and the accuracy and sensitivity of the airspeed indicator is normalized over the entire operating range of the aircraft. This non-linear use of a transducer to measure airspeed is applicable in both commercial and military grade aircraft. It is therefore possible, in accordance with a preferred embodiment of the present invention, to employ a normalizing circuit to transform the non- linear airspeed characteristic to a linear characteristic in an airspeed indicator having approximately the same resolution over its entire operating range. This linear conversion is also necessary to condition the airspeed data for use by the digital circuitry of the airspeed indicator of the present invention.

Such a configuration also yields greater accuracy, at least equal reliability, and reduced cost over the prior art.

Other objects and features of the present invention will become apparent from the following detailed description considered in conjunction with the accompanying drawings. It is to be understood, however, that the drawings are designed solely for purposes of illustration and not as a definition of the limits of

the invention, for which reference should be made to the appended claims. It should be further understood that the drawings are not necessarily drawn to scale and that, unless otherwise indicated, they are merely intended to conceptually illustrate the structures and procedures described herein.

BRIEF DESCRIPTION OF THE DRAWINGS In the drawings, wherein like reference characters denote similar elements throughout the several views: Fig. 1 is block diagram of an altitude control system configured in accordance with the present invention; Fig. 2 is a front face view of a solid state barometric altimeter of the altitude control system of Fig. 1 ; Fig. 3 is a front face view of an airspeed indicator of the altitude control system of Fig. 1; Fig. 4 is a front face view of the airspeed indicator of Fig. 3 showing the STBY legend; Fig. 5 is a front face view of the airspeed indicator of Fig. 3 showing the angle of attack display segment; Fig. 6 is a block diagram of an airspeed indicator configured in accordance with the present invention; Fig. 7 is a front face view of an altitude alerter of the altitude control system of Fig. 1; Fig. 8 is a front face view of the altitude alerter of Fig. 7 showing the MDA and temperature legends; Fig. 9 is a graphical representation of the performance of the altitude alerter of the system of Fig. 1 in the altitude alert mode; Fig. 10 is a graphical representation of the performance of the altitude alerter of the system of Fig. 1 in the MDA mode; Fig. 11 is a block diagram of a digital air data computer of the altitude control system of Fig. 1 ;

Fig. 12 is a schematic diagram of the transducer circuit of the airspeed indicator of the altitude control system of Fig.

1; and Fig. 13 is a schematic diagram of the transducer and normalizing circuit of Fig. 12.

DETAILED DESCRIPTION OF THE PRESENTLY PREFERRED EMBODIMENTS Referring now to the drawings in detail, Fig. 1 is a block diagram of an aircraft altitude control system configured in accordance with the present invention, and generally designated at 10, including redundant pilot and copilot solid state barometric altimeters (SSBA) 100,100', redundant pilot and copilot airspeed (A/S) indicators 200,200', an altitude alerter 300, and a digital air data computer (DADC) 400. The control system 10 provides three-way redundancy for altitude and airspeed data by providing three independent pressure sensing devices, i. e. sensing devices associated with each of the SSBA 100,100', the A/S indicator 200, 200'and the DADC 400, and all connected to the aircraft Pitot- static pressure system 26. The control system 10 is a dual-mode system, operable in either a NORMAL or a STANDBY mode, with the current operating mode being selected either manually or automatically. In NORMAL mode, the SSBA 100,100'and A/S indicator 200,200'merely display altitude and airspeed data sensed by the DADC 400, operating basically as repeaters for the DADC 400. However, because each A/S indicator 200,200'and SSBA 100,100'includes internal differential pressure transducers connected identically to the aircraft Pitot-static pressure system 26, each is capable of independent pressure measurement. As a result, in NORMAL mode the A/S indicators 200,200'and SSBA 100, 100'simultaneously measure static and Pitot pressure from the same inputs provided to the DADC 400.

