Appendix_A_-_Attachment_1_-_SRD.pdf

APPENDIX A, ATTACHMENT 1

B-1B

SELF-CONTAINED ATTITUDE INDICATOR (SCAI)

SYSTEM REQUIREMENTS DOCUMENT

VERSION 1.0

SUPPORTING MILESTONE B

ENGINEERING & MANUFACTURING DEVELOPMENT

And

PRODUCTION & DEPLOYMENT

January 29, 2014

Table of Contents

Revision History ........................................................................................................................... iv

1. SCOPE/INTRODUCTION

1.1. Program Description and Applicable Documents

1.2. Current Program Status

2. APPLICABLE DOCUMENTS

2.1. Order of Precedence

3. REQUIREMENTS

3.1. Functional Requirements

3.1.1. Standby Altimeter:

3.1.2. Standby Airspeed/Mach Indicator

3.1.3. Standby Attitude Indicator

3.1.4. Symbology Requirement

3.2. Environmental Conditions

3.2.1. Temperature and Pressure

3.2.2. Explosive Decompression

3.2.3. Explosive Atmosphere

3.2.4. Humidity

3.2.5. Sand and Dust

3.2.6. Salt Fog and Spray

3.2.7. Fungus

3.2.8. Solar Radiation

3.2.9. Ozone Exposure

3.2.10. Fogging

3.2.11. Warm-up

3.3. Dynamic Environment

3.3.1. Vibration

3.3.2. Shock

3.3.3. Linear Acceleration

3.4. Human Factors Requirements

3.4.1. Display Luminance

3.4.2. Display Veiwing Envelope (PVE & EVE)

3.4.3. Display Contrast Ratio

3.4.4. Touch Temperature Limits

3.5. Restrictions

3.6. General Requirements

3.6.1. Unit Interface

3.6.2. Electromagnetic Compatibility

3.6.3. Power

3.6.4. Weight

3.6.5. Input Signal Description

3.6.6. Envelope Size (Dimensions)

3.6.7. Display Controls

3.6.8. Reliability

ii

3.6.9. Environmental Conditions

3.6.10. Response Time

3.6.11. Display Defects

3.6.12. Display Refresh Rate

3.6.13. Display Resolution

3.7. Failure Reporting

3.7.1. Built-In-Test (BIT)

3.7.2. Power-Up BIT (PBIT)

3.7.3. Initiated BIT (IBIT)

3.7.4. Continuous BIT (CBIT)

3.7.5. BIT Fail

3.7.6. Fault Isolation

3.7.7. Failure of SCAI System Test Provisions

3.7.8. Fault/Failure Reporting Interface

4. VERIFICATION PROVISIONS

4.1 Verification Methods

4.1.1 Demonstration

4.1.2 Test

4.1.3 Analysis

4.1.4 Inspection

4.1.5 Special Verification Methods

5. REQUIREMENTS TRACEABILITY

5.1 Traceability to Capability Document or System Specification

5.2 Traceability to Subsystems Requirements

APPENDIX A: ACRONYMS AND DEFINITIONS

APPENDIX B: KEY PERFORMANCE PARAMETERS/KEY SYSTEM ATTRIBUTES 14

APPENDIX C: REQUIREMENTS TRACEABILITY

APPENDIX D: VERIFICATION MATRIX

Verification Matrix iii

Revision History

Date Revision Number Reason for Update

1 29 2014 1.0 Initial publication iv

1. SCOPE/INTRODUCTION

1.1. Program Description and Applicable Documents

The Oklahoma City Air Logistics Center, B-1 System Program Office (SPO) requires replacement of the current B-1B standby flight instruments. The current Standby Altimeter, Standby Airspeed/Mach Indicator, and the Standby Attitude Indicator shall be replaced with a Self-Contained Attitude Indicator (SCAI) system. The SCAI system shall incorporate a single Line Replaceable Unit (LRU) that provides the same functionality and meets the same environmental specifications of the current B-1 standby instrumentation identified above.

1.2. Current Program Status

Part obsolescence and increasing failure rates drove the B-1 SPO to begin pursuing a replacement three-in-one unit system for the standby indicators from as early as 2006. A Boeing Engineering Assignment (EA) was conducted in 2006 to establish evaluation criteria for the replacement system and to compile a matrix of the capabilities of potential candidate vendors. In 2009, another Boeing EA was conducted to produce preliminary specifications for the system and to evaluate a proposed vendor. A Request for Proposal (RFP) is currently being sent to evaluate competitive proposals for any required additional development, production, and delivery of the replacement system.

