emits.sso.esa.intemits.sso.esa.int/emits-doc/SENER/Copia.pdf · ASTRIUM Ltd and SENER Ingeniería y Sistemas S.A. – Tres Cantos 2012 The information contained in this document is

ASTRIUM Ltd and SENER Ingeniería y Sistemas S.A. – Tres Cantos 2012

The information contained in this document is confidential and restricted, and is to be used only for the purposes established in the document. No modification, exploitation, reproduction, communication to any third party, dissemination or distribution of the whole or any part of the document is permitted without the prior written consent of ASTRIUM Ltd and SENER Ingeniería y Sistemas, S.A.. Failure to respond to any request for such consent shall in no way be construed as authorization for use.

Solar Orbiter Antennas Subsystem

Low Gain Antenna Requirement Specification

Doc. No: SOL-B.SEN-RS-00004 Revision: 1 Date: 12-06-07 DRD/DRL ref.: EN-3 CI Number: 252 100 Contract No: SOL.S.ASTR.CON.00002 SENER Doc.: Doc-01541443

Name Signature & Date

Written Astrium Ltd & I. Melgar

Checked J. Ruiz de Gopegui

Accepted J. Zabala

Authorised I. Melgar

File: SOL-B.SEN-RS-00004 rev1 Low Gain Antenna Requirement Specification.doc Pages: 36

Document approved electronically through SENER documentation management system

Doc.No.: SOL-B.SEN-RS-00004 Solar Orbiter Revision: 1 Date: 12-06-07 Antennas Subsystem Page: 2


DOCUMENT CHANGE RECORD

REVISION _________

DATE

________

SECTION/PARAGRAPH AFFECTED

_______________________

INITIATING DOCUMENT/ REASON OF CHANGE/REMARKS

______________________________

0 12-05-24 Initiating document

1 12-06-07 LGA-190 and LGA-231 modified. LGA-574, LGA-575 and LGA-576

removed.



TABLE OF CONTENTS 1. INTRODUCTION ................................................................................................................. 5

1.1 Background .................................................................................................................. 5 1.2 Scope of Document ...................................................................................................... 5 1.3 Priority of Requirements ............................................................................................... 5 1.4 Priority of Documents ................................................................................................... 5 1.5 Format of Requirements .............................................................................................. 6

2. APPLICABLE AND REFERENCE DOCUMENTS .............................................................. 7 3. MISSION DESCRIPTION .................................................................................................... 8

3.1 Scientific Objectives ..................................................................................................... 8 3.2 System Description ...................................................................................................... 8

3.2.1 System elements ................................................................................................... 8 3.2.2 Subsystem Overview ............................................................................................. 9

3.3 Mission Design ........................................................................................................... 10 3.3.1 Mission Overview ................................................................................................ 10 3.3.2 Mission Phases ................................................................................................... 10

3.4 Abbreviations ............................................................................................................. 11 4. REQUIREMENTS ............................................................................................................. 12

4.1 LGA Functional Requirements ................................................................................... 12 4.1.1 General................................................................................................................ 12 4.1.2 Operating Frequency ........................................................................................... 12

4.1.2.1 Frequency Bandwidth ................................................................................ 12 4.1.3 Antenna Performance ......................................................................................... 12

4.1.3.1 Forward Gain ............................................................................................. 12 4.1.3.2 Backward Gain ........................................................................................... 13 4.1.3.3 Internal Loss .............................................................................................. 13 4.1.3.4 Gain Variation ............................................................................................ 13

4.1.4 Axial Ratio ........................................................................................................... 13 4.1.5 Polarisation ......................................................................................................... 13 4.1.6 Azimuth Ripple .................................................................................................... 13 4.1.7 VSWR .................................................................................................................. 13 4.1.8 Phase Centre ...................................................................................................... 14 4.1.9 RF Radiation Characteristic and Power Handling ............................................... 14

4.1.9.1 Multipaction and Gas Discharge ................................................................ 14 4.1.10 Thermal Protection .............................................................................................. 15 4.1.11 Anechoic Test Cap .............................................................................................. 15

4.2 LGA Operational Requirements ................................................................................. 16 4.2.1 Physical ............................................................................................................... 16

4.2.1.1 Solar Orbiter Physical Reference Frame ................................................... 16 4.2.2 Modes of Operation ............................................................................................. 16

4.3 Design Requirements ................................................................................................. 18 4.3.1 LGA Mechanical Design and Interface Requirements ........................................ 18

4.3.1.1 LGA Design Loads and Stiffness ............................................................... 18 4.3.1.2 LGA Shock and Vibration Requirements ................................................... 19

4.3.2 Mechanical Interface ........................................................................................... 19 4.3.3 Envelope Requirements ...................................................................................... 20 4.3.4 Thermal Environment .......................................................................................... 20

4.3.4.1 Overview .................................................................................................... 20 4.3.4.2 On Ground ................................................................................................. 20 4.3.4.3 Pre-Launch Phase ..................................................................................... 20



4.3.4.4 Launch and Early Orbit Phase (LEOP) ...................................................... 20 4.3.4.5 Flight Phases ............................................................................................. 21 4.3.4.6 Spacecraft Thermal Environment .............................................................. 21 4.3.4.7 External Solar Environment Loads ............................................................ 25

4.3.5 Thermal Design Requirements ............................................................................ 25 4.3.5.1 Heat Flow Requirements ........................................................................... 25 4.3.5.2 Shadow Cone ............................................................................................ 26 4.3.5.3 Temperature Uncertainties ........................................................................ 28 4.3.5.4 Monitoring & Control .................................................................................. 28 4.3.5.5 Multi Layer Insulation Requirements .......................................................... 28 4.3.5.6 Material Degradation Requirements .......................................................... 28 4.3.5.7 Thermistor and Heater Interface Requirements ......................................... 28

4.3.6 Thermal Verification ............................................................................................ 28 4.3.7 Mass .................................................................................................................... 29 4.3.8 Electrical Requirements ...................................................................................... 29

4.3.8.1 Power Consumption ................................................................................... 29 4.3.9 Software Requirements ....................................................................................... 29 4.3.10 Reliability, Availability, and Maintainability .......................................................... 29

4.3.10.1 Assessed Reliability ................................................................................... 29 4.3.10.2 Maintainability ............................................................................................ 29 4.3.10.3 Useful Life .................................................................................................. 29

4.4 Cleanliness and Contamination ................................................................................. 30 4.5 Depressurisation ........................................................................................................ 31 4.6 Radiation Environment ............................................................................................... 31 4.7 EMC ........................................................................................................................... 31