Each SSBA 100,100'and A/S indicator 200,200' constantly compares the data received from the DADC 400 with its

internally measured data. If an error that exceeds a predefined threshold deviation is detected between these data, the component (s) which detected the error automatically switch to STANDBY mode. For example, if the pilot's SSBA 100 and A/S indicator 200 switch to STANDBY, but the copilot's do not, then the DADC 400 and co-pilot's information"agree"and the deviation is in the pilot's instruments. The converse is likewise true; if both the pilot and copilot SSBA 100,100'and A/S indicator 200, 200'switch to STANDBY mode, then the deviation is in the DADC 400. Thus, even with one air data source removed from the altitude control system 10, two sources always remain, thereby continuing to satisfy the RVSM airspace redundancy requirement.

With continued reference to Fig. 1, the pilot and copilot instruments, i. e. the SSBA 100,100'and A/S indicator 200,200', each receive pneumatic static pressure (Ps) and Pitot pressure (Pt) input from the aircraft Pitot-static pressure system 26 on independent pressure lines 24', 28'and 24", 28", respectively. Data from the DADC 400 is provided to the respective A/S indicators 200,200'via separate data busses 40, 40'. The pilot-side SSBA 100 supplies a barometric setting value to the DADC 400 via line 46. The pilot and copilot SSBAs 100, 100'provide altitude data corrected for local pressure conditions to the altitude alerter 300 on lines 44,44'. Data is also provided by the DADC 400 to the altitude alerter 300 on data bus 40. The altitude alerter 300 supplies altitude preselect data to the DADC 400 via line 48.

Differential pressure transducers are provided within each A/S indicator 200,200'to independently measure static and Pitot pressure. While the required detection range for such pressure transducers typically extends between approximately 0 inHg and 8.38 inHg, the preferred embodiments of the present invention use transducers having an operating range of between approximately 0 inHg and 4.08 inHg. In accordance with the

invention, the transducers are operated as part of the altitude control system 10 in a manner by which they are overstressed, i. e. used outside of their specified operating range, resulting in a non-linear transducer output. This non-linearity is compensated for by incorporating each transducer in a normalizing circuit (see Fig. 13) which provides a linear output in response to the non- linear performance of the transducer. The possibility of early fatigue or premature failure of the transducers, may, in accordance with a further aspect of the preferred embodiments, be avoided by utilizing silicon transducers which, when overstressed, are not prone to early fatigue or failure.

With continued reference to Fig. 1 and additional reference to Fig. 2, the general operation of the various components of the altitude control system 10 of the present invention will now be described. For the sake of brevity, only the pilot-side instruments will be discussed below, it being understood that the discussion applies equally to the copilot instruments.

Fig. 2 is a front face view of the preferred embodiment of the SSBA 100 of the present invention. The SSBA 100 is a dual- mode integrally-lighted solid state altimeter having an electromechanically driven analog pointer 114 and a digital liquid-crystal display (LCD) 122. The SSBA 100 is readable in both direct sunlight and complete darkness, i. e. nightvision lighting is provided. The SSBA 100 operating mode is manually selectable between the NORMAL and STANDBY modes via a spring mode select switch or knob 110 in the lower right hand corner of the front bezel 112. In addition, and as described below, the SSBA 100 may switch automatically to from NORMAL to STANDBY mode if its self-sensed data does not agree with the data provided by the DADC 400. A barometric pressure setting adjust knob 116 ("baro knob") is provided on the front of the SSBA 100 to allow the pilot to manually enter a barometric pressure to correct pressure altitude

to local conditions. This value is displayed in a digital barometric setting display 118 that is provided as part of the LCD 122 and is also transmitted to the DADC 400 on lead 46. If the SSBA 100 detects a failure in the circuit associated with the baro knob 116, it displays and uses a fixed value of 29.92 inHg for calculation of corrected barometric altitude; in the event of a digital failure the digital barometric setting display 118 will not be affected by the barometric setting adjust knob 116. In addition, a barometric failure graphic symbol 128 is illuminated to the right of the barometric display 118 and the barometric display 118 flashes until deactivated by the pilot's toggling of the mode select knob 110 either right or left. As previously described, each SSBA 100,100'is connected to the DADC 400 via a separate data bus 40,40', and independent power busses (not shown) provide additional redundancy.

Errors in measuring static pressure may arise partly due to imperfections in the exterior surface of the aircraft or imperfections at the input opening of the static pressure source.

