Mechanical Aspects of Design, Analysis and Testing of the Nanosatellite for Earth Monitoring and...

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Mechanical Aspects of Design, Analysis and Testing of theNanosatellite for Earth Monitoring and Observation

Aerosol Monitor (NEMO-AM)

by

Dumitru Diaconu

A thesis submitted in conformity with the requirementsfor the degree of Master of Applied Science

Graduate Department of Aerospace Science and EngineeringUniversity of Toronto

Copyright by Dumitru Diaconu 2014


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Mechanical Aspects of Design, Analysis and Testing of the

Nanosatellite for Earth Monitoring and Observation

Aerosol Monitor (NEMO-AM)

Dumitru Diaconu

Master of Applied Science

Graduate Department of Aerospace Science and EngineeringUniversity of Toronto

2014

Abstract

A next generation nanosatellite bus is under development at the University of Torontos Space

Flight Laboratory (SFL), and is being used for the first time in an ambitious Earth observation

mission to identify and monitor atmospheric aerosol species. The spacecraft system brings

together novel advanced designs that expand the capability envelope of nanosatellites, with

heritage SFL technology that is presently defining the state-of-the-art in microspace applications.

The work presented in this thesis pertains primarily to the development of the structural

subsystem of the Nanosatellite for Earth Monitoring and Observation Aerosol Monitor

(NEMO-AM). Described extensively are the design and analysis efforts made by the author to

validate and finalize the structural design in order to bring it to a manufacturing-ready stage.

Subsequent work to meet the mechanical requirements of ground operations during the assembly

and testing of the spacecraft is also presented.


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iii

Acknowledgments

This dissertation marks the most significant milestone in my professional development thus far.

On the subtle side, it also holds a personal symbolic value for the experiences and intellectual

framework that became part of who I am over these past two years. It is by no means easy to find

words to express gratitude for the people that played a part (be it of a positive or negative nature)

in my personal journey. The following is not an attempt to do that, but merely a collection of

thoughts.

I am profoundly grateful to my parents for their incredible dedication, unrelenting

encouragement and support, and above all, for the passion that they seeded in me from an early

age. I would have not strived for greatness the way I do, without their wisdom and compassion.

To my sister, thanks for being an indispensible source of cheer, as well as for being an excellent

confidant and good critic of anything from my eating habits to my (debatable manifestation of)

social skills. Thank you good friends, all of you, from two continents, for painting color into my

every day. Phil, you are an inspiration and a reliable source of good times. Nicolle, your

affection and care, together with everything that makes you the wonderful person that you are,

have been both motivation and reward throughout my work efforts.

For providing me with inspiration and enjoyment through their work, and for being my companythrough a wide palette of moods and mindsets, I would like to thank Thom Yorke and Sir Patrick

Stewart.

Finally, I would like to thank the SFL family. My deep gratitude goes to Dr. Robert Zee for

believing in my potential and giving me the opportunity to be part of an environment of

excellence and invaluable access to knowledge. To Dr. Simon Grocott, thank you for doing such

a thorough job of shaping my engineering judgment and maintaining the quality of my work on

an ascending curve, as well as for being an exceptional project manager throughout my research.

I would also like to thank Stephen Mauthe, Mihail Barbu and Cordell Grant, for the lessons and

advice that they had to offer. To the rest of the SFL staff and students, I am honored to have been

working alongside all of you, thank you for your friendship and support.


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iv

Table of Contents

Acknowledgments ........................................................................................................iii

Table of Contents ..........................................................................................................iv

List of Tables ................................................................................................................vii

List of Figures .............................................................................................................viii

Chapter 1. Introduction ................................................................................................ 1

Satellite Remote Sensing of Atmospheric Aerosols ............................................ 11.1

Microspace Missions at the Space Flight Laboratory .......................................... 31.2

The Nanosatellite for Earth Monitoring and Observation Aerosol Monitor1.3

(NEMO-AM) Mission ................................................................................................... 4

Thesis Outline ..................................................................................................... 51.4

Chapter 2. Structural Subsystem Design ................................................................... 6