4.7.1 Charge Protection ............................................................................................... 32 4.8 Electrical Interface Requirements .............................................................................. 32

4.8.1 Waveguide Type ................................................................................................. 32 4.9 Ground Support Equipment Requirements ................................................................ 33

4.9.1 MGSE .................................................................................................................. 33 4.9.2 EGSE .................................................................................................................. 33

4.10 Model Requirements ................................................................................................. 33 4.10.1 Mechanical Mathematical Model Requirements .................................................. 33 4.10.2 Thermal Mathematical Model Requirements ....................................................... 33

5. GDIR APPLICABILITY MATRIX ........................................................................................ 34 6. VERIFICATION REQUIREMENTS ................................................................................... 35 7. APPENDIX A ..................................................................................................................... 36

7.1 Interface Requirement Document IRD ....................................................................... 36



1. INTRODUCTION

1.1 Background The ESA Solar Orbiter is an interdisciplinary mission to the Sun. It consists of a single spacecraft which will orbit the Sun in a moderately elliptical orbit, using a suite of advanced Remote-Sensing and In-Situ instruments to perform a detailed observation of the Sun and surrounding space. The mission will revolutionise our understanding of the various processes and mechanisms which make up the Sun and Heliosphere.

The main stakeholders are:

•The European Space Agency (ESA) who are the end-customer for the spacecraft. Solar Orbiter is planned as part of ESA Science Directorate’s (D/SRE) “Cosmic Vision” programme. ESA will also operate the spacecraft from their facility at ESOC, Darmstadt

•The scientific community, in particular the Principal Investigators (PIs) who are responsible for the development and delivery of the scientific instruments on-board the spacecraft.

•Astrium Ltd. who is the Solar Orbiter Prime Contractor and the direct customer for the equipment.

Due to programmatic and technical considerations, the Prime Contractor has identified a number of equipments which are expected to be reused from the BepiColombo programme, whose mission has certain similarities to that of Solar Orbiter. Any requirements for such "Bepi reuse" will be clearly stated in this document. The supplier will be expected to procure the units, based on an Equipment Specification to be free-issued by Prime, and to accommodate the Bepi component(s) in his equipment design and verification activities.

1.2 Scope of Document This document comprises the contractual technical requirements and constraints for the Low Gain Antenna (LGA) of the Solar Orbiter Project. This includes the functional and performance requirements as well as design and interface requirements. Programmatic requirements including testing and verification requirements, and deliverables are contained in the Statement of Work, while Product Assurance Requirements are covered in a specific document of that name.

The Solar Orbiter General Design and Interface Requirements (GDIR) document ([AD 06]) outlines the general requirements for equipments, including design, environment, interfaces, and testing. Where the GDIR is an applicable document in Section 2, the subcontractor shall be responsible for flowing down and allocating the correct GDIR requirements to the equipments that he will be specifying.

1.3 Priority of Requirements In this document, the verb form used in a technical requirement shall have the following specific meanings:

•The verbal form “shall” shall be used whenever a provision is a requirement.

•The verbal form "will" shall be used to cite things that the wider system is to provide to the equipment capability being specified.

•The verbal form “should” shall be used whenever a provision is a recommendation.

•The verbal form “may” shall be used whenever a provision is a permission.

•The verbal form “can” shall be used to indicate possibility or capability

1.4 Priority of Documents Within this equipment specification, Sections 1, 2 and 3 contain general information, document references, definitions and background material, and Section 4 contains requirements. In case of



contradiction or conflict between the content of this document and any technical annex or secondary document called herein e.g. ICD or GDIR, this document shall have precedence.

1.5 Format of Requirements This document has been generated in the DOORS tool, and extracted for publication. Requirements within this document are shown in an italic font. Information within this document is shown in normal font. Each requirement is preceded by a summary line that contains the following fields, delimited by "/".

•<DOORS Requirement Number> SUBSYS-xyz. This is a unique number, assigned by DOORS.

•<Created From> Shows parent requirement. This will either be one of the System Requirement Specifications ("SSRS-xxxx" or "OIRD-YYYY") or else "Created", in which case the requirement has emerged from the system design.

•<Intended Verification Method> "T"= Test, "A" = Analysis, "I" = Inspection, "R" = Review of Design.

All document elements, which are not presented in the format explained above are not requirements and will not be verified or tracked.

There is no difference between requirements in black text or in blue text.

Red text is used to identify optional definition of requirements to be covered by Tenderers as Tender options, as requested Option 2 of the LGA ITT Cover Letter.



2. APPLICABLE AND REFERENCE DOCUMENTS As defined in Section 2 of the Low Gain Antennas Statement of Work (SOL-B.SEN-SOW-00001).

Along the document, the identifiers [AD xx] used for applicable documents and [AS xx] used for applicable standards correspond to the tables contained in Section 2 of the LGA Statement of Work (SOL-B.SEN-SOW-00001).



3. MISSION DESCRIPTION

3.1 Scientific Objectives The Sun's atmosphere and the heliosphere represent uniquely accessible domains of space, where fundamental physical processes common to solar, astrophysical and laboratory plasmas can be studied in detail and under conditions impossible to reproduce on Earth or to study from astronomical distances.

The results from missions such as Helios, Ulysses, Yohkoh, SOHO, TRACE and RHESSI have advanced significantly our understanding of the solar corona, the associated solar wind and the three-dimensional heliosphere. Further progress is to be expected with the launch of STEREO, Solar-B, and the first of NASA’s Living With a Star (LWS) missions, the Solar Dynamics Observatory (SDO). Each of these missions has a specific focus, being part of an overall strategy of coordinated solar and heliospheric research.

An important element of this strategy, however, has yet to be implemented. We have reached the point where further in-situ measurements, now much closer to the Sun, together with high-resolution imaging and spectroscopy from a near-Sun and out-of-ecliptic perspective, promise to bring about major breakthroughs in solar and heliospheric physics. The Solar Orbiter will, through a novel orbital design and an advanced suite of scientific instruments, provide the required observations.

The unique mission profile of Solar Orbiter will, for the first time, make it possible to:

•Explore the uncharted innermost regions of our solar system;

•Study the Sun from close-up;

•Fly by the Sun tuned to its rotation and examine the solar surface and the space above from a quasi co-rotating vantage point;

•Provide images and spectral observations of the Sun’s polar regions from out of the ecliptic.