To compensate for such imperfections, static source error correction (SSEC) calibration data is obtained by measuring Pitot and static pressures under various speed, altitude and gross load conditions, for particular aircraft. This measured calibration data is stored in the SSBA 100 as SSEC values which are compared with measured pressure data (either from the DADC 400 or internally measured data) to derive SSEC pressure altitude.

The SSBA 100 includes a differential pressure transducer for detecting the differential pressure, i. e. the difference between Pitot and static pressures. The transducer receives pneumatic input from the aircraft Pitot-static pressure system 26 via line 24', from which it calculates static source error (SSE) corrected altitude. The SSBA 100 also includes a preprogrammed threshold deviation value which is used to determine when the self-sensed data and DADC 400 data differ to a degree sufficient

to trigger an automatic switchover to STANDBY operation. In use, the SSBA 100 constantly compares the altitude data that it receives from the DADC 400 with its self-sensed data; if the difference between these data exceed a predetermined threshold deviation, then the SSBA 100 automatically switches to STANDBY operation. In a preferred embodiment, the SSBA 100 will automatically switch from NORMAL to STANDBY operation when the difference between its self-sensed altitude and the altitude data received from the DADC 400 exceeds a nominal value of 150 feet.

After switching to STANDBY, a STBY legend 126 will flash.

Reversion to NORMAL operation will then not occur until the difference between the self-sensed data and DADC 400 data is less than about 125 feet.

When operating in the NORMAL mode, SSBA 100 corrects the input altitude data received from the DADC 400 to match the setting of the baro knob 116 and displays SSE corrected baro altitude via both the LCD 122 and analog pointer 114. The SSBA 100 also communicates altitude data, corrected to the baro knob 116 setting, to the altitude alerter 300 on lead 44.

When operating in the self-sensing STANDBY mode, SSBA 100 calculates SSE corrected altitude from static pressure sensed by its internal pressure transducer and displays this value via the LCD 122.

Calibration of the pressure transducer is automatically checked by the SSBA 100. In addition, the SSBA 100 also automatically validates the signal returned from the pressure transducer against a known signal range. If, when operating in STANDBY mode, the SSBA 100 detects an error in either of these signals or data, it will automatically display NORMAL mode data (if it is available from the DADC 400), and will also indicate the detection of suspect or inaccurate data by displaying"Err"on LCD 122.

Each SSBA 100 provides annunciation of excessive HOLD altitude deviation, which typically results from autopilot drift.

The SSBA 100 receives as input the state of the HOLD altitude value sent to the autopilot and continuously monitors the difference between the actual 29.92 inHg altitude and the HOLD altitude. If the difference between these values deviates beyond a predetermined limit, then a STBY legend 126 (see Fig. 2) flashes until either the HOLD altitude value is released, i. e. autopilot is disabled, or the actual 29.92 inHg altitude returns to an acceptable range.

Referring next to Fig. 3, a front face view of the preferred embodiment of the A/S indicator 200 of the present invention is there shown. The A/S indicator 200 is configured as a dual-mode integrally-lighted indicator having a non-metallic body 270 to which a metallic bezel 212 is affixed. An analog pointer 214 is provided as part of a display 224 that also includes an LCD section 222. LCD section 222 further includes an upper three-digit digital segment 226 for displaying either computed airspeed (CAS) during NORMAL mode operation or indicated airspeed (IAS) during STANDBY mode operation. A lower three-digit digital segment 228 displays either true airspeed (TAS) or ground speed (GS), user selection of which is made by toggling the control lever 210 clockwise, i. e. to the"NORM"position. The pointer 214 and upper digital segment 226 are synchronized when either CAS or IAS is being displayed. Failure of any part of LCD 222 will not affect the operation of analog pointer 214. A discrete legend 218 adjacent the lower digital segment 228 indicates whether TAS or GS is being displayed (see Fig. 4). The A/S indicator 200 display defaults to NORMAL mode, displaying TAS.

The control lever 210 is movable in the clockwise and counterclockwise directions to manually set the operating mode of the A/S indicator 200. When the lever 210 is moved in the clockwise direction to the"NORM"position, the A/S indicator 200

will either switch from STANDBY to NORMAL mode (if the A/S indicator 200 was previously operating in STANDBY mode), or such movement will toggle the lower digital segment 228 between TAS and GS.

