PARIS Aircraft Demonstrator (PAD) General Requirementsemits.sso.esa.int/emits-doc/ASTRIUMLIM/PARIS/General_Documents... · PARIS Aircraft Demonstrator ... CAN Controller Area Network

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PARIS Aircraft Demonstrator

ENG.SP.EU.03.0117 Issue C

Page 1 of 29

SP_0117_C GenRqts.doc


PARIS Aircraft Demonstrator (PAD)

General Requirements

Author Date:

Product Assurance Date:

Project Manager Date:

© EADS Astrium Limited, 2003

EADS Astrium Ltd owns the copyright of this document which is supplied in confidence and which shall not be used for any purpose other than that for which it is supplied and shall not in whole or in part be reproduced, copied, or

communicated to any person without written permission from the owner.

EADS Astrium Ltd Anchorage Road, Portsmouth, Hampshire, PO3 5PU, England



ENG.EP.EU.03.0117 Issue C

Page 2 of 29

EADS Astrium Ltd owns the copyright of this document which is supplied in confidence and which shall not be used for any purpose other than that for which it is supplied and shall not in whole or in part be reproduced, copied, or communicated to any person without written permission from the owner.



INTENTIONALLY BLANK




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CONTENTS

1. SCOPE AND INTRODUCTION ........................................................................................................... 7 1.1 Scope ......................................................................................................................................... 7 1.2 Introduction ................................................................................................................................. 7 1.3 Study Context.............................................................................................................................. 7 1.4 Objectives of Project .................................................................................................................... 7

2. APPLICABLE AND REFERENCE DOCUMENTS ................................................................................. 8 2.1 Applicable Documents ................................................................................................................. 8 2.2 Reference Documents ................................................................................................................. 8

3. PHYSICAL CHARACTE RISTICS ........................................................................................................ 9 3.1 General....................................................................................................................................... 9 3.2 Configuration............................................................................................................................... 9 3.3 Mass Properties .......................................................................................................................... 9

3.3.1 Mass ................................................................................................................................. 9 3.3.2 Mass Budget ...................................................................................................................... 9 3.3.3 Centre of Gravity ...............................................................................................................10 3.3.4 Moment of Inertia ..............................................................................................................10

3.4 Electrical Requirements ..............................................................................................................10 3.4.1 General ............................................................................................................................10

3.5 Thermal Requirements................................................................................................................10 3.5.1 General ............................................................................................................................10 3.5.2 Temperatures Limits..........................................................................................................10 3.5.3 Thermal Finishes ...............................................................................................................11 3.5.4 Cooling.............................................................................................................................11 3.5.5 Analysis............................................................................................................................11

4. MECHANICAL DESIGN AND INTERFACES.......................................................................................11 4.1 General......................................................................................................................................11 4.2 Operability..................................................................................................................................14

4.2.1 Reliability ..........................................................................................................................14 4.2.2 Safety...............................................................................................................................14 4.2.3 Useful Life.........................................................................................................................15

4.3 Environment ...............................................................................................................................15 4.3.1 On-Ground Mechanical Environment ..................................................................................15

4.4 Design and Construction .............................................................................................................16 4.4.1 General ............................................................................................................................16 4.4.2 Venting Provisions .............................................................................................................16 4.4.3 Mounting Provisions ..........................................................................................................16 4.4.4 Parts, Materials and Processes ..........................................................................................16 4.4.5 Moisture and fungus resistance..........................................................................................16 4.4.6 Corrosion prevention and control ........................................................................................17 4.4.7 Identification and marking ..................................................................................................17 4.4.8 Documents and standards .................................................................................................17

4.5 Electro-Magnetic Compatibility.....................................................................................................17 4.5.1 Bonding & Isolation............................................................................................................18

5. VERIFICATION, CERTIFICATION AND TEST...................................................................................18 5.1 General Requirements ................................................................................................................18

5.1.1 Responsibility for Tests......................................................................................................18 5.1.2 Airworthiness testing (reference AD2) .................................................................................18




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5.1.3 Certification for Airworthiness.............................................................................................19 5.2 Testing.......................................................................................................................................20

5.2.1 Test Conditions .................................................................................................................20 5.2.2 Testing for Airworthiness....................................................................................................21

6. .........................................................................................................................................................27

TABLES Table 3-1 Mass Budget .......................................................................................................................... 9 Table 3-2 Unit/ Sub-system Temperature Requirements..........................................................................11 Table 4-1 Applied Safety Factors ...........................................................................................................13 Table 4-2 Structural Failure Modes.........................................................................................................14 Table 4-3 Limit Accelerations for Ground Operations ...............................................................................15 Table 4-4 Limit Accelerations for Ground Operations ...............................................................................16 Table 5-1 EMC Applicability Matrix.........................................................................................................27