2. APPLICABLE DOCUMENTS

The iterative nature of document development creates an interrelationship between this SRD and the documents listed in Table 1. These documents provide guidance and requirements for the design, development, and validation of the system.

2.1. Order of Precedence

In the event of a conflict between this document and other documents, the requirements of this specification shall be considered as the superseding requirement. Nothing in this document, however, supersedes applicable laws and regulations unless a specific exemption has been obtained.

Table 1- Applicable Documents Guidance Documents Document, Name Responsible Org.

AFI 62-601, USAF Airworthiness HQ USAF AFI 63-101, Acquisition and Sustainment Life Cycle Management HQ USAF AFI 91202, The US Air Force Mishap Prevention Program HQ USAF AFI 99-103, Capabilities Based Test and Evaluation HQ USAF/TEP AFI 163-107, Integrated Product Support Planning and Assessment HQ USAF/ILMM AFMAN 63-119, Certification of System Readiness for Dedicated Operational Test & Evaluation SAF/AQXA AFMCI 23-103, Diminishing Manufacturing Sources and Material Shortages (DMSMS) Program DAG 4000.1, Defense Acquisition Guidebook DoD 5010.12-M, Procedures for Acquisition and Management of Technical Data OASD (P&L) DoD Directive 5000.1, The Defense Acquisition System USD (AT&L) DoD Directive 5000.2, Operation of the Defense Acquisition System USD (AT&L) DoDI 4630.8, Procedures for Interoperability and Supportability of Information Technology (IT) and National Security Systems (NSS) ASD(NII)/DoD CIO

DoDI 5000.02, Operation of the Defense Acquisition System DoDI 8500.2, Information Assurance Implementation ASD (C3I) MIL-HDBK-514, Operational Safety, Suitability, and Effectiveness for the Aeronautical Enterprise ASC/ENOI

MIL-HDBK-515, Weapon System Integrity Guide ASC/ENOI MIL-HDBK-1530, Aircraft Structural Integrity Program, General Guidelines For ASC/ENOI MIL-HDBK-1798, Mechanical Equipment and Subsystems Integrity Program ASC/ENOI MIL-STD-882D, Department of Defense Standard Practice for System Safety AFMC/SES Systems Engineering Plan Preparation Guide OUSD (AT&L) TO 00-35D-54, USAF Deficiency Reporting and Investigation System AFMC/A4UE SS2569A01, B-1B System Specification AFLCMC/WWN GS2569A03, Vibration, Noise, Acoustics, Shock, Accel. Criteria AFLCMC/WWN GS2569A05, Thermodynamic Design & Test Criteria AFLCMC/WWN L432C2060, Performance Spec, Attitude Indicator AFLCMC/WWN MIL-PRF-27197E, Performance Spec, Airspeed/Mach Indicator AFLCMC/WWN MIL-A-83212A, Performance Spec, Standby Altimeter AFLCMC/WWN MIL-STD-1787C, Aircraft Display Symbology MIL-STD-210A, Atmosphere Models MIL-STD-1472, Design Criteria Human Engineering ARINC 408, Casing and Mounting MIL-C-5541, Coating MIL-DTL-38999, Connectors MIL-STD-810B, Environmental Test Methods MIL-STD-461E, Electromagnetic Compatibility MIL-STD-464C, Electromagnetic Environmental Effects http://www.e-publishing.af.mil/pubfiles/af/99/afi99-103/afi99-103.pdf http://www.e-publishing.af.mil/pubfiles/af/63/afman63-119/afman63-119.pdf http://www.e-publishing.af.mil/pubfiles/af/63/afman63-119/afman63-119.pdf http://www.dtic.mil/whs/directives/corres/pdf/501012m_0593/p501012m.pdf http://www.dtic.mil/whs/directives/corres/pdf/d50001_051203/d50001p.pdf http://www.dtic.mil/whs/directives/corres/pdf2/i50002p.pdf http://www.dtic.mil/whs/directives/corres/pdf2/i46308p.pdf http://www.dtic.mil/whs/directives/corres/pdf2/i46308p.pdf http://www.assistdocs.com/search/document_details.cfm?ident_number=213028&StartRow=1&PaginatorPageNumber=1&doc%5Fnumber=514&search%5Fmethod=BASIC http://www.assistdocs.com/search/document_details.cfm?ident_number=213028&StartRow=1&PaginatorPageNumber=1&doc%5Fnumber=514&search%5Fmethod=BASIC

3. REQUIREMENTS

3.1. Functional Requirements

The following subparagraphs list the operational requirements of the system on the B-1B.