Structural Subsystem Requirements ................................................................... 92.1

Structural Concept .............................................................................................112.2

Structural Subsystem Components ....................................................................162.3

+X Tray .......................................................................................................162.3.1

X Tray........................................................................................................172.3.2

Reaction Wheel Brackets ............................................................................182.3.3

Instrument Mounting Brackets .....................................................................192.3.4

Communication Equipment Panel ...............................................................212.3.5

+X Solar Array Panel ...................................................................................212.3.6

+/-Y Body Panels ........................................................................................222.3.7

+/-Z Body Panels and Risers ......................................................................242.3.8

X Body Panel ............................................................................................252.3.9

Closing the Structural Subsystem Design ...................................................262.3.10

Tolerance Stacking Analysis for Magnetometer Mounting .................................272.4

Wiring Harness Design ......................................................................................352.5

NEMO-AM Wiring Harness Modeling ..........................................................382.5.1

Satellite Mock-up for Wiring Harness Design Validation .............................392.5.2


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Satellite Mass Dummy .......................................................................................412.6

Design of Mass Dummies ...........................................................................422.6.1

Chapter 3. Finite Element Modeling and Analysis ....................................................45

Structural Model Considerations ........................................................................45

3.1 Finite Element Model Structure ..........................................................................493.2

Finite Element Modeling Strategies ....................................................................533.3

Tray Modeling .............................................................................................553.3.1

Solar Array Panel Modeling .........................................................................593.3.2

Cases Definition and Boundary Conditions ........................................................643.4

Simulation Objects ......................................................................................643.4.1

Boundary Conditions ...................................................................................653.4.2

Modal Analysis ............................................................................................673.4.3

Static Analysis .............................................................................................683.4.4

Thermo-elastic Analysis ..............................................................................693.4.5

Summary of Results ...........................................................................................713.5

Modal Analysis Results ...............................................................................713.5.1

Static Analysis Results ................................................................................733.5.2

Thermo-elastic Analysis Results .................................................................743.5.3

Chapter 4. Design and Employment of Mechanical Ground Support Equipment ..75

MGSE Requirements .........................................................................................764.1

Protective MGSE ................................................................................................804.2

Solar Array Protective Panels .....................................................................804.2.1

+/- Z Protective Panels ................................................................................824.2.2

+/- Y Protective Panels ................................................................................834.2.3

Star Tracker Cover ......................................................................................844.2.4

Satellite Support and Handling MGSE ...............................................................854.3

Satellite Support Rails .................................................................................854.3.1

Rail Handles ................................................................................................874.3.2

Satellite Support Frame ..............................................................................874.3.3

Positioning MGSE ..............................................................................................894.4

Positioner Interface Baseplate ....................................................................914.4.1


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Base Support Block .....................................................................................914.4.2

45Support Block ........................................................................................924.4.3

90 Support Block .......................................................................................934.4.4

Positioning MGSE Configurations ...............................................................944.4.5

Chapter 5. Conclusion .................................................................................................95

References....................................................................................................................96


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vii

List of Tables

Table 1.1Overview of SFL nanosatellite bus capabilities ............................................ 3

Table 2.1NEMO-AM Structural Subsystem Requirements ......................................... 9

Table 2.2Magnetometer PCB manufacturing tolerances ...........................................29

Table 2.3Magnetometer PCB mounting interface tolerances .....................................30

Table 2.4Magnetometer boom tolerances .................................................................31

Table 2.5Magnetometer angular errors due to mechanical tolerances ......................34

Table 2.6Mass dummy inventory ...............................................................................44

Table 3.1 Summary of results for the comparative study on honeycomb panel

modeling ........................................................................................................................63

Table 3.2Material yield stress criteria .........................................................................73

Table 3.3Summary of maximum stresses on primary structural components ............73

Table 3.4Angular displacements under on-orbit thermal loading ...............................74

Table 4.1MGSE Requirements ..................................................................................77


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List of Figures

Figure 1.1NEMO-AM Mission Patch ........................................................................... 4

Figure 2.1XPOD Duo and generic spacecraft ............................................................12