Within the framework of the global strategy outlined above, the top-level scientific goals of the Solar Orbiter mission are to:

•Determine the properties, dynamics and interactions of plasma, fields and particles in the near-Sun heliosphere;

•Investigate the links between the solar surface, corona and inner heliosphere;

•Explore, at all latitudes, the energetics, dynamics and fine-scale structure of the Sun’s magnetized atmosphere;

•Probe the solar dynamo by observing the Sun’s high-latitude field, flows and seismic waves.

3.2 System Description

3.2.1 System elements The Solar Orbiter Mission consists of the following elements:

• The Space Segment, consisting of the Solar Orbiter spacecraft and its payload instruments, as follows:

• The Payload (P/L) which is composed of two suites of instruments:

o In-situ instruments

o Remote Sensing instruments.

Each of the instruments is to be provided by a Principal Investigator based on national funding through

an institute or an international group of institutes (see next section for further definition of the payload).



•The Spacecraft (SC), procured by ESA, and supplied by Astrium Ltd, that provides:

o Structural support

o Thermal control including heat shield

o Solar power generation

o Power conditioning and control

o Telecommunications

o Data handling

o Attitude and orbit control.

•The Launch Service Segment

•The Ground Segment.

3.2.2 Subsystem Overview The Low gain Antenna LGA will be used in the Telemetry, Tracking and Command (TT&C) subsystem of the Solar Orbiter satellite, which operates in X-band.

The fundamental functions of the LGA are:

• To receive X-band uplink signals for telecommand purposes.

•To transmit X-band downlink telemetry data signals.

The block diagram of the Solar Orbiter communication subsystem is shown in Figure 3.2:1 with the HGA, Medium Gain Antenna and two Low Gain Antennas.



X-MGA

X-LGA 1

X-LGA 2

MGAPM(1-axis)

Drv

A+B

(1-a

xis)

X-HGA

X-Bd Feed

HGAPM(2-axis)

Drv

A+B

(2-a

xis)

PYR (A+B) HDs

HDs PYR (A+B)

Status (A+B)

Status (A+B)

APME

DC

/DC

DC

/DC RT

N R

RT

N R

HGA Safe Pos.

MGA Safe Pos.

X-TC (Test)

X-TC (Test)

DST-1

X-B

and

Tran

s.A

LCL DC/DC

X-B

and

Rec

eive

r/D

emod

A

FCL DC/DC

RTNR

X-TM 1AX-TM 1B

X-TC 1AX-TC 1B

Discrete TM/TC

DST-2

X-B

and

Tran

s.B

LCL DC/DC

X-B

and

Rec

eive

r/D

emod

B

FCL DC/DC

RTNR

X-TM 2AX-TM 2B

X-TC 2AX-TC 2B

Discrete TM/TC

ZZ

C(2)C(2)

C(2)C(2)

C(2)C(2)

C(2)C(2)

Serial Packet TM/TC

Discrete TM/TC

MIL-1553B Bus

Pyro

DC Power

RFDA

SWITCHING CONTROLTC/TM

TC/TM

3 dB Coupler

X-TWTA-2

EPC

X-TWTA-1

EPC

LCL(B)

TC/TM

LCL(A)

TC/TM

Communication Sub System

Figure 3.2:1: Block diagram of Communication Sub System

3.3 Mission Design

3.3.1 Mission Overview The system baseline envisages the launch of the spacecraft on an Atlas IV series 401 from KSC. The nominal launch window is in January 2017 and is driven by the requirement to rendezvous with Venus as the first step in a series of gravity assist manoeuvres to bring the SC to the operational orbit around the Sun.

The duration of the entire mission (up until the end of the Extended Mission Phase) will be approximately 10 years.

The mission design is such that the transfer is entirely ballistic, with no deterministic component required. Accordingly, the only delta V requirements are those associated with initial launch error correction and pre/post GAM corrections.

3.3.2 Mission Phases The main phases during ground based integration and testing at System Level are:



•AIT related life events

•Transportation to launching area;

•Conditioning and tests;

•Installation on launcher;

•Pre-launch phase.

The main flight phases are:

• Launch and Early Operations Phase (LEOP): The period starting from the launch event through to the successful completion of the launcher error correction TCM and acquisition of nominal attitude.

• Near Earth Commissioning Phase (NECP): The period starting with the end of the LEOP, and lasting until no longer than T0 + 2 months.

• Cruise Phase (CP): The period between the end of the NECP (T0 + 2 months) and the entry of the Spacecraft into the initial operational orbit with the targeted pericenter altitude at VGAM-2.

• Nominal Mission Phase (NMP): The period between VGAM-2 and the first set of three science windows after VGAM-4.

• Extended Mission Phase (EMP): The period between the end of the first 3 Science Windows after VGAM-4 and the first set of three science windows after VGAM-6.

3.4 Abbreviations The acronyms used in this document are listed in [AD 10].



4. REQUIREMENTS

4.1 LGA Functional Requirements

4.1.1 General LGA-79 / Derived from TTC-139 / R

All performance characteristics defined in this section as a worst case value shall be met through the lifetime of the spacecraft, environmental conditions, interfaces and shall include measurement accuracy.

In addition to the requirements specified in this document the LGA shall be designed to meet the relevant sections of the GDIR [AD 06].

LGA-143 / TTC-1746 / R

Each LGA shall comprise as a minimum:

•Radiating element (e.g. helix or horn).

•Interface provisions with support bracket that attaches the LGA to the S/C structure.

The support bracket is not part of the LGA equipment. The Tenderer shall consider that this element is Customer’s responsibility.

LGA-558 / Derived from TTC-5255 / R

Single point failures shall be removed wherever possible, where single point failure cannot be removed these will be identified and summarised in a single point Failure/Critical items list. A rationale for retention shall be provided to Astrium for approval.

4.1.2 Operating Frequency LGA-145 / TTC-132 / T,A

The operating frequency of the LGA shall operate in the X-band and is as follows:

• X-band transmit: 8400-8450 MHz.

• X-band receive: 7145 - 7190 MHz.

The exact frequency of operation is TBD

4.1.2.1 Frequency Bandwidth LGA-546 / Created / T

The bandwidth of operation for transmit and receive band shall be greater than 10.9 MHz (TBC).

4.1.3 Antenna Performance

4.1.3.1 Forward Gain LGA-81 / Derived from TTC-4552 / T,A

BASELINE definition of LGA-81 requirement (for baseline LGA Tender):

The Gain of the LGA at the receive frequency band shall be greater than -1 dBi for 0° ≤ ϕ ≤ 360° and for 0° ≤ θ ≤ 95°.

