Appendix A A-13 Figure A-1. Example of Air Data System/Autopilot Configuration A .6.2.4.1 The air data sensor is a single LRU activated in altitude hold mode when the pilot presses an ALT HOLD button, after reaching the desired cruise FL. It is not tied to either ADC or other components comprising the ADS.
The air data sensor provides ∆H information to the autopilot so the airplane can maintain the desired altitude. In some configurations, the pilot further provides FL information to the autopilot by manually selecting the displayed altitude (either pilot’s or copilot’s).
A.6.2.4.2 Airplanes equipped with the avionics configuration shown in Figure A-1 may show compliance as follows:
1. The airplane must maintain altitude to within ±65 ft of the acquired altitude as required by item 1 under paragraph A.4.1.4. For RVSM compliance, the ∆H signal must be accurate enough such that the airplane maintains the required ±65 ft altitude deviation specification. This may be substantiated with flight test data and/or manufacturer’s specification data.
2. T he altitude alerter should function if the air data sensor /ADC fails. If the altitude alert function is not operational, altitude hold performance must be monitored manually.
3. The air data sensor should be compensated such that an airspeed change at a cruise FL is not interpreted by the system as change in altitude, causing altitude hold deviation in excess of ±65 ft.
4. The altimetry systems meet the RVSM accuracy requirements specified in paragraphs A.5.2.1 and A.5.2.2, or paragraph A.5.3.2, as appropriate.
5. All other requirements set forth in this AC, as appropriate, are satisfied.
Appendix A A-14 A.6.2.5 Single ADC input to the autopilot: On a large number of older airplane models, the avionics configuration is such that one ADC provides altitude hold information to the autopilot (see Figure A-2). In most models, a second ADC is also present, or provisions exist so a second can be installed.
Figure A-2. Single Air-Data Computer Configuration for Autopilot Input A .6.2.5.1 Airplanes equipped with the avionics configuration shown in Figure A-2 may show compliance as follows:
1. The airplane must maintain altitude to within ±65 ft of the acquired altitude required by item 1 under paragraph A.4.1.4.
This may be substantiated with flight test data or manufacturer’s specification data.
2. T he altitude alerter should function i f either ADS or ADC fails.
If the altitude alert function is not operational, altitude hold performance must be monitored manually.
3. If ADC 1 fails, the airplane must be controlled manually until air traffic control (ATC) contingency procedures are executed.
Annunciation should be provided if the pilot deviates ±300 ft from desired altitude. This annunciation must be provided automatically by the altitude alert system. If the altitude alert system is not functioning, altitude hold performance must be monitored manually.
4. The altimetry systems meet the RVSM accuracy requirements specified in paragraphs A.5.2.1 and A.5.2.2, or paragraph A.5.3.2, as appropriate.
5. All other requirements in this AC, as appropriate, are satisfied.
A.6.3 O perational Restrictions . Applicants should be aware operational restrictions and/or changes may also be required for aircraft with avionics configurations shown in Figures A-1 and A-2 , to meet all RVSM requirements.
Appendix A A-15 A.7 Altimetry System Performance Substantiation.
A.7.1 Flight Testing: Group and Non-Gro up Aircraft.
A.7.1.1 Where precision flight calibrations are used to quantify or verify altimetry system performance, they may be accomplished by any of the following methods. Flight calibrations should only be performed once appropriate ground checks have been completed, and the certifying authority should agree to the number of flight test conditions. Uncertainties in application of the method must be assessed and taken into account in the data package.
1. Precision tracking radar in conjunction with pressure calibration of the atmosphere at test altitude.
2. Trailing cone.
3. Pacer aircraft. The pacer aircraft must have been directly calibrated to a known standard. It is not acceptable to calibrate a pacer aircraft by another pacer aircraft.
4. Any other method acceptable to the FAA or approving authority.
Note: Data acquired using elements from the RVSM monitoring program, such as a ground-based height monitoring unit (HMU) or Aircraft Geometric Height Measurement Element (AGHME), or a portable Global Positioning System (GPS)-based monitoring unit (GMU), is not acceptable for substantiating the ASE performance specified in paragraphs A.5.2 and A.5.3.
A.7.1.2 ASE will generally vary with flight condition. The data package should provide coverage of the RVSM envelope sufficient to define the largest errors in the basic and full RVSM envelopes. Note that, in the case of Group approval, the worst flight condition may be different for each of the requirements of paragraphs A.5.2.1 and A.5.2.2, and each should be evaluated. Similarly, for Non-Group approval, the worst flight condition may be different for each of the requirements of paragraph A.5.3.2 and each should be evaluated.