FIGURES Figure 1-1 Block schematic of the Airborne Demonstrator......................................................................... 8 Figure 4-1 Typical Configuration ...........................................................................................................12 Figure 5-1 Standard Random Vibration Test Curves for Equipment installed in Fixed Wing Aircraft with

Turbojet of Turbofan Engines .........................................................................................................22 Figure 5-2 Shock Measuring system Frequency Response......................................................................23 Figure 5-3 Terminal Saw-tooth Shock Pulse Configuration and its Tolerance Limits ..................................24 Figure 5-4 Thermal Test Cycle Profile ....................................................................................................25 Figure 5-5 Altitude Test .........................................................................................................................26




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ABBREVIATIONS

AC Alternating Current ESA European Space Agency ACT Advanced CMOS TTL -compatible

logic device ESTEC European Space Research and

Technology Centre

AD Applicable Document EU European Union ADD Architectural Design Document

(Software) FACT Framework Architecture for

Communication Technologies

AFT Abbreviated Functional Test FEM Finite Element Model AIT Assembly Integration and Test FESS Front -End Subsystem AIV Assembly, Integration and Verification GSE Ground Support Equipment

ASAR Advanced Synthetic Aperture Radar I/F Interface ASIC Application Specific Integrated Circuit IC Integrated Circuit BBE Base Band Equipment ICD Interface Control Doc/ Dwg BOL Beginning Of Life ID Identification C&C Command & Control IFE Intermediate Freq Equipment CAN Controller Area Network INS Inertial Navigation System CE Conducted Emissions IRF Impulse Response Function CMOS Complementary Metal Oxide

Semiconductor ISO International Standards Organisation

CNES Centre National d'Etudes Spatiales ITU International Telecommunications Union CORE Common Radar Electronics LNA Low Noise Amplifier CPC not Applicable LS Low Side (Drive Transistor) CRC Cyclic Redundancy Check LSAR L-Band Synthetic Aperture Radar CRESS Core Radar Electronics Subsystem L-SAR L-Band Synthetic Aperture Radar CS Cornerstone mission/ Cooler System LSB Least Significant Bit DAC Digital to Analogue Converter MAC Media Access Control DC Direct Current MGSE Mechanical Ground Support Equipment DCL Declared Components List MOI Moment of Inertia DEG not Applicable MS Material Specification DLC Data Length Code/ Digital Loop

Carrier MSB Most Significant Bit

DM Demonstrator Model N/A Not Applicable ECSS European Co-operation for Space

Standardisation PCB Printed Circuit Board

EEPROM Electrically Erasable Programmable Read Only Memory

PDM Pre-Development Model

EIA Electronic Industries Association PRF Pulse Repetition Frequency EIRP Equivalent Isotropic Radiated Power PRI Pulse Repetition Interval EMC Electro-Magnetic Compatibility PSU Power Supply Unit EOL End Of Life RAM Random Access Memory RF Radio Frequency RTR RT Remote Terminal RTR Remote Transmission Request-bit RX Receiver SAR Synthetic Aperture Radar SI System Imaging/Service Information




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SINDA Systems Improved Numerical Differencing Analyser

SLIO Serial Link I/O SP Sample Plan SPF Single Point Failure SSB Single Sideband STP Shielded Twisted Pair T/R Transmit/Receive TBC To Be Confirmed TBD To Be Determined TCS Thermal Control subsystem TCU Tile Control Unit TMTC Telemetry/Telecommand TPL Tribology Parts Library TR Tracking TTL Transistor-Transistor Logic TX Transmit UK United Kingdom VIP Video Interface Port WCA Worst Case Analysis




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1. SCOPE AND INTRODUCTION

1.1 Scope

This document defines the requirements for the general Requirements and Environments for Airworthiness which will be required for the Airborne Demonstrator. This demonstrator will use prototype hardware which will be produced to de-risk the development for an potential space borne instrument. 1.2 Introduction

Traditional nadir-looking altimeters have very limited capabilities to observe Mesoscale Ocean Topography. Mesoscale phenomena require higher spatial-temporal sampling in the order of 50 km – 7 days. The PARIS (Passive Reflectometry and Interferometric System) concept was conceived by ESA in 1993 as an alternative technique to increase the sampling of the ocean surface from a single Low Earth Orbit (LEO) platform. It consists in combining direct and ocean reflected signals from opportunity sources, such as Global Navigation Satellite System (GNSS) signals, to perform ocean altimetry. By-products of this processing are ocean wind, significant wave height and ionospheric Total Electron Content (TEC) over the ocean. The present activity is part of a mini-project that aims at the definition of a PARIS sensor primarily used for mesoscale ocean altimetry. 1.3 Study Context

The present activity titled PARIS Technology Concept Definition –Phase 2 (Phase 2 of TRP ETD -137.A) is part of the PARIS mini-project (TRP ETD-137). 1.4 Objectives of Project

The project is intended to provide Technology demonstrations of critical technologies which will be required for a space PARIS Mission. To achieve this end three technology projects will provide hardware which will be integrated into an Airborne Demonstrator. The overall system which will be used in the Airborne Demonstrator (PADS) is shown in Figure 1-1. This shows the major hardware items which will be required. Of these the following will be subject to Pre-development activities:

• Up-looking GPS Receiver • Multi-beam Array • Beamformer and Receiver • Signal Processor

The demonstrator items which are specified in this specification are intended to reduce the technology risks in a flight program. To achieve this the designs shall consider the implications of design for flight by choosing techniques and devices which ultimately can be used for space hardware.