The control lever 210 may also be manually moved in a counterclockwise direction to the"STBY"position to operate the A/S indicator 200 in STANDBY mode and cause a STBY flag 240 to flash. When operating in this mode, the A/S indicator 200 employs internal pressure sensing to compute IAS from data received directly from the aircraft Pitot-static pressure system 26.

A movable command bug 262 is provided on the front face of the A/S indicator 200 for selective movement about its periphery. The bug 262 is movable by selective manual rotation of a bug setting knob 264.

When operating in its NORMAL mode, the A/S indicator 200 displays CAS, maximum operating velocity (Vmo), and TAS or GS from data supplied by the DADC 400. CAS is displayed on the upper digital segment 226 and by the analog pointer 214 against a non- linear dial scale. The lower digital segment 228 shows either TAS or GS. By switching control lever 210 to the"NORM"position, the lower digital segment 228 will alternate between a display of TAS and GS. The display of TAS is accompanied by a discrete TAS indicating legend 218. Transition from the display of TAS to GS will not, however, occur if GS data is invalid or not available; if TAS data is invalid or unavailable, the lower digital segment 228 will show"---"in TAS display mode. In NORMAL mode, Vmo is presented as a segmented arc 220 along a scale provided as part of the LCD 222 representing a"not to exceed"velocity. The arc extends between 175 and 340 knots and is composed in the preferred embodiment of 41 equally sized liquid-crystal elements. At least one Vmo segment is active when Vmo is valid. If Vmo data is not available or if the A/S indicator 200 is unable to process the

available Vmo data, the segmented arc 220 displays a repeating pattern of three segments on, three segments off.

In STANDBY mode, the A/S indicator 200 employs the integral pressure transducer to determine and display IAS. The TAS/GS lower digital segment 228 and Vmo segmented arc 220 are blank when operating in STANDBY mode (See Fig. 4). Each A/S indicator 200,200'is connected to the DADC 400 via a separate data bus 40,40', and independent power busses (not shown) provide additional redundancy.

Automatic conversion from NORMAL to STANDBY mode occurs if the A/S indicator 200 does not receive CAS data from the DADC 400 over digital bus 40 or if such data is inoperative or determined to be invalid, as indicated by specific bits in the DADC 400 data.

The A/S indicator 200 of the present invention includes an integral pressure transducer to facilitate its operation in STANDBY mode. The transducer is operable over an airspeed range of between approximately 0 and 400 knots for commercial grade aircraft and between approximately 0 and 720 knots for military grade aircraft. Calibration of the pressure transducer is automatically checked by the A/S indicator 200. The A/S indicator 200 also automatically checks or validates the signal returned from the pressure transducer against a known signal range and checks the pressure transducer temperature sensor against a known signal range. If, when operating in STANDBY mode, the A/S indicator 200 detects a calibration data fault or that the returned pressure or temperature signal is out of the predefined range, indicator 200 will automatically display NORMAL mode CAS data (if it is available from the DADC 400) and an"Err"legend in the upper digital display 228 to indicate the error condition.

Referring next to Figs. 12 and 13, the transducer 270 that is integral or associated with the A/S indicator 200 of the present invention forms a part of a normalizing circuit 282. The

transducer 270 is preferably a differential pressure transducer machined from silicon, such for example, as EG&G IC Sensors model number 1210A-002. The transducer 270 is preferably normally operable at pressures between approximately 0 inHg and 4.08 inHg, its normal, intended (by its manufacturer) operating range in which its performance or output is substantially linear with changes in pressure. However, and in accordance with the present invention, the transducer 270 is operatively subjected to pressures between approximately 0 inHg and 33.6 inHg. As a result, the pressure transducer 270 is operated beyond its specified limits, i. e. overstressed, and its performance at pressures above 4.08 inHg (or whatever the intended operating range) is consequently non-linear.