OPTIONAL definition of LGA-81 requirement (for LGA Tender Option 2):

The Gain of the LGA at the receive frequency band shall be greater than -1 dBi for 0° ≤ ϕ ≤ 360° and for 0° ≤ θ ≤ 95°.



LGA-82 / Derived from TTC-4550 / T,A

BASELINE definition of LGA-82 requirement (for baseline LGA Tender):

The Gain of the LGA at the transmit frequency band shall be: • -3.0 dBi < LGA < 6.5 dBi for 0° ≤ ϕ ≤ 360° and for 0° ≤ θ ≤ 95°.

OPTIONAL definition of LGA-82 requirement (for LGA Tender Option 2):

The Gain of the LGA at the transmit frequency band shall be: • -1.0 dBi < LGA<6.5 dBi for 0° ≤ ϕ ≤ 360° and for 0° ≤ θ ≤ 95°. • Lower than 3.5dBi for 0° ≤ ϕ ≤ 360° and for 85° ≤ θ ≤ 95°. • Lower than 4.5dBi for 0° ≤ ϕ ≤ 360° and for 0° ≤ θ ≤ 85°.

4.1.3.2 Backward Gain LGA-150 / Created / T,A

Backward Gain

The gain of the LGA at the receive and transmit frequency band, for 0° ≤ φ ≤ 360° and for 120° ≤ θ ≤ 180° shall be less than -15dBi.

4.1.3.3 Internal Loss LGA-151 / Created / R

The losses of the internal antenna feed shall be included in the antenna gain for receive and transmit frequency band.

4.1.3.4 Gain Variation LGA-154 / Created / T,A

The gain variation with frequency shall be ≤0.15dB over receive and transmit frequency range.

4.1.4 Axial Ratio LGA-158 / TTC-1430 / T,A

The axial ratio of the LGA for transmit and receive frequency range shall not exceed 4dB.

4.1.5 Polarisation LGA-148 / TTC-134 / T

The polarisation shall be RHC for receive and transmit frequencies.

4.1.6 Azimuth Ripple LGA-160 / Created / T,A

The antenna gain ripple in azimuth plane shall be less than ±1dB. This does include any space craft structure influence on the antenna radiation pattern

4.1.7 VSWR LGA-162 / Created / T,A

The VSWR shall be less than 1.22:1 over the receive and transmit frequency range.



4.1.8 Phase Centre LGA-590 / TTC-5375 / A

The knowledge of the phase centre of the LGA shall be identified on the ICD with an accuracy of tbd mm.

4.1.9 RF Radiation Characteristic and Power Handling LGA-165 / Created / T,A

The LGA shall be able to withstand and operate with maximum input power of 48.45dBm (TBC), without any degradation in performance. The power handling shall be demonstrated over the full temperature range.

4.1.9.1 Multipaction and Gas Discharge LGA-166 / Derived from TTC-5259 / R

The LGA shall operate without the presence of multipaction breakdown during all mission phases and ground integration.


The LGA shall operate without the presence of multipaction or gas discharge breakdown during all mission phases and ground integration.

LGA-537 / TTC-1213 / R

Multipaction:

The LGA shall be designed to provide a margin of at least 8 dB between the worst case operational power level and the power level at which multipaction breakdown is initiated. This requirement shall be satisfied over the complete temperature range and may be demonstrated by analysis. The worst case operational power is defined in requirement LGA-165.

Corona: The LGA shall be designed to provide a margin of at least 6 dB by analysis or 4dB by test between the worst case operational power level and the power level at which corona breakdown is initiated. This requirement shall be satisfied over the complete temperature range and shall be demonstrated by analysis and test if required. For Corona testing equipments manufacture to the same set of RF design drawings shall be considered to be part of the same batch. The worst case operational power is defined in requirement LGA-165.

LGA-538 / TTC-1214 / T

If requirement LGA-537 cannot be demonstrated by analysis, the verification by one of the following options is acceptable:

• a qualification model shall be subjected to a design verification test during which a breakdown margin of at least 6 dB shall be established.

• a batch acceptance model shall be subjected to a design verification test during which a breakdown margin of at least 4 dB shall be established.

• each flight item shall be subjected to a unit acceptance test during which a breakdown margin of at least 3 dB shall be established.

If the multipaction margin demonstrated by analysis is insufficient (8dB) as per the multipaction standard ECSS-E-20-01A [AS 04] then testing will be carried out.

Multipaction testing and Corona testing will be carried out only on the sections showing non compliance.

Corona/Ionization breakdown testing will be performed over temperature, should it be required. Alternatively a test at ambient temperature supported by analysis to show that there is adequate margin over the temperature range.



LGA-YYY

All the components and equipments of the LGA chain shall be free of any risk of Gas discharge (Corona) at the maximum specified operating RF power over the full pressure range during:

1. the depressurization of the RF components and equipments at launch environmental conditions,

2. the pressurization due to out-gassing of the spacecraft in orbit,

3. ground testing at ambient pressure, and

LGA-539 / TTC-1215 / R

Electron seeding shall be used during the verification test. The use of electronic seeding is not mandatory for a CW test.

LGA-542 / TTC-1218 / R

The design and test requirements shall apply:

•With waveforms representative of operational use applied to or generated within the equipment.

•Taking into account the venting characteristics of hollow structures where RF fields are present.

4.1.10 Thermal Protection LGA-169 / Created / R

Any necessary thermal protection hardware to protect the antenna against thermal stress shall not degrade the antenna performance and shall be provided.

4.1.11 Anechoic Test Cap LGA-171 / TTC-1760 / R

Anechoic test caps shall be provided in order to allow full power antenna functional testing in a standard clean room area and in thermal vacuum over the full thermal environment (without additional absorbers).

LGA-172 / TTC-1761 / T

The anechoic test cap shall provide the following characteristics:

•Input / output attenuation: 35.0 dB ± 5 dB.

•VSWR at test cap connector / flange: < 1.5 : 1.

•VSWR at antenna interface: < 1.2 : 1.

RF-power handling: 49 dBm, emitted by antenna.

LGA-543 / TTC-5303 / T

The RF leakage of the Antenna / anechoic test cap assembly shall ensure that a power density of 1 mW/cm² is not exceeded in a distance of 10 cm around the Antenna when transmitting 49dBm into the Antenna interface.

LGA-544 / TTC-1762 / R

The RF performances of the Test Cap shall be guaranteed over the whole specified Transmit and Receive frequencies bands as defined in Section 4.1.2, without the need of tuning or modifications.

LGA-173 / TTC-1763 / R

The test cap shall provide coaxial connectors of type SMA female.