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60 deg Raydome

AntennaLNAs & Phase Shifters

Pressure BulkheadCables

RX

Processor

Mass Storage

Upwardlooking RX

Upward Looking Antenna

PowerConverter

AircraftSupply

115V 400Hz

28VdcController Man Machine Interface

BFN

Electronic Assembly

FixingsAnd

Av Mounts Cabin Floor

INS

Part of aircraftsystem


60 deg Raydome



RX

Processor

Mass Storage

Upwardlooking RX


PowerConverter

AircraftSupply

115V 400Hz


BFN

Electronic Assembly

FixingsAnd


INS



Figure 1-1 Block schematic of the Airborne Demonstrator

2. APPLICABLE AND REFERENCE DOCUMENTS

2.1 Applicable Documents

AD1 Mil-Std-461 Electromagnetic Emission and Susceptibility Requirements for the Control of Electromagnetic Interference

AD2 RTCA/DO-160D Environmental Conditions and Test Procedures for Airborne Equipment

2.2 Reference Documents

RD1




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3. PHYSICAL CHARACTERISTICS

3.1 General

The hardware being developed will be used in an airborne demonstrator. However, it is intended that the results will represent a stage towards Spaceborne designs of important sections of the PARIS Sensor. Therefore, the electrical, mechanical and software should be capable of enhancement for a potential future space mission. The developed hardware will be integrated into a complete sensor which will be capable of being operated in an airborne environment. This infers certain levels of certification for airworthiness. 3.2 Configuration

The Multibeam antenna shall be mounted on the aircraft fuselage and will be behind a special radome to protect it from the external environment. All electronic equipment will be capable of being mounted within standard 19 inch racking which shall be assumed to be part of the aircraft equipment. Each equipment shall include its own power conditioners and any cooling fans which may be required for cooling. Alarm systems shall be provided in the event of the cooling fans failing. This alarm may be associated with an over temperature being measured. 3.3 Mass Properties

3.3.1 Mass

The overall mass shall be consistent with the budget provided in Table 3-1. This will not include infrastructure components of the aircraft such as the Radome, racking and Inertial Sensing Unit etc. 3.3.2 Mass Budget

The equipments shall meet the mass requirements as defined in the Table 3-1

Table 3-1 Mass Budget

Item Name Mass Budget (kg) Comments Up looking antenna 0.5 Navigation GPS

Receive 1

Signal Processor 18 Assumes hard-disc mass memory

Beam Former Receiver

8

Control Interface 5 Laptop PC 3 Multi Beam Array 27 Including Radome

and box

Inertial Sensing System

0 Not included in budget

Harness to Upward Looking antenna

2 Part of BFNRX

Main Electrical Harness

5

Harness to Multi beam array from

2 Part of BFNRX




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RX Miscellaneous 10 Margin 10 15% Total 91.5 3.3.3 Centre of Gravity

Not applicable to the demonstrator 3.3.4 Moment of Inertia

Not applicable to the demonstrator 3.4 Electrical Requirements

3.4.1 General

The aircraft shall supply 28 Vdc from a regulated source. The suppliers shall assume that some further conditioning shall be required. Other supplies may be available. Each supplier shall provide power requirements for each of the major units. 3.5 Thermal Requirements

3.5.1 General

The internal thermal design of the unit shall be the responsibility of the unit supplier. The thermal design is based on the following conditions. The unit/ subsystem shall be designed in such a way that component and junction temperatures shall not exceed the de-rated allowables for space application when subjected to the Qualification environment interface temperatures. Material and processes shall not exceed their de-rated allowables when subjected to the qualification environment interface temperatures. The equipment will be mounted into standard 19 inch racking and cooling will be by forced air by means of electric fans within the equipments or racking. 3.5.2 Temperatures Limits

The unit/ sub-system temperature limits are defined in Table 3-2 Temperatures and are measured on the external structure of the relevant units/ sub-system. The environmental temperature of the air within the aircraft shall be as per Table 3-2. Internal




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TEMPERATURE RANGE MINIMUM DEG C MAXIMUM DEG C Ambient +21 +23

Operating +0 +30 Acceptance -5 +35 Qualification Not applicable to the

demonstrator Not applicable to the

demonstrator Cold-switch-on -10 Non-operating -20 +55

Table 3-2 Unit/ Sub-system Temperature Requirements

These temperatures may not apply to the antenna if this is mounted outside of the aircraft in a non-pressurised radome. These temperatures will be in the range –80 to 50 deg C. 3.5.3 Thermal Finishes