Figure 2.2NEMO-AM Concept design ........................................................................13

Figure 2.3Tray architecture in the GNB (left) and the NEMO-AM (right) ....................14

Figure 2.4External layout of NEMO-AM at Critical Design Review ............................15

Figure 2.5Internal layout of NEMO-AM at Critical Design Review .............................15

Figure 2.6+X Tray Assembly ......................................................................................16

Figure 2.7-X Tray Assembly.......................................................................................17

Figure 2.8Reaction Wheel Brackets Assembly ..........................................................19

Figure 2.9Optical Instrument Assembly .....................................................................20

Figure 2.10Communication Equipment Panel ............................................................21

Figure 2.11+X Solar Array Panel (+X side) ................................................................22

Figure 2.12+Y Panel ..................................................................................................23

Figure 2.13-Y Panel ...................................................................................................23

Figure 2.14+/-Z Panels and Risers ............................................................................24

Figure 2.15-X Panel ...................................................................................................25

Figure 2.16Finalized internal layout of the NEMO-AM ...............................................26

Figure 2.17Finalized external layout of the NEMO-AM ..............................................26

Figure 2.18Magnetometer Boom Mounting ................................................................28

Figure 2.19Magnetometer PCB dimensions (left) and reference system (right) ........28

Figure 2.20Magnetometer PCB manufacturing tolerances ........................................29

Figure 2.21Magnetometer PCB mounting interface tolerances..................................30

Figure 2.22Magnetometer boom tolerances ..............................................................31

Figure 2.23Magnetometer boom calibration reference features (green) ....................32

Figure 2.24DC-DC Converter wiring harness specification ........................................37

Figure 2.253D model of the wiring harness between the NEMO-AM trays ................39

Figure 2.26Structural mock-up components for wiring harness design validation......40

Figure 2.27Wiring harness mock-up on the PCB stack (left) and between the satellite

trays (right) .....................................................................................................................41

Figure 2.28Mass dummy of Battery assembly (left) and PCB stack (right) ................43


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Figure 2.29Mass dummies of Reaction wheels (left) and Payload (right) ..................43

Figure 2.30Fit check assembly with satellite trays and payload mass dummy ..........44

Figure 3.1Structure of FEA elements in NX 8.0 .........................................................50

Figure 3.2 Triangular discretization techniques based on the Delaunay criterion (left)

and the Advancing Front Method (right) .........................................................................57

Figure 3.3Accurate FEM of honeycomb panel ...........................................................61

Figure 3.4Composite element 2D mesh of honeycomb panel ...................................62

Figure 3.5Equivalent 3D mesh of honeycomb panel ..................................................62

Figure 3.6 Plot of deflection results for the accurate model (left), 2D mesh (middle)

and equivalent 3D mesh (right) of the honeycomb composite panel..............................63

Figure 3.7Representation of the thermal gradient field definition on the satellite FEM

(red = 35C to blue = 5C) ..............................................................................................70

Figure 3.8First mode of 181.5 Hz on the solar array panel (local mode) ...................71

Figure 3.9Second mode of 192.5 Hz driven by the payload mass .............................72

Figure 3.10Third mode of 208.9 Hz on the reaction wheel brackets ..........................72

Figure 3.11 Deflected shape of the satellite trays, payload and reaction wheel

brackets under thermal loads present during an on-orbit imaging campaign .................74

Figure 4.1+X solar array protective panel ..................................................................81

Figure 4.2-X solar array protective panel ...................................................................81

Figure 4.3Solar array protective panels in use ...........................................................81

Figure 4.4-Z (left) and +Z (right) protective panels .....................................................82

Figure 4.5+/-Z protective panels in use ......................................................................83

Figure 4.6+Y (left) andY (right) protective panels ...................................................84

Figure 4.7-Y protective panel in use ..........................................................................84

Figure 4.8Star Tracker protective cover in use ..........................................................85

Figure 4.9Satellite support rails in use .......................................................................86

Figure 4.10Rail handles in use ...................................................................................87

Figure 4.11Satellite support frame .............................................................................88