An important object and feature of the invention is that the A/S indicator 200 be capable of detecting pressure changes with the same degree of accuracy over the full operating range of the aircraft. To accomplish this, the transducer 270 of the A/S indicator 200 of the present invention includes a bridge circuit 298 that forms a part of the normalizing circuit 282. The outputs 288,290 (Fig. 13) of the bridge circuit 298 are connected to the non-inverting input 304 and inverting input 306, respectively, of a differential signal amplifier 272. A feedback loop 280 having a feedback resistor 312 is connected between the output 276 of the signal amplifier 272 and the inverting input 292 of a drive amplifier 284, the non-inverting input 300 of which is connected to a voltage source 274. In a preferred embodiment, the voltage source 274 provides a reference voltage of approximately 2.5 volts DC at input 300. The output 302 of drive amplifier 284 is connected to the bridge circuit 298 at node 286 and operates as a voltage source thereto.

The DC gain of drive amplifier 284 is very large and theoretically infinite. As a result, there exists a virtual short circuit between its inverting and non-inverting input terminals

292,300 so that the voltage at each terminal is approximately 2.5 volts DC. Output node 310 and feedback loop 280 are therefore summed at node 308 to drive the voltage at the inverting input 292 to 2.5 volts DC; increases in pressure, producing increases in the output 276 of signal amplifier 272, accordingly correspondingly reduce the amount of voltage necessary from node 310 to maintain 2.5 volts DC at non-inverting input 292 and also reduce the voltage output from drive amplifier 284. With a smaller voltage input to bridge circuit 298 at node 286, the relatively large changes in pressure that accompany relatively small changes in airspeed during non-linear operation of the transducer will produce smaller differences in the voltages at nodes 288 and 290.

The sensitivity of the bridge circuit 298 therefore decreases as pressure increases, countering the increasingly non-linear relationship between pressure and airspeed at higher pressures.

The signal amplifier 272 responds to changes in pressure by detecting the voltage difference between nodes 288 and 290; the voltage at these nodes is proportional to the transducer 270 resistance which is, in turn, proportional to pressure. As the difference between the voltage at nodes 288 and 290 increases, indicating increasing pressure, the output 276 of differential amplifier 272 also increases as does the voltage across feedback resistor 312 and accordingly, the feedback voltage component at node 308.

When coupled with the non-linear output of the transducer at pressures above 4.08 inHg, the signal amplifier 272 of the present invention produces an output that approaches linearity and that provides a substantially consistent response to changes in pressure over the entire operating range of the aircraft. This configuration also yields greater accuracy, at least equal reliability, and reduced cost over prior art arrangements and, by using a silicon transducer, no adverse effects such as fatigue, early failure, etc., result from

overstressing of the transducer through operation outside of its normal operating range.

Although the transducer 270 of A/S indicator 200 of the present invention is specified by the manufacturer as being intended for operation at pressures between approximately 0 and 4.08 inHg, it is, in a preferred embodiment of the present invention, operable over a pressure range of between approximately 0 and 33.6 inHg. The specified operating range is accordingly appreciably narrower than the aircraft operating range and, as a result, appreciably narrower than that of typical pressure transducers used in such applications, i. e. transducers are usually selected for their ability to operate linearly over the entire operating range of the aircraft. The advantages of using such a transducer in accordance with the present invention include increased resolution. In addition, the effects of instrument drift, which are likewise related to the full-scale operating range of the transducer, are reduced in accordance with the invention due to the reduced specified operating range of the transducer 270 of the present invention. Consequently, calibration of the transducer 270 is required less often than with transducers employed in prior art devices. Finally, the cost of the transducer increases with its operable pressure range so that the present invention utilizes less costly transducers.

In an alternative embodiment, the A/S indicator 200 of the present invention can provide windshear alert data to the pilot via visual and audible warnings. The A/S indicator 200 may use either self-sensed or externally-provided inertia longitudinal speed and IAS data to warn of an impending windshear condition during take-off and approach. When such conditions exist, a voice warning of"Windshear"accompanies a cyan"WS" legend (not shown) displayed on the LCD 222 of A/S indicator 200.

A/S indicator 200 may also be used during take-off to ensure that the required take-off airspeed is achieved for the

available length of runway. If the indicator 200 detects that airspeed is below the minimum required for take-off, an audible alarm warns the pilot of an impending"refuse take-off"condition.