Not relevant to the demonstrator 3.5.4 Cooling

The equipment will be mounted in a 19 inch rack assembly and the cooling will be in the form of forced cooling using electric fans. 3.5.5 Analysis

Sufficient analysis shall be carried out to ensure that all components and requirements stay within the specified temperatures. . 4. MECHANICAL DESIGN AND INTERFACES

4.1 General

The mechanical design shall be suitable for use in a pressurised or non-pressurised aircraft environment. The Antenna will be located behind a Radome which will be mounted in the lower part of the fuselage of the aircraft. Cables from this assembly will need to enter the cabin so it is expected that a special bulkhead will be required to pass the electrical connections from the cabin to the antenna.




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60 deg Raydome



RX

Processor

Mass Storage

Upwardlooking RX


PowerConverter

AircraftSupply

115V 400Hz


BFN

Electronic Assembly

FixingsAnd


INS



60 deg Raydome



RX

Processor

Mass Storage

Upwardlooking RX


PowerConverter

AircraftSupply

115V 400Hz


BFN

Electronic Assembly

FixingsAnd


INS



Figure 4-1 Typical Configuration

The Figure 4-1 shows a typical aircraft configuration. The Downward looking antenna is mounted behind the Radome and the signals passed into the cabin where the beams are formed in the BFN and then passed to the Receiver and so on. The electrical assembly will contain the electronic circuits for the:

• Upward Looking GPS Receiver • Downward Looking Receiver • Signal Processor • Mass Storage (Part of Signal Processor) • Beam Forming Network • Instrument Control Unit

The controller may be a Laptop Pcin conjunction with an Interface Unit which is then connected to the Electronic Assembly The Electronic Assembly will contain the functions as described and this will be mounted to the aircraft via anti-vibration mounts if necessary. 4.1.1.1 Mechanical Analysis

Sufficient mechanical analysis shall be carried out to show that the equipment shall not cause any hazard once it is operated inside an aircraft. 4.1.1.2 Safety factors

To account for uncertainties of load factor evaluation, material data and analysis, as well as to avoid the influence of manufacturing inconsistencies, safety factors shall be applied as shown in Table 4-1.

FACTOR TYPE FACTOR REMARKS

Qualification factor (jQ) Limit loads x 1.25. Qualification test levels Yield Factor (jY) Qualification loads x 1.25 Avoids permanent deformation

Ultimate Factor (jU) Qualification loads x 2 Avoids failure by rupture or buckling




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Insert Factor (jINS) Ultimate load x 1.5 For use in fixation insert design

Table 4-1 Applied Safety Factors

Notes on safety factors a) Failure resulting from local effects, e.g. stress concentration at cut-outs, inconsistencies of joining

elements shall be avoided by the application of an additional safety factor (jADD) to the qualification load. The use and magnitude of the factor shall be justified by prior experience or test data.

b) If the unit structure is verified by analysis only, then a factor of 1.25 for yield and 2 for ultimate load

conditions shall be applied to the qualification load. 4.1.1.3 Margin of safety

The unit shall be of adequate strength to withstand, without yielding or failing, all credible combinations of the specified loads and associated environment including assembly, integration, handling and transportation. All structural components shall have positive margins of safety (MS ≥ 0) for all load conditions and environments specified herein. 4.1.1.4 Evaluation of safety margins

For demonstration of the actual strength, a margin of safety evaluation is required. The evaluation shall be used in all cases of structural loading against surpassing the yield point and/or ultimate strength of materials as well as failure due to instability. The relevant surplus percentage has to be shown. The evaluation schematic is defined below:-

[ ]%100.1.

−=

σ

σ

jMs f

Where: σf = Material characteristics at failure point, eg material strength data,

Minimum value or critical instability data (allowables from material data sheet). σ = Maximum load/stress due to applied load. J = Relevant cumulative safety factor including jQ, jY, jU, jADD and jINS.

Material strength data for metal alloys as well as composites with a 99% statistical probability at a confidence level of 95% (A values) shall be used for MS evaluation. The long-term stability, moisture absorption and creep effects shall be considered appropriately. In case of bonded joints and composite materials, which have high strength variation due to the manufacturing process, the relevant design criteria shall be respected. The special failure modes, such as de-lamination, shall be avoided by application of the relevant failure criteria, design principles and product assurance activities.