A.7.2 ASE Variability. In order to evaluate a system against the ASE performance statements established by the Review of the General Concept of Separation Panel (RGCSP) (see Appendix D), it is necessary to quan tify the mean and three SD values for ASE, expressed as ASEmean and ASE3 SD. In order to do this, it is necessary to account for the different ways in which variations in ASE can arise. The factors affecting ASE are as follows and should be considered in the ASE evaluation:
1. Unit-to-un it variability of avionics.
2. Effect of environmental operating conditions on avionics.
3. Effect of transducer and/or avionics component error drift over time.
Appendix A A-16
4. Effect of flight operating condition on SSE.
5. Airframe-to-airframe variability of SSE, including the following:
• Skin waviness, skin splices/joints, access panels, and radome fit/fair.
• Pitot-static probe variation. This includes manufacturing variation, installation variation, and probe degradation (erosion/corrosion) over time.
• Static port variation (for aircraft configured with static sources flush to the skin surface). Sources of variation include port step-height, degradation, and static port condition.
• SmartProbes© (integrated ADC/pitot-static probe). Smart probes are sensitive to installation variation. They are also capable of complex SSEC algorithms that are a function of several variables, all of which may be affected by probe condition and installation.
A.7.2.1 Assessment of ASE, whether based on measured or predicted data, must include the factors listed above in items 1 through 5. The effect of item 4 as a variable can be eliminated by evaluating ASE at the most adverse flight condition in an RVSM flight envelope.
A.7.2.2 This document does not specify separate limits for the various error sources contributing to the mean and variable components of ASE as long as the overall ASE accuracy requirements of paragraph A.5.2 or A.5.3 are met. For example, in the case of Group approval, the smaller the mean of the Group and the more stringent the avionics standard, the larger the available allowance for SSE variations. In all cases, present the tradeoff adopted in the data package in the form of an error budget including all significant error sources.
A.8 Altimetry System Component Error Budget.
A .8.1 General. The ASE budget demonstrates the allocation of tolerances among the various parts of the altimetry system is, for the particular data package, consistent with the overall statistical ASE requirements. These individual tolerances within the ASE budget represent the maximum error levels for each of the ADS components contributing to ASE. These error levels form the basis of the maintenance procedures used to substantiate the RVSM airworthiness compliance status of Group or Non-Group aircraft.
The component error evaluation should be assessed at the worst flight condition in the basic and full envelope.
A.8.2 ASE Components.
A.8.2.1 General. Figure A-3, Altimetry System Error and Its Components, shows the breakdown of total ASE into its main components, with each error block representing the error associated with one of the functions needed to generate a display of pressure altitude. This breakdown encompasses all ASEs that can
Appendix A A-17 occur, although different system architectures may combine the components in slightly different ways.
A.8.2.1.1 The “Actual Pressure Altitude” is the pressure altitude corresponding to the undisturbed ambient pressure.
A.8.2.1.2 “Static Source Error” is the difference between the undisturbed ambient pressure and the pressure within the static port at the input end of the static pressure line.
A.8.2.1.3 “Static Line Error” is any difference in pressure along the length of the line.
A.8.2.1.4 “Pressure Measurement & Conversion” is the error associated with the processes of transducing the pneumatic input seen by the avionics and converting the resulting pressure signal into altitude. As drawn, Figure A-3 represents a self-sensing altimeter system in which the pressure measurement and altitude conversion functions would not normally be separable. In an ADC system, the two functions would be separate and SSEC would probably then be applied before pressure altitude (Hp) was calculated.
A.8.2.1.5 “Perfect SSEC” would be that correction which compensated exactly for the S SE actually present at any time. If such a correction could be applied, then the resulting value of Hp calculated by the system would differ from the actual altitude only by the static line error plus the pressure measurement and conversion error. In general, this cannot be achieved, so although the “Actual SSEC” can be expected to reduce the effect of SSE, it will do so imperfectly.
A.8.2.1.6 “Residual Static Source Error” is applicable only in systems applying an avionic SSEC. It is the difference between the SSE and the correction actually applied. The corrected value of Hp will therefore differ from actual pressure altitude by the sum of static line error, pressure measurement and conversion error, and residual SSE.
A.8.2.1.7 The baro- correction error and the display error occur between Hp and the displayed altitude. Figure A-3 represents their sequence for a self-sensing altimeter system. ADC systems can implement baro-correction in a number of ways that would slightly modify this part of the block diagram, but the errors would still be associated with either the baro-correction function or the display function. The only exception is those systems that can be switched to operate the display directly from the Hp signal. These systems can eliminate baro-correction error where standard ground pressure setting is used, as in RVSM operations.
Appendix A A-18 Figure A-3. Altimetry System Error and Its Components A.8.2.2 SSE Components. The component parts of SSE are presented in Table A-2, Static Source Error, with the factors controlling their magnitude.