Referring now to Fig. 5, when TAS drops below 250 knots the A/S indicator 200 will automatically reconfigure to display Angle of Attack (AoA) by using the segmented arc 220 of LCD 222 adjacent the 250 to 450 knot markers. A scale of 0 to 1 displays AoA as a percent of stall. If AoA information is not supplied by the DADC 400, the AoA segment 220 is blank (see Fig. 3) since the A/S indicator 200 cannot self-sense AoA data. When a sensed stall condition is impending, a voice warning will be emitted to alert the pilot to such condition.

In a preferred embodiment, the display 224 of the A/S indicator 200 is configured so that a pilot can readily detect, read and understand the data when displayed under any lighting conditions ranging from full-sunshine to complete darkness.

Fig. 7 depicts an altitude alerter 300 of the present invention and which includes a non-metallic body 330 having a front bezel 312 and an LCD 320 having upper and lower digital display segments 322,324, respectively. The altitude alerter 300 provides a single aural advisory configuration mode and two user- selectable operating modes: altitude alerting, to both audibly and visually annunciate conditions of approach to, and departure from, a user selected altitude, and minimum descent altitude alerting (MDA) to indicate approach to a user-selected minimum altitude.

The altitude alerter 300 also displays cabin altitude as determined by an integral pressure transducer, or SAT as received from the DADC 400 on data bus 40. The pressure transducer is capable of self-sensing the ambient air pressure over a range of between approximately 0 and 50,000 feet. Alerter 300 also receives as an input the SSEC corrected barometric altitude from the pilot SSBA 100, and includes a bidirectional serial interface (not shown) over which messages may be transmitted and received

including, without limitation, active mode (i. e. MDA or altitude alert), actual baro corrected altitude (as a DADC 400 repeater), selected target altitude, SAT and cabin altitude.

The current operating mode is selected by depressing an illuminated pushbutton 310. If no alerts are active on the altitude alerter 300, then pushbutton 310 may be depressed for less than or about one second to toggle the alerter 300 between its operating modes. If the pushbutton 310 is depressed for more than one second, then the altitude alerter 300 toggles the display mode from cabin altitude to static air temperature (SAT); when displaying SAT, the altitude alerter 300 will revert back to cabin altitude display after five seconds. If pushbutton 310 is depressed for more than 8.75 seconds, then altitude alerter 300 assumes a failed pushbutton 310 condition and reverts operation of the alerter 300 to altitude alert mode. When the altitude, MDA, or cabin pressure alerts are displayed, depressing pushbutton 310 will extinguish the displays.

A rotary knob 314 on the front of the altitude alerter 300 is operable to increase or decrease the displayed altitude value for altitude alert or MDA functions; knob 314 may be used to adjust the altitude alert range between approximately 0 and 50,000 feet, and to adjust the MDA range to between approximately 0 and 15,000 feet. If altitude and MDA alert advisories are active, then rotation of knob 314 will extinguish the advisories until the preset approach conditions are subsequently met.

LCD 320 includes a five-digit upper segment 322 that provides a direct reading of the selected altitude in either the altitude alert or MDA modes. In its lower segment 324, it displays a digital indication of cabin altitude to 500 foot resolution at altitudes between 0 and 15,000 feet, and a resolution of 1000 feet at altitudes between 15,000 and 50,000 feet. A discrete legend (not shown) accompanies the cabin altitude display. Likewise, when SAT is displayed in the lower

segment 324, it is accompanied by a discrete"°C"legend 340, as seen in Fig. 8. Failure of any one display area of the altitude alerter 300 will not affect the performance or operation of the remaining display areas.

When configured for operation in the aural advisory mode, altitude alerter 300 of the present invention provides a mechanism through which an aural warning interface can be configured for selection of a tone or voice advisory. The configuration mode is entered during power-up of the instrument by applying power while pushbutton 310 is depressed. After completing a display test, alerter 300 sets the aural advisory configuration mode active if pushbutton 310 remains depressed at the conclusion of the display test. The type of advisory, i. e. tone or voice, is selectable by rotating the rotary knob 314; turning knob 314 clockwise selects voice while counterclockwise rotation selects a tone. By then depressing pushbutton 310 once again, the configuration mode transitions from advisory selection to volume control. If the pushbutton 310 is not depressed again following power-up, alerter 300 exits the configuration mode after approximately nine seconds of inactivity of rotary knob 314.