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4.1.1.5 Material design allowable

Material strengths and other mechanical and physical properties shall be selected from authorised sources of reference, such as British Standards, and from contractor test values when appropriate. Strength allowables and other mechanical properties used shall be appropriate to the loading conditions, design environments and stress states for each structural member. Allowable material strengths used in design shall include the effects of temperature and time associated with the design environment. 4.1.1.6 Failure modes

The unit shall withstand all anticipated loads and environments without experiencing failures as defined in Table 4-2

FAILURE MODE ALLOWED ABOVE YIELD LOAD ULTIMATE LOAD Rupture X Instability (Global Buckling) X Yielding X Excessive elastic deformations (exceedance of allowed clearances)

X

Gapping of bolted connections X

Table 4-2 Structural Failure Modes

For verification of gapping, the maximum friction coefficient shall be used. For stress analyses of fasteners, the minimum friction coefficient shall be used. This requirement applies for the determination of pre-torque moments of bolts/screws in the thread. Limit load = design load exclusive of any safety factors. 4.1.1.7 Design loads

The mechanical environment shall be as defined for vibration and shock as defined in 4.3. 4.2 Operability

4.2.1 Reliability

4.2.1.1 Assessed Reliability

Not relevant to demonstrator 4.2.1.2 Single Point Failure

Not relevant to demonstrator 4.2.2 Safety

A safety assessment shall be carried out and potential hazards identified.




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4.2.2.1 Product Safety

Shall be compliant with the requirements of air safety. The equipment is expected to reside in a pressurised cabin so they should not be subjected to the extremes of environments. However, vibration, shock, temperature and reductions in pressure will be experienced. In the case of the antenna it may reside outside of the pressurised cabin so its pressure and temperature environment will be more severe. 4.2.3 Useful Life

The PADS shall be capable of operating for up to 5 hrs in one operation. The use of the PADS shall be in a semi-operational manner. For the purpose of this program the life can be assumed to be up to two years with a flight frequency of up to once per week. TBC. The equipment may be used between missions for test purposes up to 7 hours per day for up to once per week over the life. 4.3 Environment

4.3.1 On-Ground Mechanical Environment

4.3.1.1 Transportation and Handling loads

The unit transportation container and MGSE will be exposed to dynamic loads during road and air transportation and ground handling operations. The loads are defined for the load bed of the transportation vehicle. Transportation and handling the loads shall be applied as defined in Table 4-3. For handling operations the following loads shall be considered: Shock: equivalent to a fall from 100mm height when one edge of the Unit container remains on the ground. 4.3.1.2 MGSE Loads

The MGSE and transportation container shall be designed to limit the loads at the centre of gravity of the unit to less than the values defined in Table 4-4 when subjected to the environment specified in this specification

OPERATION X (+ AFT)

Y (+ STBD)

Z (+ UP)

REMARKS

(g) Scalar value Handling ± 1.0 ± 0.75 +2.0 / -0.7 Road Transport ± 1.1 ± 1.2 -3.0 / +1.0 Quasi-static Air Transport ± 1.5 ± 1.5 +2.2 / -1.5 Ship Transport ± 0.3 ± 0.5 +1.8 / -0.4 For MGSE only Transport shock load ± 4g (for a saw tooth form of 20 milliseconds duration)

Table 4-3 Limit Accelerations for Ground Operations

(Unattenuated Worst Case Input to MGSE)




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OPERATION X (+ AFT)

Y (+ STBD)

Z (+ UP)

REMARKS

(g) Scalar value Handling ± 1.0 ± 0.75 +2.0 / -0.7 Road Transport ± 0.8 ± 0.8 ± 2.0 Quasi-static Air Transport ± 1.5 ± 1.5 +2.2 / -1.5 Ship Transport ± 0.3 ± 0.5 +1.8 / -0.4 For MGSE only Transport shock load ± 1.5 ± 1.5 ± 2.0

Table 4-4 Limit Accelerations for Ground Operations

(Worse Case Input Allowables to Unit)

4.4 Design and Construction

4.4.1 General

The design and construction shall take into account to purpose of the demonstration and the need to produce an instrument which will be installed onto an aircraft. On the basis that the instrument will be used periodically then the installation and de-installation shall be considered during the design. 4.4.2 Venting Provisions

Provision shall be made for adequate venting of the internal free air space of the unit, in order to maintain internal and external pressure equilibrium. As a guideline the venting apertures should be less than 1.5mm diameter and have a depth to diameter ratio of at least 3:1. Typical thermal vacuum conditions of ≥ 1 x 10-5

Torr may be experienced. 4.4.3 Mounting Provisions

To be compatible with mounting in a 19 inch rack. The interface shall be agreed with the customer. 4.4.4 Parts, Materials and Processes

4.4.4.1 Parts

All critical parts used in the design of the unit shall be declared in a parts list. 4.4.4.2 Materials

All critical materials shall be declared in a materials list. 4.4.4.3 Processes

All critical processes shall be declared in a processes list. 4.4.5 Moisture and fungus resistance

Moisture traps shall be avoided. Parts and materials that may deteriorate when exposed to the climatic and environmental conditions during storage or service usage shall be protected against such deterioration.