A.8.2.2.1 Th e reference SSE is the best estimate of actual SSE, for a single aircraft or an aircraft Group, obtained from flight calibration measurements. It is variable with operating condition, characteristically reducing to a family of W/δ curves that are functions of Mach. It includes the effect of any aerodynamic compensation incorporated in the design, and once it has been determined, the reference SSE is fixed for the single aircraft or Group, although it may be revised if substantiated with subsequent data.
A.8.2.2.2 The test techniques used to derive the reference SSE will have some measurement uncertainty associated with them, even though known instrumentation errors will normally be eliminated from the data. For trailing-cone measurements, the uncertainty arises from limitations on pressure measurement accuracy, calibration of the trailing-cone installation,
Appendix A A-19 and variability in installations where more than one is used. Once the reference SSE has been determined, the actual measurement error is fixed, but as it is unknown, it can only be handled within the ASE budget as an estimated uncertainty.
A.8.2.2.3 The airframe variability and pitot-static probe/static port variability components arise from differences between the individual airframe and pitot-static probe/static port, and the example(s) of airframe and probe/port used to derive the reference SSE.
A.8.2.3 Residual SSE.
A.8.2.3.1 Figure A-3 p resents the components and factors. Residual SSE consists of those error components that make actual SSE different from the reference value (components 2), 3), and 4) from Table A-2), plus the amount by which the actual SSEC differs from the value that would correct the reference value exactly (components 2)a), 2)b), and 2)c) from Table A-3, Residual Static Source Error (Aircraft with Avionic Static Source Error Correction)).
A.8.2.3.2 T here will generally be a difference between the SSEC that would exactly compensate the reference SSE, and the SSEC that the avionics is designed to apply. This arises from practical avionics design limitations. The resulting Table A-3 error component 2)a) will therefore be fixed, for a particular flight condition, for the single aircraft or Group. Additional variable errors 2)b) and 2)c) arise from those factors causing a particular set of avionics to apply an actual SSEC that differs from its design value.
A.8.2.3.3 The relationship between perfect SSEC, reference SSEC, design SSEC, and actual SSEC is illustrated in Figure A-4, Static S ource Error/Static Source Error Correction Relationships for Altimetry System Error Where Static Line, Pressure Measurement, and Conversion Errors Are Zero, for the case where static line errors and pressure measurements and conversion errors are taken as zero.
A.8.2.3.4 A ccount for factors creating variability of SSE relative to the reference characteristic in two ways: first, as noted for the SSE itself in Table A-2, and second, for its effect on the application of SSEC as in factor 2)a)i) of Table A-3. Similarly, account for the static pressure measurement error in two separate ways: the main effect will be via the “pressure measurement and conversion,” but a secondary effect will be via factor 2)a)ii) of Table A-3.
Appendix A A-20 Table A-2. Static Source Error (Cause: Aerodynamic Disturbance to Free-Stream Conditions) Factors Error Components Airframe Effects Operating Condition (M, Hp, ∝ , β)* Geometry: S hape of airframe Location of static sources Variations of surface contour near the sources Variations in fit of nearby doors, skin panels, or other items
1) Reference SSE values from flight calibration measurements.
2) Uncertainty of flight calibration measurements.
3) Airframe-to-airframe variability.
4) Probe/port -to-probe/port variability.
Probe/Port Effects Operating Condition (M, Hp, ∝ , β)* Geometry: Shape of probe/port Manufacturing variations Installation variations
*M Mach, speed;
Hp pressure altitude;
∝ angle of attack (AOA);
β yaw (sideslip).
Appendix A A-21 Table A-3. Residual Static Source Error (Aircraft with Avionic Static Source Error Correction) (Cause: Difference Between the Static Source Error Correction Actually Applied and the Actual Static Source Error) Factors Error Components
1) As for SSE.
PLUS
2) Source of input data for SSEC function:
a) Where SSEC is a function of Mach:
i) PS sensing: difference in SSEC from reference SSE.
ii) PS measurement: pressure transduction error.
iii) PT errors: mainly pressure transduction error.
b) Where SSEC is a function of angle of attack (AOA):
i) G eometric effects on alpha:
• Sensor tolerances.
• Installation tolerances.
• Local surface variations.
ii) Measurement error:
• Angle transducer accuracy.
3) Implementation of SSEC function:
a) Calculation of SSEC from input data.
b) Combination of SSEC with uncorrected height.
1) Static Source Error Components 2),
3), and 4) from Table A-2.
PLUS 2a) Approximation in fitting design SSEC to flight calibration reference SSE.
2b) Effect of production variability (sensors and avionics) on achieving design SSEC.
2c) Effect of operating environment (sensors and avionics) on achieving design SSEC.