Detailed requirements, along with source information, can be found in the Requirements Breakdown Matrix (RBM)

3.1.1. Standby Altimeter:

3.1.2. Standby Airspeed/Mach Indicator

1 Airspeed Indication Indicate KIAS/KCAS with reference to an airspeed scale

2 Airspeed Scale Markings Display 5/10/100 knot increments 3 Airspeed Scale Range Allow for an airspeed range of 0 knots through 850 knots 4 Maximum Allowable Airspeed Indicate maximum allowable airspeed which shall be adjustable from 600 to 800 Knots

5 Airspeed Marker Indicate an airspeed reference 6 Airspeed Marker Set Allow for setting of a desired airspeed reference 7 Mach Number Indication Indicate Mach number 8 Mach Number Indication Range Allow for a Mach number range of 0.5 through 2.2 Mach

1 Altitude Indication Indicate altitude in feet with reference to altitude scale 2 Altitude Increments Display 10/100/1,000/10,000 foot increments 3 Altitude Range Allow for an barometric altitude range of -1,000 feet through +50,000 feet 4 Barometric Set Allow for barometric pressure adjustment 5 Barometric Set Indication Display barometric pressure adjustment values 6 Barometric Set Range Allow for barometric pressure adjustment range of

28.10 through 31.00 inches of Hg in increments of hundredths of an inch of Hg

3.1.3. Standby Attitude Indicator

1 Pitch Angle Range Measure pitch attitude climb/dive from 0 through 90 degrees.

2 Pitch Angle Indication Indicate pitch angle in 5-degree increments.

3 Bank Angle Range Measure bank angles left and right from 0 through 90 degrees.

4 Bank Angle Indication Indicate bank angle increments 0, 10, 20, 30, 60, and 90 degrees

5 Horizon Symbol Indicate Sky/Earth horizon.

6 OFF/Fail Indication Indicate loss of power, low voltage, failure of power conversion, or when attitude data is incorrect. Also indicate the Cage function engaged, if applicable.

7 Aircraft Symbol Indicate the aircraft attitude in relation to the earth's surface.

8 Caging/Pitch Trim Allow for manual fast erection of indicator gyro and pitch adjustment of +10 to -5 degrees of aircraft symbol.

9 Slip/Rate of turn Indication

Indicate aircraft lateral acceleration.

3.1.4. Symbology Requirement

Symbology shall be in accordance with MIL-STD-1787C, Aircraft Display Symbology.

3.2. Environmental Conditions

The system shall provide the specified performance when subjected to the following natural and induced environmental conditions in accordance with MIL-STD-810B.

Detailed requirements, along with source information, can be found in the Requirements Breakdown Matrix (RBM).

3.2.1. Temperature and Pressure

The system shall be designed to be compatible with the surrounding air temperature and pressure criteria specified in the table below. Normal and emergency conditions relate to the Environmental Control System (ECS) operation. The system shall operate during the 5-minute duration emergency temperature or pressure conditions. The compartment pressure test must last for duration of 1 hour.

Environmental Parameters Temperature °F (°C)

Pressure (lb/in2)

Ground Non-operating -40 to +200 (-40 to +93)

10.1 to 15.4

Operating -40 to +160 (-40 to +71)

10.1 to 15.4

Compartment pressure test -- 27.0 In-flight Normal (Operating and Non-operating)

Minimum pressure -40 to +160 (-40 to +71)

10.4

Maximum pressure -40 to +160 (-40 to +71)

15.4

Emergency (Operating and Non-operating)

5-minute duration -40 to +160 (-40 to +71)

2.14 to 15.4

Extended duration -40 to +160 (-40 to +71)

2.7 to 5.45

Transient 1.0° F/sec Explosive Decompression

3.2.2. Explosive Decompression

The system must be in a pressurized compartment and shall withstand a decompression of the surrounding air from an initial pressure of 10.9 lb/in2 a to a final pressure of 2.14 lb/in2 a in 0.20 seconds.