LGA-174 / TTC-1764 / T

The repeatability of measurements using the anechoic test cap shall be better than ± 1 dB.

LGAs and Anechoic Test Caps shall have a reference mounting mark to always be interconnected in the same relative position between each other during testing at antenna and at S/C level.

4.2 LGA Operational Requirements LGA-94 / TTC-139 / R

All performance characteristics defined in this section as a worst case value shall be met through the lifetime of the spacecraft, environmental conditions, interfaces and shall include measurement accuracy.

4.2.1 Physical The Spacecraft Co-ordinate Systems are axis reference frames physically attached to the respective spacecraft.

LGA-176 / TTC-1502 / R

All reference frames shall be right-handed orthogonal triads.

4.2.1.1 Solar Orbiter Physical Reference Frame LGA-84 / TTC-1523 / R

The Solar Orbiter Physical Reference Frame (PRF) coordinate system shall be as defined in Figure 4.2:1. The +X axis is sun pointing during the nominal mission.

Figure 4.2:1:Solar Orbiter Coordinate System

4.2.2 Modes of Operation The LGAs will be mounted outside S/C as shown in the Figure 4.2:2, and Figure 4.2:3 (TBC )



LGA 1

LGA 2

+Z

+Y

LGA 1

LGA 2

+Z

+Y

+Z

+Y

Figure 4.2:2: LGA accommodation on Solar Orbiter (TBC)



Figure 4.2:3: LGA interface to MLI (TBC)

4.3 Design Requirements

4.3.1 LGA Mechanical Design and Interface Requirements LGA-588 / Derived from TTC-5373 / T,A

The LGA shall be able to withstand all mechanical static and dynamic loads encountered during its entire lifetime, including:

Manufacturing,

Handling,

Transportation,

Testing,

Launch

In orbit operations.

4.3.1.1 LGA Design Loads and Stiffness LGA-103 / TTC-5259 / T

The LGA shall be designed such that its first natural frequency is above 140 Hz (TBC) in fixed base configuration.



4.3.1.2 LGA Shock and Vibration Requirements LGA-217 / TTC-5297 / T,A

Sine Environment

The LGA shall be tested to Sine Vibration Qualification levels in accordance with the requirement as stated in Table 4.3:1, The qualification levels are applicable at the interface point with the spacecraft (all values are TBC).

Frequency[Hz] Input level5-20 ±9.9mm

20-100 16g5-20 ±18.6mm

20-100 30gsweep rate : 2 oct/min

Along boresight axis

Transverse to boresight axis

Sine Vibration Test loads (qual level)

Table 4.3:1: LGA Sine vibration Qual levels LGA-219 / TTC-5299 / T,A

Random Environment

The LGA shall be subjected to Random Vibration, Qualification levels are in accordance with requirement as stated in Table 4.3:2, (all values are TBC).

All Three Axis

Frequency[Hz] QUALI/PQ

20 - 100 +12dB/oct100-300 1.5g2/Hz300-650 -15dB/oct650-850 0.03g2/Hz850-2000 -6dB/octOverall level 21.4 grms

Random input LGA

Table 4.3:2: LGA Random Qual vibration levels LGA-221 / TTC-4555 / T,A

Shock Environment

LGA shall be shock tested in accordance with the levels defined in the GDIR GDI-4254 [AD 06]

4.3.2 Mechanical Interface LGA-214 / Created / R

The LGA shall meet the mechanical interface and envelope volume requirement specified in the IRD see appendix A



4.3.3 Envelope Requirements LGA-117 / Created / I,R

The LGA including mounting feet and connectors shall not exceed the following dimensions:

•Diameter <157mm (TBC)

•Length<325mm (TBC)

with the +Z and -Z Plane being the mounting plane of the unit, see Figure 4.2:2.

4.3.4 Thermal Environment

4.3.4.1 Overview The LGA is to be able to withstand all the nominal thermal environments encountered during its entire lifetime, whilst meeting all performance requirements relevant to the specific phases outlined below:

•Integration, transportation, testing, and storage.

•Spacecraft preparation at the launch site,

•Pre-launch phase with the spacecraft encapsulated in the launch vehicle,

•Ascent phase

•In-orbit operation phase from launcher separation until the end of the extended mission.

•Safe Mode

4.3.4.2 On Ground LGA level AIT, storage and transportation climatic conditions are covered by the GDIR [AD 06].

4.3.4.3 Pre-Launch Phase LGA-240 / Derived from TTC-4728 / A

At the launch site, the LGA flight hardware shall be compatible with the thermal climatic conditions as defined in Table 4.3:3.

Temperature Range Relative Humidity

Pre-Launch Phase 4 to 30 °C (TBC) ≤60 % (TBC)

Table 4.3:3: Thermal Environmental Conditions for the Pre-Launch Phase

4.3.4.4 Launch and Early Orbit Phase (LEOP) The LEOP duration is 7 days, and the LEOP will comprise the following principal events:

•launch

•separation

•sun acquisition

•solar array and HGA deployment

•launch error correction manoeuvre

•platform commissioning.



LGA-254 / Derived from TTC-4731 / A

The LGA shall be compatible with the launcher thermal environments as defined in Table 4.3:4, considering both when spacecraft is under the fairing and after fairing jettison.

Flux Density

Net flux density radiated by the fairing before fairing jettison ≤1000 W/m2 (TBC)

Aerothermal flux exposure after fairing jettison ≤1135 W/m2 (TBC)

Table 4.3:4: Thermal Environmental Conditions for the Launch Phase

4.3.4.5 Flight Phases LGA-279 / Derived from TTC-4744, TTC-4883, TTC-4749, TTC-4750, TTC-4755, TTC-4756, TTC-4759 / A,R

The LGA shall be compatible with the enveloping thermal environmental conditions, as specified in Table 4.3:5.

Minimum Distance Flux

Maximum Distance Flux Spacecraft Pointing

Far Sun 2140W/m2 (0.8AU) (TBC)

605W/m2 (1.5AU) (TBC)

Any with respect to Sun (TBC)

Near Sun 17500W/m2 (0.28AU) (TBC)

2130W/m2 (0.8AU) (TBC)

+Xsc Sun Pointing (TBC)

Table 4.3:5: Thermal Environmental Conditions

Solar flux at 1 AU as specified in [AS 24] (1366 ± 3 W/m2) is to be used for the LGA thermal design and analysis.

The solar flux varies by the inverse of the sun distance squared (1/r2) with respect to the solar flux value at 1 AU.

4.3.4.6 Spacecraft Thermal Environment LGA-326 / Derived from TTC-4777 / A

The LGAs shall be thermally designed using panel conductive interface temperatures as presented in Figure 4.3:1.