During volume adjustment mode, clockwise rotation of rotary knob 314 increases the volume, and counterclockwise rotation decreases the volume. If the alerter 300 detects a minimum of approximately four seconds of inactivity of knob 314 following the volume adjustment mode, it automatically enters altitude alert or MDA mode. The altitude alerter 300 retains advisory type and volume adjustment settings in the absence of power to the unit.

In altitude alert mode, altitude alerter 300 of the present invention provides visual and audio indications of certain altitude situations in relation to an operator-selected target altitude. Cabin altitude is constantly displayed in the lower segment 324, as illustrated in Fig. 7. When configured for altitude alert, alerter 300 monitors approach, either ascending or

descending, to a selected altitude and departure and/or deviation from the preselected alert altitude. The upper limit of the alert altitude is 1,000 feet, +/-25 feet, above the alert altitude; when the aircraft passes through this altitude, alerter 300 sounds an audio alert and illuminates pushbutton 310, and when the aircraft approaches to within 200 feet, +/-25 feet, of the alert altitude, the pushbutton light is extinguished. A profile of the operation of the alerter 300 configured in altitude alert mode is shown in Fig. 9. The preselected target is retained by altitude alerter 300 in the absence of power.

In MDA mode, altitude alerter 300 of the present invention provides visual and audio indications of certain altitude situations relative to an operator-selected MDA target.

Cabin altitude is constantly displayed in the lower segment 324, as illustrated in Fig. 7, and in MDA legend 328 is displayed in the lower segment 324 next to the cabin altitude data. Three distinct operational situations may exist for MDA mode: 1) descent to within 200 feet of the selected MDA altitude; 2) descent to within 100 feet of the selected MDA altitude; and 3) descent below the selected MDA altitude. For the first condition, the alerter 300 provides voice messages at 100 feet above target level if the aural advisory is set to voice mode; otherwise, a warning tone sounds. When the aircraft reaches 200 feet above the selected MDA altitude, pushbutton 310 illuminates; upon subsequent descent to within 100 feet of the selected MDA altitude, the voice message or audio tone is emitted and the illuminated pushbutton 310 is extinguished. As the aircraft passes through the selected MDA altitude, altitude alerter 300 sounds a voice message or an audio tone and flashes the light of pushbutton 310 three times at a rate of approximately 1/3 second on and 1/3 second off. If the aircraft then climbs above the level set in the second condition, i. e. 100 feet above MDA altitude, pushbutton 310 re-illuminates and the voice message or audio tone sounds; ascending through the

upper limit extinguishes the light of pushbutton 310. For the first and second conditions indicated above, the altitude alerter 300 of the invention is subject to hysteresis of between approximately +/-20 feet. A profile of the operation of the alerter 300 in MDA mode is shown in Fig. 10. The preselected target altitude is retained in the absence of power.

Altitude alerter 300 displays cabin altitude in the lower segment 328 of LCD 320 in both altitude alert and MDA modes.

An internal pressure transducer derives cabin altitude to a resolution of approximately 500 feet at altitudes below 15,000 feet, and to approximately 1,000 feet resolution at altitudes above 15,000 feet. The alerter 300 issues a visual warning in the form of a flashing cabin altitude display when the self-sensed cabin altitude is at or above 10,000 feet; the flashing ceases when the sensed cabin altitude returns to or below 9,500 feet.

Altitude alerter 300 may also have its target altitude and/or operating mode remotely set, in which use a remote legend 326 (Fig. 7) is displayed in the lower segment 324 of LCD 320.

Manual activation of rotary knob 314 extinguishes remote legend 326.

Shown in Fig. 11 is a block diagram of the preferred embodiment of the DADC 400 of the present invention. DADC 400 receives temperature data from the aircraft temperature probe 22 via line 20 at a rate of not less than approximately twice per second. DADC 400 also receives a barometric setting from the pilot SSBA 100 and provides a barometric correction accurate to within 10 feet over a valid correction range of between approximately 28.1 and 31.00 inHg.