3.2.3. Explosive Atmosphere

The system shall be designed for operation in an explosive atmosphere without causing ignition of such atmosphere. If sealing or flame-arresting construction is necessary to meet this requirement, all system components shall contain an explosure occurring within the sealed enclosure.

3.2.4. Humidity

The system designed range for humidity in the surrounding air shall be from zero to 182 grains of water vapor per pound of dry air. The maximum value of 182 grains per pound is equivalent to an 85°F (29°C) dew point temperature, which corresponds to the climatic extreme specified in MIL-STD-210A.

3.2.5. Sand and Dust

The system shall be designed for exposure for an environment of particle sizes from 0.1 to 1,000 micrometers in diameter, with predominant diameters less than 300 micrometers.

The sand and dust concentration shall be as specified in MIL-STD-210A.

3.2.6. Salt Fog and Spray

The system shall be designed for exposure to salt-sea atmosphere having particular sizes as small as 1.0 micrometer in diameter.

3.2.7. Fungus

The system shall be designed for exposure to fungus growth as encountered in tropical climates.

3.2.8. Solar Radiation

The forward crew station system display face shall withstand solar radiant energy at a maximum rate of 254 BTU / hour / ft2 at sea level. Above, 30,000 feet altitude, the forward crew station system display face maximum solar radiant energy rate shall be 307 BTU / hour/ ft2.

3.2.9. Ozone Exposure

All system components shall withstand, without deterioration, exposure to an ozone concentration from 0.10 part per million (p/m) by volume at sea level to 6 p/m at 60,000 ft.

3.2.10. Fogging

The indicator shall be designed and constructed so the face of the indicator does not become fogged after operation for a period of 1 hour at a temperature of 131°F, when an ice cube is rubbed on the glass face for a period of no less than 1 minute.

3.2.11. Warm-up

The system warm-up time shall not exceed 3 minutes at any temperature from -40°F to 160°F (-40°C to 71°C). Warm-up shall be defined as the transition from a non-operational condition, power off or standby, to an operational condition which meets the requirements of this specification.

3.3. Dynamic Environment

3.3.1. Vibration

The system shall perform as specified while exposed to the following random vibration ranges, for test duration of 10 hours per axis using the test tolerances of MIL-STD-810B (Change Notice 2) Method 514.1 Procedure II, Part 3.

Frequency range

(HZ)

Acceleration Power Spectral

Density 20 to 300

300 to 1000 1000 to 2000

+3 dB per Octave .02g2/Hz

-6db per Octave

3.3.2. Shock

The system shall perform as specified when subjected to a saw tooth shock pulse of 20g peak magnitude with duration of 11 (±1) milliseconds. The system shall not represent a hazard during or after exposure to a crash safety shock requirement of a saw tooth shock pulse of 40G peak magnitude with duration of 11 (±1) milliseconds with the unit powered as for normal operation, using the test tolerances of MIL-STD-810B (Change Notice 2) Method 516.1 Procedure I and II.

3.3.3. Linear Acceleration

The system shall perform as specified during and after exposure to the following levels of acceleration. These levels shall be applied individually along the three mutually perpendicular aircraft axes for duration of 1 minute. (Note: The installed system indicator shall sit tilted 10 degrees from the horizon with the connector end down.) In addition, the system shall meet the following crash safety acceleration requirements applied individually along the three mutually perpendicular aircraft axes for a minimum duration of 1 minute, with the system powered for all tests.

Functional Acceleration (Based Upon Installed Location Near Yf210) Normally Mounted Indicator

Vertical Axis (Height) 4.1 g upward 2.8 g downward Longitudinal Axis (Width) 4.2 g aft 0.8 forward

Lateral Axis (Length) +/- 1.2 g Crash Safety Acceleration

Normally Mounted Indicator Vertical Axis (Height) 25.0 g upward 10.0 g downward

Longitudinal Axis (Width) 40.0 g aft 9.0 g forward Lateral Axis (Length) +/- 10.0 g

3.4. Human Factors Requirements

The system shall be operationally compatible with a ‘gloved hand’. Bezel buttons, knobs and other controls shall be operable by the user per MIL-STD-1472. The system shall be NVIS Class B compatible with a luminance between 0.02 to 12.0 fL. The system shall provide the following optical characteristics.

3.4.1. Display Luminance

The display shall have a dimming range from 0.02 ftL to 250FtL.