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Materials, which are not nutrients for fungi, shall be used wherever practicable. Where nutrient materials have to be used outside of hermetically sealed containers, such materials shall be treated with fungicidal agent. Qualification of a particular composition shall be determined by test method 508 of MIL-STD-810, environmental test methods. 4.4.6 Corrosion prevention and control

4.4.6.1 Corrosion resistance

The design shall include provisions for corrosion prevention and control. Materials used shall resist damage from moisture, with particular emphasis on parts that lose their protective coatings during normal course of inspection, maintenance and periodic testing. 4.4.6.2 Dissimilar metals

The use of dissimilar metals shall be controlled in accordance with MIL-STD-889, dissimilar metals. 4.4.6.3 Protective treatments

Any materials used in the construction of the unit that may be subjected to deterioration when exposed to the climatic and environmental conditions likely to occur during any phase of the programme, shall be protected against such deterioration. Particular attention shall be paid to earth bonding requirements where bonding is achieved through the unit-mounting interface 4.4.6.4 Coatings and adhesives

Coatings and adhesives shall be capable of withstanding the environmental conditions defined in this specification without performance degradation. 4.4.7 Identification and marking

Each unit furnished under this specification shall be marked with a unique visual identification code. This shall not affect any aspect of the performance. 4.4.8 Documents and standards

Documentation used in the design, manufacture and test of the unit shall conform to general configuration requirements. Drawings and other documents required for the manufacture and test of the unit shall be listed in the unit drawing tree. All design and manufacturing standards and techniques shall be approved by the customer. Any change from the approved baseline shall only be implemented following formal configuration control procedure. 4.5 Electro-Magnetic Compatibility

TBD




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4.5.1 Bonding & Isolation

TBD 5. VERIFICATION, CERTIFICATION AND TEST

5.1 General Requirements

The design, manufacture and test of the PADS shall be verified as specified in this section by using established company standards, methods and procedures. The testing for certification for airworthiness shall be carried out as defined in the following sections. The supplier shall propose a series of analyses and testing based on the requirements of this specification. 5.1.1 Responsibility for Tests

The Supplier shall be responsible for all tests. 5.1.2 Airworthiness testing (reference AD2)

The PADS shall be designed to meet the environment as specified in RTCA/DO-160D to the extent defined in this document. It shall be required to provide Qualification according to the above referenced specification. The environments and testing shall be in accordance with the following table.

Conditions Section of RTCA/DO-160D

Category Required

Description of Tests to be Conducted for Qualification

Temperature and Altitude

4.0 through to 4.6.3

A1 Equipment Temperature range within Cabin assumed to be 0 to 55deg C at minimum Cabin Pressure equivalent to 4600m (170 kPa). It is possible that the Antenna could see a more severe range if it is not contained within the Cabin TBD. Emergency depressurisation is not required

Temperature Variation 5.0 C Internal Cabin. The Antenna may see more sevre environment if outside the Cabin (Category A)

Humidity 6.0 A Operational Shock and Crash Safety

7.0 A No requirement for Crash Safety testing is assumed

Vibration 8.0 Standard Vibration –

Sine or Random

Shock Analysis or test may be applicable

See 5.2.2.2 of this specification

Explosion 9.0 Not applicable

Waterproofness 10.0 Not applicable




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Fluids susceptibility 11.0 Not applicable

Sand and Dust 12.0 Not applicable

Fungus 13.0 Not applicable

Salt spray 14.0 Not applicable

Magnetic Field 15.0 Not applicable

Power Input 16.0 Not applicable

Preferred approach is that a regulated supply is already provided onboard so this test will have been covered by the Certification of the regulator which will be part of the aircraft

Voltage spike 17.0 Not applicable

Preferred approach is that a regulated supply is already provided onboard so this test will have been covered by the Certification of the regulator which will be part of the aircraft

Audio Frequency Susceptibility

18.0 Not applicable

This is associated with ac power supplies which will not be used

Induced Signal Susceptibility

19.0 Not applicable

Covered by design of the equipment and will not influence the aircraft

Radio Frequency Susceptibility

20.0 S Reduced level

Radio Frequency Emissions

21.0 U

Lightning Induced Transient Susceptibility

22.0 TBC May need to demonstrate survival under these conditions. But not performance

Lightning Direct Effects

23.0 Not applicable

Icing 24.0 Electrostatic Discharge

25.0

Other Tests 5.1.3 Certification for Airworthiness

The PADS shall be qualified for aircraft operations. The equipment shall be designed to meet the defined environments and a Environment Qualification Form provided as evidence according to the requirements of RTCA/DO-160D. The following is a typical requirement for meeting the airworthiness requirements. In the event the actual requirements will need to be agreed with the airworthiness authority who may require some adjustments to the requirements as defined in this specification. However, the stated requirements shall be used for the purposes of the proposed project.