Min Max

Conductive Interface Temperature (°C) -50.0 (TBC) +70.0 (TBC)

Table 4.3:6: Conductive Interface Temperatures for Externally Mounted Units/Equipment LGA-337 / Derived from TTC-4804 / A

The LGAs shall be thermally designed considering radiative spacecraft interface temperatures as presented in Figure 4.3:1 (TBC).



Heat Shield

Feedthrough Inner surface

Corner Feedthrough Outer surface

Heat Shield Rear

Heat Shield Layup

Temperature, °C (1)

0.28AU, Sun

Pointing -MAX (2) 0.952AU, Sun

Pointing-MIN (2) 1.5AU, Sun

Pointing-MIN (2) Any orientation (Above 0.8AU,

including VGAM) – MAX (3) Spacecraft +X External Wall 110 -90 -108 220 when illuminated, -50

when in shadow

-X External Wall -50 -170 -175 220 when illuminated, -50 when in shadow

+Y External Wall 50 -180 -195 220 when illuminated, -50 when in shadow

-Y External Wall 50 -180 -195 220 when illuminated, -50 when in shadow



Temperature, °C (1)

0.28AU, Sun

Pointing -MAX (2) 0.952AU, Sun

Pointing-MIN (2) 1.5AU, Sun

Pointing-MIN (2) Any orientation (Above 0.8AU,

including VGAM) – MAX (3) +Z External Wall 50 -170 -180 220 when illuminated, -50

when in shadow

-Z External Wall 120 -130 -165 220 when illuminated, -50 when in shadow

Heatshield & Feedthroughs Heat Shield Rear Surface 130 -70 -100 240

HIS Feedthrough 280 0 -160 50 PAS Feedthrough 280 0 -160 50 Appendages HGA - Sun side 450 70 -5 100 HGA - Anti Sun side 250 -220 -260 -80 Solar Array - Sun side 175 10 -25 100 Solar Array - Anti Sun side 140 0 -31 70

RPW Antenna 350 90 -85 170 RPW Antenna boom 350 90 -85 170 Instrument Boom MLI -180 -190 -195 190

Figure 4.3:1: Solar Orbiter Enveloping Temperatures for External Spacecraft Surfaces (TBC)

Note (1): These temperatures apply to the external surfaces only.For Structure Reference Point (SRP) temperatures, please refer to the EID-A Annex 3.

Note (2): The -Z external wall temperature stated gives the max global temperature across the wall. The progression of the HGA design may result in local hotspots based on reflections which may result in higher temperatures. This will be addressed in further iterations of this table. For information, the maximum qualification temperature of the MLI in this wall is 290°C.

Note (3): The temperatures for any orientation above 0.8AU are provided as there may be a need to orientate the spacecraft such that the heatshield is not facing the sun. This temperatures provide an enveloping environment, and will be the highest forseen for the spacecraft walls, heat shield rear surface and boom. This option is still being explored, and these values will be updated as the mission design evolves.




The LGA shall be thermally designed considering external spacecraft surface optical properties as presented in Table 4.3:7 (TBC):

Surface Region on surface ε α EOL Finish

+Z External Wall MLI 0,84 0,93 Diffuse Radiator 0,79 0,23 Specular

-Z External Wall MLI 0,84 0,93 Diffuse Radiator 0,79 0,23 Specular

+Y External Wall MLI 0,84 0,93 Diffuse Radiator 0,79 0,23 Specular

-Y External Wall MLI 0,84 0,93 Diffuse Radiator 0,79 0,23 Specular

+X External Wall MLI 0,05 0,14 Specular -X External Wall MLI 0,84 0,93 Diffuse HGA Coating 0,8 0,5 Diffuse Heat Shield Panel Rear - overhang Low Emissivity Finish 0,05 0,14 Specular Heat Shield Panel Rear - centre OSR 0,79 0,23 Diffuse Solar Array Yoke OSR 0,79 0,23 Specular Solar array Panels Cells/OSR 0,794 0,595 Specular Instrument Boom MLI 0,84 0,93 Diffuse HIS & PAS Feedthrough - outer surface Black Keplacoat 0,8 0,95 Diffuse HIS & PAS Feedthrough - inner surface Black Keplacoat 0,8 0,95 Diffuse

Note: As radiator positions are TBC, the optical properties chosen should provide a conservative environment for analysis. As the RPW is still evolving, assume black body properties. This will be updated as the design progresses.

Regions of Low and High Emissivity on Heat shield Support Panel

Spacecraft

Low emissivity f inish on

OSR

x

y

z



Solar Array pointing (View is of MY Wall):

Table 4.3:7: Solar Orbiter Optical Properties of External Spacecraft Surfaces (TBC)

The operating and non operating design temperature ranges are defined in the GDIR [AD 06].

4.3.4.7 External Solar Environment Loads LGA-571 / Derived from TTC-5240 / A

The LGA shall be able to survive direct sun exposure at steady state conditions for a sun distance less than 0.8 AU (TBC).


The LGA shall be able to survive with no external thermal loads at steady state conditions for a sun distance greater than 0.8 AU (TBC).


The LGA shall be compatible with a peak reflective solar intensity of 2500 W/m2 (TBC), at steady state conditions for a sun distance less than 0.8 AU.

LGA-573 result from reflections of solar flux off spacecraft appendages exposed to the sun onto the LGA.

4.3.5 Thermal Design Requirements

4.3.5.1 Heat Flow Requirements LGA-454 / Derived from TTC-4811 / T

Conductive heat flow between LGA and spacecraft shall be ≤1W (TBC).



4.3.5.2 Shadow Cone LGA-461 / Derived from TTC-4808 / A

The LGA shall be designed to withstand continuous sun exposure according to the shadow cone specified in Table 4.3:8 and Figure 4.3:2.


The LGA shall be designed to withstand transient sun exposure for 20 seconds (TBC) according to the shadow cone specified in Table 4.3:8 and Figure 4.3:2.

Description from Figure 4.3:2

Detailed description

Angle to edge of shadow cone from Nominal

direction of solar flux

Associated

Shadow Cone Requirement

Full protection from direct solar flux

Parts of spacecraft within this shadow cone are always fully

protected against direct solar flux by the Heatshield.

σ > 7.3°

Transient direct solar flux exposure

Parts of spacecraft within this part of the shadow cone may be

exposed to direct solar flux for short periods of time (20sec.

TBC).

3.3° > σ > 7.3° LGA-462

Steady state direct solar flux exposure

Parts of the spacecraft outside the shadow cone is always exposed to

direct solar flux.