3.4.2. Display Viewing Envelope (PVE & EVE)

The viewing angles for the display must be usable at 0° - 40° laterally and 0° - 10° vertically in high ambient conditions up to the following goal: 10,000 fC ambient illumination and 2000 fL specular glare.

3.4.3. Display Contrast Ratio

Display Contrast Ratio Low Ambient Minimum Contrast Ratio: 70: 1 ((white-black)/Black) on

Central Viewing Axis Percept # Gray Shades (min): 64 or more per primary color High Ambient Contrast Ratio ((White-Black)/Black): 9:1 over viewing angle Perceptible # Gray Shades (min): 14 or more per primary color

3.4.4. Touch Temperature Limits

The system during normal operation must not expose personnel to surface temperatures greater than or less than those shown below:

Exposure Metal Glass Plastic Momentary (Hot) 60°C or 140°F 68°C or 154°F 85°C or 185°F Prolonged (Hot) 49°C or 120°F N/A 69°C or 156°F Momentary (Cold) 0°C or 32°F 0°C or 32°F 0°C or 32°F Prolonged (Cold) 0°C or 32°F 0°C or 32°F 0°C or 32°F

3.5. Restrictions

No internal cooling fans are allowed.

3.6. General Requirements

3.6.1. Unit Interface

The system installation must mate with the following connectors:

• PITOT connector (7/16-20UNF-3B)

• STATIC connector (7/16-20UNF-3B)

• ES27484T12F35S Electrical Plug (Deviations from the connector specified shall be entertained.)

3.6.2. Electromagnetic Compatibility

The system must demonstrate compliance with the electromagnetic effects compatibility requirements per MIL-STD-464C and MIL-STD-461E for class A1 equipment.

3.6.3. Power

Total power required shall not exceed 50 watts. The system shall use the following MIL- STD-704E compliant power:

• 1 Amp, 28 V DC

• 2 Amp, 28 V DC

3.6.4. Weight

The system shall weigh no more than 5 pounds.

3.6.5. Input Signal Description

• Standby Altimeter input PS.

• Standby Airspeed/Mach Indicator inputs PS and PT.

3.6.6. Indicator Envelope Size (Dimensions)

The case of the unit shall be similar in concept to a standard Air Transport Indicator (ATI) flange-mounted vertical scale case per ARINC specification 408. However, the physical size shall be such that it shall fit in design standard cutout LP032-0024-0001.

The mounting flange of the unit should be square and shall not exceed 3.24 by 3.24 inches. The unit shall have a maximum depth of 9.05 inches measured from the mounting flange surface that contacts the front face of the aircraft instrument panel.

Including all electrical connector(s) the entire indicator envelope size shall have a maximum depth of 9.75 inches. The unit shall be secured to the aircraft instrument panel by four (4) No. 6-32 captive screws conforming to the dimensions in design standard cutout LP032-0024-0001. The electrical bonding surface where the mounting flange contacts the front face of the aircraft instrument panel shall have a MIL-C-5541, Class 1A chemical film surface treatment. Electrical connectors shall be Series III MIL-DTL- 38999. Deviations from the connector specified shall be entertained. The envelope shall include the type of connectors for electrical power and signals.

3.6.7. Display Controls

At a minimum, luminance control must be provided for any edge or panel lighting. The illumination control must be phase modulated 0 to 5V AC 400 Hz.

3.6.8. Reliability

Mean-Time-Between-Failure (MTBF) shall be projected to be at least 10,000 hours.

3.6.9. Environmental Conditions

The system must be naturally cooled.

3.6.10. Response Time

The display response time while surrounded by the extreme ambient temperatures shall be no longer than the following:

T>0°C, 2 minutes 0°C >T>-18°C, 4 minutes -18°C >T>-40°C, 5 minutes

3.6.11. Display Defects

The total number of failures shall be less than 3 elements per million.

3.6.12. Display Refresh Rate

The minimum display refresh rate shall be 60 Hz.

3.6.13. Display Resolution

The display resolution shall fully support all format and test information presented on the display for viewing distances from 26.0 to 48.0 inches.

3.7. Failure Reporting

3.7.1. Built-In-Test (BIT)

The system shall support Built-In Test (BIT). The system shall report status in all modes of operation during flight or on the ground without being supplemented by support equipment via the PCI bus. The system shall verify acceptable performance, detect failures, identify affected functions, and isolate failures within the system. A failure shall be defined as any condition wherein the performance requirements of this specification are not fulfilled. The BIT shall be capable of detecting 95 percent of testable board failures. The BIT interface including control and BIT reporting shall be provided via Seller-provided library routines hosted on the system. The system shall report the BIT status via a call to the Application Program Interface (API) which is provided by the Seller.