Appendix A Figure A-4. Static Source Error/Static Source Error Correction Relationships for Altimetry System Error Where Static Line, Pressure Measurement, and Conversion Errors Are Zero Actual Altitude Static Source Error Components
(2), (3) and (4) (see Table A-2) Residual Static Source Error SSE Component 2(a) (see Table A-3) Residual Static Source Error (Total) Perfect Residual Static Source SSEC Components 2(b) & 2(c) (see Table A-3) Reference Hp (Corrected) SSEC Design SSEC Actual SSEC Hp (Uncorrected) A.8.2.3.5 Static line errors arise from leaks and pneumatic lags. In level cruise, these can be made negligible for a system correctly designed and correctly installed.
A.8.2.3.6 Pressure measurement and conversion error:
1. The functional elements are static pressure transduction (which may be mechanical, electromechanical, or solid-state) and the conversion of pressure signal to pressure altitude. The error components are:
• Calibration uncertainty;
• Nominal design performance;
• Unit-to-unit manufacturing variations; and
• Effect of operating environment.
A-22
Appendix A A-23
2. The equipment specification usually covers the combined effect of the error components. If the value of pressure measurements and conversion error used in the error budget is the worst-case specification value, then it is not necessary to assess the above components separately. However, calibration uncertainty, nominal design performance, and effect of operating environment can all contribute to bias errors within the equipment tolerance. Therefore, if it is desired to take statistical account of the likely spread of errors within the tolerance band, it will be necessary to assess their likely interaction for the particular hardware design under consideration.
3. It is particularly important to ensure the specified environmental performance is adequate for the intended application.
A.8.2.3.7 Baro-setting error is defined as the difference between the value displayed and the value applied within the system. For RVSM operation, the value displayed should always be International Standard Atmosphere (ISA) standard ground pressure, but setting mistakes, although part of TVE, are not components of ASE.
1. The components of the baro-setting error are:
• Resolution of setting knob/display (“Setability”);
• Transduction of displayed value; and
• Application of transduced value.
2. The applicability of these factors and the way they combine depends on the particular system architecture.
3. For systems in which the display is remote from the pressure measurement function there may be elements of the transduction and/or application or transduced value error components arising from the need to transmit and receive the setting between the two locations.
A.8.2.3.8 Imperfect conversion from altitude signal to display causes display error. The components are:
• Conversion of display input signal;
• Graticule/format accuracy; and
• Readability.
Note: In self-sensing altimeters, the first of these would normally be separate from the pressure measurement and conversion error.
Appendix A A-24 A.8.3 A SE Component Error Budget: Group Approval. Where approval is sought for an aircraft Group, the data package must be sufficient to show the requirements of paragraphs A.5.1 and A.5.2 are met. Because of the statistical nature of these requirements, the content of the data package may vary considerably from Group to Group. Paragraph A.8 should serve as a guide to properly account for ASE sources.
A.8.3.1 E stablish the mean and airframe-to-airframe variability of ASE based on precision flight test calibration of a number of aircraft. Where analytical methods are available, it may be possible to enhance the flight test database and to track subsequent change in the mean and variability based on geometric inspections and bench tests or any other method acceptable to the approving authority. In the case of derivative aircraft, it may be possible to utilize data from the parent as part of the database (e.g., a fuselage stretch where the only difference in ASEmean between Groups could be reliably accounted for by analytical means).
A.8.3.2 All avionics equipment contributing to ASE must be identified by function and part number. The applicant must demonstrate the avionics equipment can meet the requirements established according to the error budget when operating the equipment in the environmental conditions expected to be met during RVSM operations.
A.8.3.3 Assess the aircraft-to-aircraft variability of each error source. The error assessment may take various forms as appropriate to the nature and magnitude of the source and the type of data available. For example, for some error sources (especially small ones) it may be acceptable to use specification values to represent 3 SD. For other error sources (especially larger ones), a more comprehensive assessment may be required; this is especially true for airframe error sources where “specification” values of ASE contribution may not have been previously established.
A.8.3.4 In many cases, one or more of the major ASE sources will be aerodynamic in nature (such as variations in the aircraft surface contour near the static pressure source). If evaluation of these errors is based on geometric measurements, substantiation should be provided that the methodology used is adequate to ensure compliance. (See Figure A-6, Compliance Demonstration Ground-To-Flight Test Correlation Process Example.)
A.8.3.5 I n showing compliance with the overall requirements, combine the component error sources in an appropriate manner. In most cases, this will involve the algebraic summation of the mean components of the errors, root sum square (RSS) combination of the variable components of the errors, and summation of the RSS value with the absolute value of the overall mean. Be sure the RSS combines only variable component error sources independent of each other.