Figure 4.12Satellite support frame use for integration with the XPOD Duo (blue) .....89

Figure 4.13Standard test setup in the anechoic chamber ..........................................90

Figure 4.14RF positioner interface baseplate ............................................................91

Figure 4.15Base support block ..................................................................................92


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Figure 4.1645Support block .....................................................................................93

Figure 4.1790Support block .....................................................................................93

Figure 4.18MGSE setup to position satellite at 0eft (middle) and 90(right) tilt

angles for the RF antenna pattern test ...........................................................................94


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1

Chapter 1

Introduction

Sateite Remote Sensing of Atmospheric Aerosos1.1

The subject of accelerated global climate change triggered by human activities is no longer a

topic of dispute within the international scientific community [1][2][3]. A strong consensus has

been built in the past several years among climate scientists worldwide, agreeing that current and

past practices in fields like agriculture, industry and transportation, are the main contributors to

the climatological trends that have been observed over the past century [4]. The two primary

phenomena considered by researchers to be the drivers of the current increased rates in climate

change are global warming and global dimming. While global warming is attributed to

greenhouse gases which trap heat within Earths atmosphere, global dimming has the opposite

effect through aerosols (suspension of particulates in the atmosphere) which cause an increase in

Earths albedo and a reduction in the amount of solar radiation that reaches the surface. Because

of this coupling, the magnitudes of these two phenomena have been masked until recently.

Greenhouse gases have a relatively homogenous distribution around the globe, and a lifetime of

nearly 100 years in the atmosphere, making their characterization feasible from ground-based

facilities. By contrast, measurement of anthropogenic aerosols is considerably more difficult, due

to their short lifetime (approx. one week), which gives their distribution a heterogeneous quality

in both space and time. Under these circumstances, an opportunity arises for remote sensing

satellites to provide service through their global coverage capabilities and favorable location for

capturing and analyzing the properties of the solar radiation reflected by Earths atmosphere.


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CHAPTER 1. INTRODUCTION 2

The limited number of satellites that have been employed so far in the study of atmospheric

aerosols is primarily due to the novelty of this field of study, which has only started to move

from the exploratory phase to the global quantitative phase in the last decade [5]. Notable

missions that carried sensors dedicated to aerosol investigation are:

- ADEOS I (Advanced Earth Observing Satellite I) launched in 1996 and ADEOS II

launched in 2002, two approx. 3,500 kg satellites developed by the Japanese national

space agency JAXA, carrying POLDER (POLarization and Directionality of the Earth's

Reflectances), the first instrument designed specifically for atmospheric aerosol

identification, by the French space agency CNES (Centre National dEtudes Spatiales);

- Terra (EOS AM-1), a NASA-built 4,864 kg satellite for Earth observation, launched in

1999; among its payloads, the instruments MISR (Multi-angle Imaging

SpectroRadiometer) and MODIS (Moderate-resolution Imaging Spectroradiometer)

measure aerosol properties over continents and oceans;

- PARASOL (Polarization and Anisotropy of Reflectances for Atmospheric Sciences

coupled with Observations from a Lidar), a 120 kg microsatellite launched in 2004 by

CNES, carrying the POLDER instrument;

- CALIPSO (Cloud-Aerosol Lidar and Infrared Pathfinder Satellite Observations), a 585

kg mini-satellite jointly built by NASA and CNES and launched in 2006, carrying three

instruments dedicated to atmospheric aerosol and cloud investigations;

This list suggests the relatively large scale and inherent high costs of space missions that have

been aimed at anthropogenic aerosol detection thus far. Considering the localized properties of

aerosol masses across regions of the globe and the implicit desire for dedicated observation of

critical areas, the need emerges for more accessible remote sensing satellite missions that can

help individual nations monitor and manage the aerosol emissions above their territory.