3.7.2. Power-Up BIT (PBIT)

The system shall perform a self test prior to initialization at power up when initiated via the API. Upon power up and the API initiation, the system shall report back the status of the PBIT system operation. The standard PBIT execution time, as delivered, shall not exceed 500 milli-seconds.

3.7.3. Initiated BIT (IBIT)

The Initiated BIT (IBIT) shall be performed when initiated via the API. The system shall be able to perform test on, but not limited to, RAM, register, PCI bus interface, and internal system link loop back. The system shall conduct an IBIT when initiated via calls to the API. Normal system operation may be interrupted during IBIT. This test shall verify the proper operation of all modes of system operation.

3.7.4. Continuous BIT (CBIT)

The system shall provide a non-disruptive CBIT that is executed as a background task or upon calls to the API. The CBIT capability shall be provided for automatic real-time determination of system performance. The CBIT shall detect failures and report affected modes and functions. CBIT shall not interrupt normal system operation or cause any other un-commanded system operation. The CBIT shall at a minimum report message or block count, bad message or block count. The frequency of CBIT shall be between 0.5 seconds and 10 seconds.

3.7.5. BIT Fail

Whenever a system failure is detected, the system shall transmit confirmation of the failure to the aircrew via audible and/or visual warning within 3 seconds.

3.7.6. Fault Isolation

The system shall isolate 95 percent of detected failures, based on failure rate. The system isolation shall be accomplished by PBIT, CBIT and IBIT.

3.7.7. Failure of system Test Provisions

Failure of system test provisions shall not cause failure or degradation of system operational performance.

3.7.8. Fault/Failure Reporting Interface

The system shall be capable of displaying fault and failure information (BIT and Power- On Self-Test) to the user.

4. VERIFICATION PROVISIONS

4.1 Verification Methods

The requirements stated within this document must be verified to show that they have been met by the system. The following methods shall be utilized to perform requirement verification. A verification matrix tracing requirements to verification methods can be found in Appendix D. The Requirements Breakdown Matrix contains a detailed test strategy for each requirement.

4.1.1 Demonstration

Demonstration is the operation of the system, subsystem, or a part of the system that relies on observable functional operation not requiring use of instrumentation, special test equipment, or subsequent analysis.

4.1.2 Test

Test is the operation of the system, subsystem, or a part of the system, using instrumentation or other special test equipment to collect data for later analysis. Test can be performed on the aircraft as a ground test or during aircraft flight.

4.1.3 Analysis

Analysis is the processing of accumulated data obtained from other qualification methods. Examples are reduction interpolation, or extrapolation of test results.

4.1.4 Inspection

Inspection is the visual examination of system components, documentation, etc.

4.1.5 Special Verification Methods

Special verification methods are used when verification cannot be accomplished by demonstration, test, analysis, or inspection alone. Special verification methods for the system or subsystem include but are not limited to special tools, techniques, procedures, facilities, acceptance limits, use of standard samples, preproduction or periodic samples, pilot models, or pilot lots.

5. REQUIREMENTS TRACEABILITY

System requirements stem from the B-1B System Specification SS2569A01, the legacy indicator specifications (L432C2060, MIL-PRF-27197E, and MIL-A-83212A), as well as a risk reducing engineering analysis. Requirement traceability is important to ensure that the system performs as specified and can complete the intended mission.

5.1 Traceability to Capability Document or System Specification System requirements stem from the B-1B System Specification SS2569A01, the legacy indicator specifications (L432C2060, MIL-PRF-27197E, and MIL-A-83212A). The requirements in this SRD form a direct link to the above documents. Contractor developed documents shall provide a direct, traceable link to the requirements held in this document. Each requirement shall be tracked by the contractor in a database, such as the Dynamic Object Oriented Requirements System (DOORS) or equivalent, and shall be reviewed at each technical review. In addition, each requirement is tracked in a Requirements Breakdown Matrix (RBM) which shows the requirement/derived requirement and its traceability to the appropriate requirements document.