σ < 3.3° LGA-461

Table 4.3:8: Shadow Cone



Figure 4.3:2: Shadow Cone Due to the size of the sun in comparison with the spacecraft and Safe Mode conditions the edge of the shadow cone is not parallel with the nominal direction of solar flux (see Figure 4.3:2).



4.3.5.3 Temperature Uncertainties LGA-489 / TTC-4820 / A

Temperature uncertainties, as defined in Section 3.2.1.41 of [AS 05], shall be analytically calculated and fully justified by sensitivity analysis for all items, according to Annex B of [AS 38].

Temperature uncertainties are to be agreed with the customer.

4.3.5.4 Monitoring & Control LGA-500 / Derived from TTC-4830 / R

The Thermal Control of the LGA shall be single-point failure tolerant.

4.3.5.5 Multi Layer Insulation Requirements LGA-506 / Derived from TTC-4279 / T,R

The LGA MLI used on the Flight Model of the LGA shall be able to withstand at least 2 times the complete set of equipment level mechanical and thermal qualification tests plus one launch campaign without the need for any refurbishment.


Removable LGA MLI items shall ensure compatibility with 10 mounting and dismounting operations of its interfaces with the spacecraft without degradation of performance.


The LGA MLI configuration shall be compatible with possible reflections from appendages of high intensity solar flux.

LGA-509 / TTC-5200 / T,R

The thermal hardware coatings shall be designed such that performance, as measured at ambient conditions on ground, are maintained in the space environment.

4.3.5.6 Material Degradation Requirements LGA-512 / Derived from TTC-4275 / A

Environmentally induced degradation of materials and their properties shall be taken into account in the LGA thermal design and analysis.

LGA-513 / TTC-4276 / T,R

Use of materials with non-stable thermo-optical properties shall be justified and assumptions about their ageing law caused by natural and induced (contamination) environment shall be substantiated by test.

4.3.5.7 Thermistor and Heater Interface Requirements LGA-515 / Derived from TTC-4850 / R

No interrnal unit thermistors shall be required for the LGA.

The LGA interfaces with the spacecraft will be monitored at spacecraft level.


No heater circuit shall be required for the LGA.

4.3.6 Thermal Verification LGA-568 / Derived from TTC-5218 / T

Conditions during LGA thermal vacuum testing shall envelope the thermal flight environment.




The LGA thermal mathematical model shall be correlated against at least two thermal vacuum test phases.

4.3.7 Mass LGA-115 / TTC-1886 / T

The X-band LGA Mass, excluding the support bracket, and including maturity margin shall be less than 0.9Kg (TBC)

4.3.8 Electrical Requirements

4.3.8.1 Power Consumption None

4.3.9 Software Requirements Not applicable

4.3.10 Reliability, Availability, and Maintainability

4.3.10.1 Assessed Reliability LGA-181 / Derived from TTC-5208 / A

The failure rate and single point failure, of the LGA meeting the performance requirements of this specification throughout the operational lifetime, shall comply with the FMECA, which shall have conducted in accordance with the flight standard documentation defining the construction of the equipment using flight approved parts, materials and processes.

The reliability for LGA shall be greater than 0.9995 over the operational life as defined in Section 4.3.10.3.

4.3.10.2 Maintainability LGA-184 / TTC-5210 / R

The LGA shall meet the requirements of this specification throughout the lifetime defined in Section 4.3.10.3 without the need for any regular maintenance.

LGA-185 / TTC-5211 / R

Fault diagnosis and accessibility for repair and rework shall be taken into consideration in the design of the LGA

4.3.10.3 Useful Life LGA-188 / TTC-5213 / R

The LGA shall meet all the requirements of this specification throughout operational life of 7.1years.


The LGA shall meet all the requirements of this specification throughout a 3.1 years extended life.

Ground storage / Life time



LGA-190 / Derived from SSRS-1241 / R

The LGA shall meet all the requirements of this specification after a minimum on ground storage in a controlled facility for up to two years or until 31st December 2018, whichever is the longer.

4.4 Cleanliness and Contamination LGA-194 / TTC-4582 / R

The Cleanliness and Contamination requirements for the LGA are as defined in:

• PA requirements for Subcontractors [AD 04]

• Cleanliness Requirement Specification [AD 05]

• GDIR [AD 06]

This includes requirements for the Handling, storage and transportation.

LGA-547 / Created / R

The LGA (and its GSE if any) shall be designed to minimize any areas of settling and accumulation of molecular and particulate contamination. Such areas shall be designed for easy access for cleaning.


The LGA shall be designed to minimize generation or emission of particles during operation of the equipment.

LGA-549 / Created / T

The LGA shall be designed, by appropriate material selection, to minimize organic and other outgassing. It shall be subjected to vacuum bake-out, at a suitable facility approved by ASU, immediately before delivery at ASU. The sub-contractor shall be responsible for arranging this and detailed instructions shall be issued to control these activities when the materials comprising the equipment, and their maximum temperature limits, have been identified. The vacuum bake-out will be monitored in real-time by means of a QCM (Quartz Crystal Monitor) until the outgassing rate reaches a significantly reduced and almost steady state ( i.e.; until the rate of change as measured by the QCM is ≤ to 15 Hz/hr/hr (Δ - Δ)* ) (TBC).

* If delta is the outgassing rate in Hz/hr (averaged over 20-minute intervals), then ' Δ - Δ ' is the change in the outgassing rate in Hz/hr/hr.


The mean level of molecular contamination of the exterior surfaces of the equipment shall not exceed 50 ng / cm2 at delivery to ASU. The actual level of surface molecular contamination attained at delivery at ASU will be determined by surface wipe testing iaw section 5.3.3 of ECSS-Q-ST-70-05C.


The mean level of particulate contamination of the exterior surfaces of the equipment shall not exceed 300 mm2 / m2 (“visibly clean”) at delivery to ASU.


As a minimum the following control, monitoring and implementation measures up to delivery at ASU shall be implemented:

• Selection of materials with very low out-gassing

• Vacuum bake-out of the equipment

• Optimised design and positioning of venting apertures

• Monitoring with PFO and molecular witness plates during equipment AIT




The contractor shall prepare a Cleanliness and Contamination Control Plan in accordance with requirement 5.1.2.2a of ECSS-Q-ST-70-01. The plan shall explicitly set out the measures to be undertaken by the contractor, and lower tier suppliers to the contractor, in order that compliance with the cleanliness requirements stated in this document can be demonstrated. It shall include a detailed contamination budget for all activities up to delivery at ASU.