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5.2 Testing

5.2.1 Test Conditions

5.2.1.1 Ambient Test Conditions

1. Unless otherwise specified in the applicable test, all test shall be performed under ambient conditions, which are:

a. Temperature +22±3 deg C b. Pressure 96 to 103.5 kPa c. Relative Humidity 35 to 60% 5.2.1.2 Tolerance of Test Parameters

The maximum allowable tolerance for test parameters shall be as follows unless otherwise specified within this specification. The tolerance specifies the allowable range within which the specified test level may vary and is exclusive of instrument accuracy. a. Temperature Maximum -0, +3 deg C Minimum -3, 0 deg C Tolerance between

Max and Min ±3 deg C

b. Barometric

Pressure

>=100Pa ±10% <100Pa and >0.04Pa ±50% <0.04 Pa ±80% c. Relative

Humidity ±5%

d. Acceleration ±10% e. Vibration

Frequency ±2% or ±1Hz

whichever id the least

f. Sinusoidal Amplitude

±10%

g. Sweep Rate ±5%

h. Random Vibration Acceleration

Power Spectral Density

20 to 500Hz ±1.5 dB

>500Hz ±3.0 dB Random

overall g rms ±1%

i. Force ±5% j. Weight ±1 g k. Dimensions ±0.123 mm l. COG

measurements ±0.5 mm

m. MOI measurements

±0.05kgm2




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n. Time ±3% 5.2.1.3 Temperature Stabilisation

The rate of change of temperature shall not change by more than 5 deg/ minute. 5.2.2 Testing for Airworthiness

5.2.2.1 Vibration

5.2.2.1.1 Fixture

The transmission and cross-talk between axes of the test fixture when loaded by the test article shall be determined prior to the formal tests. The transmission amplification shall not exceed ±3 dB and the cross-talk between axes shall not exceed the input. 5.2.2.1.2 Sine Survey

The objective of this test is to verify that no unexpected resonance’s have occurred due to design or issues associated with manufacture and quality. The conditions for the test shall be as follows: a. Amplitude 0.5g zero to peak b. Frequency Range 5 Hz to 2 kHz c. Sweep Rate 2 octaves/ minute d. Direction X, Y and Z axes e. Equipment state Non-operating 5.2.2.1.3 Random Vibration

5.2.2.1.3.1 General

The standard vibration tests, as defined in RTCA/DO-160D, shall demonstrate that equipment will meet its functional performance requirements during normal operating conditions of the aircraft. 5.2.2.1.3.2 Procedure for The Random Vibration Test

In each of the equipment's three orthogonal axes, perform the following test sequence:

a. Performance a sine survey

b. With the equipment operating, apply the appropriate performance level test Acceleration Power Spectral Density (APSD) for a minimum of one-hour-per-axis to DETERMINE COMPLIANCE WITH APPLICABLE EQUIPMENT PERFORMANCE STANDARDS. During this vibration period, also perform an APSD analysis of the vibration response measurements on the equipment. The levels shall be as defined in Figure 5-1.

c. Repeat the sine survey. Any changes in vibration resonant frequencies shall be noted on the environmental Qualification form . At the completion of the test, the equipment shall be inspected and shall show no evidence of structural failure of any internal or external component.




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Figure 5-1 Standard Random Vibration Test Curves for Equipment installed in Fixed Wing Aircraft with Turbojet of Turbofan Engines

5.2.2.2 Shock Tests

5.2.2.2.1 Purpose of the Tests

The operational shock test verifies that the equipment will continue to function within performance standards after exposure to shocks experienced during normal aircraft operations. These shocks may occur during taxiing, landing or when the aircraft encounters sudden gusts in flight. This test applies to all equipment installed on fixed- wing aircraft and helicopters. Two operational shock test curves are provided; a standard 11 msec pulse and a low frequency 20 msec pulse. The 20 msec shock is applicable to very rough runways and will not be relevant to this mission. The 11 msec shock level shall be applied as the equipment is regarded as Category A.

Appropriate Level is Level




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5.2.2.2.2 Test Procedure

Secure the equipment to a shock table by means of a rigid test fixture and mounting means intended for use in service installations. The mounting of the equipment should include those non-structural connections that are a normal part of the installation. The accelerometer used to measure or control the input shock pulse shall be placed as close as practicable to an equipment attachment point. With the equipment operating and with its temperature stabilized, apply to the test item three shocks in each orientation having a terminal saw-tooth wave shape with an acceleration peak value of six (6) g’s. The nominal pulse duration shall be 11 ms for standard shock testing. The characteristics of instrumentation used to demonstrate compliance and the shock pulse tolerance limits are shown in Figure 5-2and Figure 5-3, respectively. An equivalent shock response spectrum may replace the terminal saw-tooth wave shape. After application of the shocks, DETERMINE COMPLIANCE WITH APPLICABLE EOUIPMENT PERFORMANCE SPECIFICATION. When using a conventional drop shock machine, the equipment shall be shock tested in the following orientations:

a. Normal upright. b. Suspended upside down. C. At orientations such that the first major orthogonal axis of the equipment successively forms angles of +90 degrees and -90 degrees (two orientations) with the plane of the table. d. At orientations such that the second major orthogonal axis of the equipment successively forms angles of +90 degrees and -90 degrees (two orientations) with the plane of the table.