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Microspace Missions at the Space Fight Laboratory1.2

The Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies

is a research facility that possesses end-to-end capability for space mission implementation using

nanosatellites (satellites of mass


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The Nanosateite for Earth Monitoring and Observation 1.3Aeroso Monitor NEMO-AM Mission

In the context of the global climatological shifts caused primarily by certain areas of human

activity, an increasing number of nations are making efforts to better understand and control theircontribution to these changes. With respect to aerosol concentration in the atmosphere, southern

Asia is among the regions that exhibit the most elevated levels [8] [9]. This has led to a

collaborative agreement between the Indian Space Research Organization (ISRO) and SFL, for

the development of an aerosol detecting nanosatellite. In order to address the demanding mission

requirements, the NEMO bus concept was developed. Offering leading-edge payload carrying

capability in the nanosatellite field, this bus design is first employed in the NEMO-AM (Aerosol

Monitor) mission.

The primary objective of the NEMO-AM mission is to identify aerosol species and

concentrations in the troposphere above India [10]. This goal is achieved through the

incorporation into the satellite bus of an advanced optical payload. The device is a

custom-designed multi-spectral dual-polarization instrument, which will operate by taking a

series of exposures of the target area at different illumination angles, using three

charged-coupled device (CCD) imaging sensors to simultaneously capture images in the red,

blue and near-infrared spectral bands. A team of scientists at ISRO will subsequently process the

data, using a number of algorithms to isolate and analyze the solar radiation reflected by the

atmosphere above the targeted geographical region.

Figure 1.1NEMO-AM Mission Patch


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Thesis Outine1.4

The goal of this thesis is to document the design and analysis work done for the structural

subsystem of the NEMO-AM nanosatellite and for the mechanical ground support equipment

(MGSE) to be used during the assembly and testing of the spacecraft. The majority of

mechanical components presented in the following chapters have gone through several design

iterations, driven by the evolution of the satellite subsystems and the associated revised

requirements. The author focuses on his contribution, the lessons learned and the general

techniques that can be derived and applied to similar problems.

Following this introductory chapter, which contextualizes the work of the author, Chapter 2

elaborates on the design of the structural subsystem of the NEMO-AM spacecraft. A brief

description is given of the conceptual and preliminary designs of the satellite, and a list of

requirements is provided in order to better define the parameters of the design challenge related

to the NEMO-AM mission. The bulk of this chapter consists of a systematic view of the final

versions of the structural subsystem components, along with the description of several major

mechanical design tasks completed by the author.

A presentation of the structural modeling and analysis work that was done for NEMO-AM using

the finite element method is given in Chapter 3. In order to document this segment of work while

also establishing a set of guidelines applicable to other similar problems, modeling rules and

techniques are discussed in detail. The final section of this chapter provides the structural

analysis results for a number of mechanical loading scenarios relevant to the mission.

Chapter 4 of this thesis turns to the design of mechanical ground support equipment. The first

section discusses equipment design requirements derived from the assembly, integration and

testing (AIT) plan. The subsequent sections consist of a part-by-part presentation of the MGSE

components developed by the author, providing information on design considerations and

intended modes of use.

Lastly, Chapter 5 summarizes and draws conclusions on the authors accomplishments

throughout the two-year experience of designing and building the structural subsystem and

mechanical ground support segment for the NEMO-AM spacecraft.


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Chapter 2

Structural Subsystem Design

The main function of a satellite structure is to provide mechanical support for the sum ofcomponents needed to achieve the objectives of a space mission. In the field of microspace

applications, one of the keys to mission performance is flexibility and responsiveness in the

development process. Therefore, it is commonly the case that the design of the satellite structural

subsystem will be driven by a set of semi-rigid requirements. These requirements are defined in

the initial phases of concept exploration and preliminary design, and are actively revised and

adjusted thereafter to accommodate changes in the architecture and layout of other satellite

subsystems. Despite this flexibility, design changes are not applied lightly. Trade-off studies and

design investigations are typically carried out in every instance of design revision. The outcome

of this practice is achieving an optimum balance between cost and attainable objectives within

the definition of a space mission enabled by a small satellite.