4.5 Depressurisation LGA-231 / TTC-4070 / R

The LGA (including waveguide section) shall have suitable venting provision to meet the depressurisation requirements defined in the GDIR [AD 06] and in the PA requirements document [AD 04].

This includes requirements for handling, storage and transportation.

4.6 Radiation Environment LGA-195 / TTC-4586 / R

The solar orbiter radiation environment for the LGA is defined in GDIR [AD 06].

LGA-196 / TTC-4567 / T,A

The LGA shall be designed to withstand and perform to its specification when exposed to a UV flux over the mission lifetime as indicated in the Table 4.6:1.

5x10-210-100EUV

0.2100-150UV

0.6<180UV

2250180-400Near UV

Worst case Flux (W/m2)Wavelength (nm)Type

5x10-210-100EUV

0.2100-150UV

0.6<180UV

2250180-400Near UV

Worst case Flux (W/m2)Wavelength (nm)Type

Table 4.6:1: UV Flux LGA-198 / TTC-1993 / R

Micro-meteorite resistivity

The LGA shall be able to withstand the micro-meteorite environment specified in the GDIR [AD 06], throughout the mission and with a probability greater than 0.99.

4.7 EMC LGA-199 / TTC-4575 / R

The LGA shall be electromagnetically compatible, both internally and with the spacecraft, test, pre-launch, launch and operational orbital environments. This shall be achieved by demonstrating compliance as applicable with the requirements as defined in the GDIR [AD 06].

LGA-200 / TTC-4576 / T

EMC tests shall be conducted in accordance with GDIR [AD 06] and procedures defined in [AS 27].

LGA-201 / TTC-4569 / T,R

The LGA shall be designed to meet the ElectroStatic Discharge (ESD) requirements as defined in GDIR [AD 06]. Particular attention should be paid to minimise the risk of differential charging.



LGA-202 / TTC-167 / T,R

Radiated Emission

The max. radiated Emissions of the LGA impinging on spacecraft equipment shall be compatible with the limits as defined in GDIR [AD 06]. The performance shall be met at satellite level, when antennas are mounted on the spacecraft.

The magnetic moment for LGA is given below. In order to achieve the requirement it may be necessary to consider orienting the elements to achieve some level of magnetic cancellation. i.e. positioning each element at 180° with respect to the other.

LGA-205 / TTC-4579 / T,A

The magnetic moment of the LGA elements at nominal power provision shall be <30 mAm2 (TBC).


The LGA shall be demagnetisation prior to delivery.

LGA-564 / Derived from TTC-5253 / T

The three axis magnet dipole of the LGA shall be characterised by tests.

4.7.1 Charge Protection LGA-207 / TTC-4706 / T,A

In order to achieve a conductive outer surface of the spacecraft, the sheet resistance of materials of any outside surface shall be less than 5K Ohm / square (TBC) and bonded to structure with a resistance less than 10 Ohm.

LGA-232 / TTC-5201 / A

Higher surface resistivity is acceptable only if it can be demonstrated that the differential charge of the surface when exposed to space plasma is less than 1 Volt (TBC).

LGA-233 / TTC-5202 / T

To prevent internal charging, parts not exposed to external plasma and not grounded to the structure shall be limited in size and bonded to the spacecraft structure with a resistance lower than 1 MOhm.

LGA-234 / TTC-5203 / A,I

All external/internal metallic parts without area consideration (such as metallic labels, baseplates, straps, insulated electrical circuits ....) and intrinsically conductive parts (like carbon) that do not perform any electrical function shall be grounded to the main structure by a DC resistance lower than 1kOhm. Floating metallic parts are strictly prohibited without any area consideration.

LGA-235 / TTC-5204 / R

There shall be no harness dielectric exposed to space plasma environment.

4.8 Electrical Interface Requirements LGA-191 / Created / R

The interfaces shall comply to the GDIR [AD 06].

4.8.1 Waveguide Type LGA-193 / TTC-1897 / R

The LGA X-Band waveguides and elements interfaces shall be WR-112 type.



4.9 Ground Support Equipment Requirements LGA-131 / TTC-4131 / R

The equipment MGSE, EGSE, and OGSE which is deliverable to Prime for handling, transportation, integration to the spacecraft and testing as well as for any deployment test at System Level are TBD.

4.9.1 MGSE LGA-580 / TTC-5274 / R

The supplier shall ensure that any supplied MGSE conforms to [AD 16] MGSE General requirements specification.

LGA-582 / TTC-5276 / R

The MGSE shall conform to the same magnetic & EMC requirements as the flight hardware to which it will be attached.

LGA-583 / TTC-5277 / R

The ICDs for the MGSE shall be reviewed and approved by Astrium before the MGSE shall be manufactured.

4.9.2 EGSE LGA-585 / TTC-5279 / R

The supplier shall provide any EGSE that is required to test, verify or otherwise support the LGA.

LGA-586 / TTC-5280 / R

The supplier shall ensure that any supplied EGSE conforms to [AD 17] EGSE General requirements specification.

LGA-587 / TTC-5281 / R

The supplier shall ensure that in the event of the EGSE being required to support equipment activities immediately prior to launch (i.e. at the launch site) any mains powered EGSE supplied shall be capable of running from a 110V mains supply, as specified in [AD 17] section 6.2.2.

4.10 Model Requirements

4.10.1 Mechanical Mathematical Model Requirements LGA-133 / Created / R

The Mechanical models developed at equipment Level are deliverable to Prime for analysis at System Level and shall conform to [AD 12].

4.10.2 Thermal Mathematical Model Requirements LGA-530 / TTC-4838 / R

All deliverable Detailed Thermal Mathematical Models shall conform to the thermal model interchange requirements as specified in Annex A and B of [AS 05].

LGA-531 / TTC-4839 / R

All deliverable Reduced Thermal Mathematical Models shall conform to [AD 14].



5. GDIR APPLICABILITY MATRIX GDIR [AD 06] is fully applicable.



6. VERIFICATION REQUIREMENTS LGA-209 / TTC-1799 / R

Functional and performance parameters derived from above requirements shall be measured/ verified under relevant boundary conditions.

LGA-210 / TTC-1800 / T

The subcontractor shall demonstrate by means of functional and performance tests that there is no degradation caused by environmental testing.

LGA-211 / TTC-1801 / T

The sub contractor shall demonstrate the requirements in this specification are achieved by performing tests over the full environment test conditions.



7. APPENDIX A

7.1 Interface Requirement Document IRD ASU will supply the IRD of the LGA.