Figure 5-2 Shock Measuring system Frequency Response




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Figure 5-3 Terminal Saw-tooth Shock Pulse Configuration and its Tolerance Limits

5.2.2.3 Temperature and Altitude Testing

5.2.2.3.1 Purpose of Test

These tests determine the performance characteristics of equipment at the applicable categories for the temperatures and altitudes. 5.2.2.3.2 Category

The Category shall be A1 Equipment intended for installation in a controlled temperature and pressurized location, on an aircraft within which pressures are normally no lower than the altitude equivalent of 15,000 ft (4,600 m) Mean Sea

Level (MSL), is identified as Category A1 . This category may also be applicable to equipment installed in temperature controlled but unpressurized locations on an aircraft that operates at altitudes no higher than 15,000 ft (4,600 m) MSL.




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5.2.2.3.3 Test Procedure

The temperature test has been designed to combine several individual tests. The tests include:

• Cold storage • Cold switch-on • Cold Operation • Hot storage • Hot operation • Temperature variation • Cooling failure

5.2.2.3.4 Thermal Tests

Thermal environment tests shall be carried out under following conditions and those as defined below:

1. Pressure: Atmospheric ambient 2. Relative Humidity: (purge with dry Nitrogen for cold)

Ambient

High-Operational

High Storage-non-operational

Low -Operational

Low StorageNon operational

Cooling Failure Test

10 deg/Min

10 deg/Min

10 deg/Min

operational

10 deg/Min

operational

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab

Stab3hrs

Non-op

Stab3hrs

Non-op

1 hroperational

ColdSwitch-on

3hrsoperational

3hrsoperational

10 deg/minoperational

A

B C

D E F

G H

J K L

M N P

Q R S

T U V30mins

Ambient

High-Operational

High Storage-non-operational

Low -Operational

Low StorageNon operational

Cooling Failure Test

10 deg/Min

10 deg/Min

10 deg/Min

operational

10 deg/Min

operational

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab3hrs

Non-op

Stab

Stab3hrs

Non-op

Stab3hrs

Non-op

1 hroperational

ColdSwitch-on

3hrsoperational

3hrsoperational

10 deg/minoperational

A

B C

D E F

G H

J K L

M N P

Q R S

T U V30mins

Figure 5-4 Thermal Test Cycle Profile

Figure 5-4 defines the temperature testing profile. 5.2.2.3.5 Altitude Test

Conduct this test at ambient temperature. Operate the equipment at maximum duty cycle. Decrease the pressure in the test chamber to the appropriate maximum operating altitude specified in m. This will be 12000m (TBC) Allow the equipment temperature to stabilize. Maintain this pressure for at least two hours. Determine compliance of the equipment performance during the two-hour period or at the maximum duty cycle, whichever is longer. The test profile is shown graphically in Figure 5-5




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Figure 5-5 Altitude Test

5.2.2.3.6 Decompression Test

This test is not required as the equipment does not have to perform correctly during an emergency descent. 5.2.2.3.7 Power Supply Test

The electrical supplies used will be special to the Sensor and will be separate from the Aircraft systems. Hence, these tests will be carried out as part of the functional testing of the Sensor. 5.2.2.3.8 Voltage Spikes

As per 5.2.2.3.7 5.2.2.3.9 Audio Frequency Audio Susceptibility

As per 5.2.2.3.7 5.2.2.3.10 Induced Frequency Susceptibility

As the system is not part of the aircraft system and the harnesses will be separate from the aircraft system this test is not required. 5.2.2.3.11 Radio Frequency Susceptibility

This should not be high as the sensor will be placed in the cabin area of an all metal aircraft. The test levels are TBD




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5.2.2.3.12 Emission of Radio Frequency Energy

The Equipment is not designed to transmit. The modules will be contained in screened housings so any EMC testing will be carried out at Sensor level. 5.2.2.3.13 Lightning induced Transient Susceptibility

This is not relevant to this equipment 5.2.2.3.14 Lightning Direct Effects

This is not relevant to this equipment 5.2.2.3.15 Electrostatic Discharge

This is not relevant for airworthiness and will be addressed at equipment level 5.2.2.3.16 EMC Tests

The EMC requirements shall be as defined in the Mil-Std-461E ( AD1) and according to the Table 5-1.

Table 5-1 EMC Applicability Matrix

Requirement ApplicabilityEquipment/

Subsystems installed on aircraft

A L A S S A A A A A Where A is applicable L is limited according to the Mil-Std-461E S is specific to project needs TBD 6.




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INTENTIONALLY BLANK




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DOCUMENT CHANGE DETAILS

ISSUE CHANGE AUTHORITY CLASS RELEVANT INFORMATION/INSTRUCTIONS

A II First issue

B II Minor amendments

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