Static pressure is input to DADC 400 through line 24 from the aircraft Pitot-static system 26 and is sensed by a pressure transducer integral to DADC 400. DADC 400 reads the transducer value at a rate of not less than 16 times per second.

Total or Pitot pressure is input through line 28 to DADC 400 from

the aircraft Pitot-static pressure system 26 and is detected by a separate integral transducer; DADC 400 also reads this transducer value at a rate of not less than 16 times per second.

Temperature probe 22 is operable for measuring temperature over a range of between approximately-100 and +350° C, with DADC 400 accepting as an input a temperature in the range of approximately-100 to +50° C. The pilot SSBA 100 provides a barometric pressure setting through line 46 to DADC 400 in the form of a variable DC voltage which corresponds to a barometric pressure range of approximately 28.01 inHg to 31.00 inHg (approximately-1727 ft. to +983 ft). Digital altitude preselect data is supplied through line 48 by the altitude alerter 300 to DADC 400 on an RS-232 serial bus (not shown) at a rate of approximately 9,600 bits per second (bps) and is updated approximately every 1/8 second.

With continued reference to Fig. 11, the data outputs from DADC 400 which are of interest to the present invention will now be discussed in detail. Digital serial data is supplied by DADC 400 to the pilot and copilot SSBA 100,100'and A/S indicator 200,200'on separate identical serial busses 40,40'at a rate of 12,500 bps and are updated every 1/32 of a second. The data carried on the busses 40,40'and sent to the A/S indicator 200, 200'includes GS, pressure altitude, baro corrected altitude, IAS, Vmo, TAS, altitude rate and SAT. Only pressure altitude data is sent to the pilot SSBA 100,100', while baro corrected altitude and SAT are provided to the altitude alerter 300 over bus 40.

When operating in its NORMAL mode, the DADC 400 of the altitude control system senses static pressure Ps, and Pitot pressure Pt, which are used to compute indicated airspeed IAS (Pt minus Ps) and pressure altitude. Errors in the measurement of static pressure may arise partly due to imperfection of the static source location. To compensate for such errors, DADC 400 includes preprogrammed SSEC data referred which represents measured values

for defects at the static pressure source for the particular aircraft. This error data is typically developed by flying the aircraft under a variety of conditions, e. g. altitude, speed, load, etc., and comparing the measured pilot and static pressures to known values. DADC 400 uses the preprogrammed SSEC data to compute corrected pressure altitude and CAS. Also stored in firmware in DADC 400 are Vu5 curés based on data measured for the aircraft under various flight conditions and configurations; such Vmo curves are well known in the art. The various SSEC and Vmo curves are by way of example selectable by pin-jumpers in the A/S indicator 200 mating connector (not shown). In a preferred embodiment, four jumper positions are provided for a total of eight different SSEC/Vm curves. These SSEC stored data are compared with measured pressure data to derive SSEC pressure altitude and CAS, where CAS refers to the theoretically correct indicated airspeed that has been properly adjusted to compensate for static source errors.

DADC 400 receives barometric pressure setting as input from the pilot side SSBA 100 and computes baro corrected altitude therefrom. DADC 400 then transmits SAT and baro corrected altitude to altitude alerter 300 and, on data busses 40.40', SSEC pressure altitude to the SSBA 100,100', and CAS, true airspeed (TAS), GS, AoA, mach and Vmo to the A/S indicators 200,200'.

If DADC 400 fails or if the difference between the DADC 400 measured air data and the A/S indicators 200,200'and SSBA 100,100'measured data exceeds a preset threshold, then the altitude control system automatically switches from NORMAL to STANDBY operating mode.

While there have shown and described and pointed out fundamental novel features of the invention as applied to preferred embodiments thereof, it will be understood that various omissions and substitutions and changes in the form and details of the devices illustrated, and in their operation, may be made by

those skilled in the art without departing from the spirit of the invention. For example, it is expressly intended that all combinations of those elements and/or method steps which perform substantially the same function in substantially the same way to achieve the same results are within the scope of the invention.

It is the intention, therefore, to be limited only as indicated by the scope of the claims appended hereto.