5.2 Traceability to Subsystems Requirements

N/A

APPENDIX A: ACRONYMS AND DEFINITIONS

A/C Aircraft

AFB Air Force Base

API Application Program Interface

ATI Air Transport Indicator

CBIT Continuous Built-in-Test

BIT Built-in-Test

DT Developmental Test

EA Engineering Assignment

EMI/EMC Electromagnetic Interference/Electromagnetic Compatibility

IBIT Initial Built-in-Test

KCAS Knots Calibrated Airspeed

KIAS Knots Indicated Airspeed

KPP Key Performance Parameter

KSA Key System Attribute

MHz Megahertz

MTBF Mean Time Between Failure

NM Nautical Miles

OC-ALC Oklahoma City Air Logistics Center

OT Operational Test

PBIT Power-up Built-in-Test

PCI Peripheral Component Interconnect

RBM Requirements Breakdown Matrix

SCAI Self-Contained Attitude Indicator

SPO System Program Office

SRD Systems Requirement Document

APPENDIX B: KEY PERFORMANCE PARAMETERS/KEY SYSTEM ATTRIBUTES

Key Performance Parameters:

KPP Title Description

KPP SCAI Performance Requirements

The system must meet the performance requirements of all three individual standby indicators: attitude indicator, altitude indicator, and airspeed/mach indicator.

KPP SCAI

Environmental Conditions

The system must be able to withstand all environmental condition requirements for the forward crew station which include the following:

temperature, pressure, explosive decompression, explosive atmosphere, humidity, sand and dust, salt fog and spray, fungus, solar radiation, ozone exposure, fogging, and warm-up.

KPP SCAI Dynamic Environment The system must be able to withstand all existing dynamic environment requirements: vibration, shock, and linear acceleration.

KPP SCAI Dimensional Requirements

Installation of the system must involve no modification to the existing structure. The indicator must fit within a footprint that does not exceed 3.24”x3.24” with a depth no longer than 9.05”.

APPENDIX C: REQUIREMENTS TRACEABILITY

Requirements Breakdown Matrix (RBM)

Requirements traceability is a key component to the Systems Engineering Process. The RBM serves as the matrix that shows traceability between the requirement and its corresponding requirements document. Refer to the SCAI system RBM for requirements traceability.

APPENDIX D: VERIFICATION MATRIX

The following matrix depicts how each requirement of the SRD shall be verified. The legend below indicates what type of validation(s) shall be performed. Verification methods are defined in Paragraph 4. The Requirements Breakdown Matrix contains a detailed test strategy for each requirement.

D - Demonstration:

T - Test A - Analysis I - Inspection

PARAGRAPH DESCRIPTION D T A I

3.1 Functional Requirements

3.1.1 Standby Altimeter X

3.1.2 Standby Airspeed/Mach Indicator X

3.1.3 Standby Altitude Indicator X

3.1.4 Symbology Requirement X

3.2 Environmental Conditions

3.2.1 Temperature and Pressure X

3.2.2 Explosive Decompression X

3.2.3 Explosive Atmosphere X X

3.2.4 Humidity X X

3.2.5 Sand and Dust X X

3.2.6 Salt Fog and Spray X X

3.2.7 Fungus X

3.2.8 Solar Radiation X

3.2.9 Ozone Exposure X

3.2.10 Fogging X

3.2.11 Warm-up X

3.3 Dynamic Environment

3.3.1 Vibration X

3.3.2 Shock X

3.3.3 Linear Acceleration X

3.4 Human Factors Requirements X

3.4.1 Display Luminance X

3.4.2 Display Viewing Envelope X X

3.4.3 Display Contrast Ratio X X

3.4.4 Touch Temperature Limits X X

3.5 Restrictions X

3.6 General Requirements

3.6.1 Unit Interface X

3.6.2 Electromagnetic Compatibility X X

3.6.3 Power X

3.6.4 Weight X

3.6.5 Input Signal Description X

3.6.6 Envelope Size X

3.6.7 Display Controls X X

3.6.8 Reliability X

3.6.9 Environmental Conditions X X

3.6.10 Response Time X

3.6.11 Display Defects X X

3.6.12 Display Refresh Rate X

3.6.13 Display Resolution X

3.7 Failure Reporting

3.7.1 Built-in-Test (BIT) X

3.7.2 Power-up BIT X

3.7.3 Initiated BIT X

3.7.4 Continuous BIT X

3.7.5 BIT Fail X

3.7.6 Fault Isolation X X

3.7.7 Failure of SCAI system Test Provisions X X

3.7.8 Fault/Failure Reporting Interface X