This general trait of adaptability that characterizes the requirements of a small satellite mission

was present throughout the structural design work conducted by the author. According to

common practice in the field of space mission design [11], the NEMO-AM mission development

followed a course that can be broken down into the following phases:

Identification of Needs (Phase 0) this step precedes the design process, and involves

recognizing a need in a field of human activity that could be met through a satellite application;

Conceptual Design (Phase A)targeted needs are analyzed and studies are made to propose

a feasible mission concept;


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CHAPTER 2. STRUCTURAL SUBSYSTEM DESIGN 7

- a System Requirements Review (SRR) is done at the end of this phase, releasing sets of

requirements that will drive the design of each satellite subsystem;

Preliminary Design (Phase B) the concept is expanded as satellite subsystems are

developed to a stage where they can demonstrate the capability of meeting their specific

requirements;

- a Preliminary Design Review (PDR) assesses each subsystem design at the end of this

phase;

Detailed Design (Phase C)detailed design is done for all satellite subsystems;

- a Critical Design Review (CDR) takes place before steps are made towards subsystem

component procurement and assembly;

Assembly, Integration and Testing (Phase D) satellite components are manufactured,

assembled and tested, coalescing into a flight-ready satellite;

- After the completion of this phase, the satellite is deployed by a launch vehicle (LV) into

Earth orbit;

Operations and Support (Phase E) the satellite becomes engaged in achieving the mission

objectives; this phase ends with decommissioning and/or deorbiting of the satellite.

The authors contributions to the NEMO-AM structural subsystem development commenced

shortly after CDR. Despite the significant portion of component-level design that was finalized at

this stage, initiation of the Assembly, Integration and Testing (AIT) phase was delayed by a

number of critical post-CDR decisions and revisions. Under these circumstances, efforts in

structural design and analysis were made by the author throughout the two-year research

program to meet the supplementary requirements emerging in every satellite subsystem. The

following is a list of major tasks that were accomplished, organized according to the spacecraft

segment that necessitated the additional design work:

a. Power Subsystem:

- re-distributed solar-cells and re-routed wiring on the satellite body panels;

-provided structural accommodation for an additional power-board and updated the wiring

harness design;


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b. Payload Subsystem

- generated a new design of supporting structure for the instrument computer boards;

- redesigned the payload mounting interface to match the revised instrument design;

-proposed several design concepts for optical instrument structural components;

c. Attitude Determination and Control Subsystem

- redesigned the Star Tracker mounting interface;

- relocated several panel-mounted Sun Sensors;

- redesigned the GPS antenna mounting bracket;

d. Communication Subsystem

- expanded the structural subsystem to provide accommodation for additional uplink

components (S-Band cavity filters, S-Band down-converter and co-axial two-way splitter);

- redesigned the S-Band patch antenna mounting brackets;

e. Thermal Control Subsystem

- adjusted the design of structural components in order to meet specific conductivity

requirements;

- designed spacers and mounting interfaces to improve thermal coupling or de-coupling

between various electrical components and the bus structure;

f. Satellite Deployment System

- to ensure compatibility with the SFL-built XPOD Duo (eXoadaptable PyrOless Deployer)

satellite deployment system, design adjustments were made to the solar array supporting

structure.

As each design task is considered in more detail, the progression of iterations gains complexity,

revealing inputs and drivers that frequently exert opposing tendencies on the structural design.

These common situations call for compromises between the parameters of various subsystems,

creating an optimized satellite system from subsystems which do not necessarily exhibit

maximum performance when examined individually. Interactions between the structural

subsystem and other satellite subsystems were present throughout the entire mission

development process, and the reduced size and close collaboration of the NEMO-AM

engineering team proved essential in meeting every design challenge and finding a solution.


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Structura Subsystem Requirements2.1

A number of high-level considerations were devised at the beginning of the preliminary design

stage of the NEMO-AM structural subsystem [12], and were formulated in such a manner that

they would maintain their relevance as the design grew in complexity:

- the structural subsystem will be designed to accommodate all satellite components as

outlined in the system architecture diagram;

- the design of the structure shall allow testing of modules/sub-assemblies and subsystems

independent of the rest of the satellite, including but not limited to:

o Electrical functionality

o Structural functionality

o RF functionality

o Thermal functionality

o Instrument (optical) functionality

- the satellite structure shall comply with the Polar Satellite Launch Vehicle (PSLV)

requirements as stipulated in the most recent version of the Nanosatellite Launch Service

Interface Control Document (ICD).

Table 2.1 derives and lists a detailed set of structural subsystem requirements, outlining the main

drivers for the design and analysis work that was done by the author.

Table 2.1NEMO-AM Structural Subsystem Requirements

No. Requirement Description

General Requirements

1.The spacecraft structure shall be compatible with the XPOD Duo deployment system.-Source: Fundamental characteristic based on the mission design concept and the

selected launch scenario (inside XPOD Duo attached to Antrix PSLV).

2.

The spacecraft primary structure shall have a dual-tray architecture.

-Source: The use of a GNB-like structural concept is a measure of risk and cost

reduction by utilizing previously developed and flight-proven designs.


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3.

The spacecraft structure shall provide GNB-derived mechanical interfaces for GNB

heritage electronic components, where practicable.

-Source: Measure of risk and cost reduction by utilizing previously developed and

flight-proven designs.

4.

The spacecraft structure shall be composed of materials that are compatible with a high

vacuum environment, as deemed relevant for accommodating a high-performance

optical payload. This includes a total mass loss (TML) of


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Launch Vehicle Requirements

10.

The spacecraft shall have its first natural frequency (FNF) above the frequency of

90Hz.

-Source: Antrix PSLV Nanosatellite Launch Service ICD Section 10 Environment on

the Payload.

11.

The spacecraft structure shall be designed to have a positive margin of safety against

stress under a steady-state acceleration of 11.6G in all axes.


the Payload; derived from the root sum square of the maximum longitudinal and

lateral quasi-static loads on the PSLV, and augmented by a safety factor of 1.25.

12.

The spacecraft shall be subjected to qualification and acceptance vibration tests in

accordance with the PSLV requirements. These tests will include:

o Sinusoidal Vibration Test

o Random Vibration Test

o Shock Test


the Payload.

Structura Concept2.2

The NEMO-AM structure is designed under the influence of two opposing factors, one

originating in the XPOD Duo accommodation constraints, which determine fixed spacecraft

dimensions along certain directions, and one originating in payload carrying capability, which

tends to exert an outwards push on the volume envelope.

The deployment principle used by SFL in its various XPOD designs consists of a push-out

mechanism that applies the necessary force to cause a nanosatellite contained within the volume

of the XPOD to slide along four rails and be ejected, achieving separation from the last stage of

the LV (launch vehicle). A simplified view of this system is shown inFigure 2.1.


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Figure 2.1XPOD Duo and generic spacecraft

An important structural design feature derived from the XPOD Duo compatibility requirement is

the presence of spherical ball-cup interfaces on the spacecraft rails (close-up view inFigure 2.1).

The design concept behind this type of interface has been proven in previous missions based on

the Generic Nanosatellite Bus (GNB) and is thus a reliable solution that employs heritage SFL

technology. There are a total of eight interfaces, on the top and bottom ends of the four

View A

View A

Guide rail Gaps to allow

smooth slidingXPOD Duo

Generic spacecraft

XPOD Duo main spring

Door

Pusher plateBaseplate

Ball-Cup interfaces to constrain lateral

motion between XPOD and spacecraft

Z

X

Y

ToolingBall

Spherical

Cup


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spacecraft rails. These interfaces constitute the only points of mechanical contact between the

NEMO-AM spacecraft and the XPOD Duo structure, and are thus the main load paths that

transfer both quasi-static and dynamic loads between the two integrated systems.

The initial concept for the NEMO-AM spacecraft was developed at a stage where little design

work had been completed for the optical instrument in the payload. As a result, the spacecraft

conformed to XPOD Duo accommodation requirements with no special design features apart

from the main solar array panel which, due to its large size driven by power generation

requirements, was located outside the XPOD interior envelope [13]. This early concept is shown

inFigure 2.2.

Figure 2.2NEMO-AM Concept design

Usage of past SFL designs is substantial within the NEMO-AM structural subsystem, and can be

noted both in the overall layout of the spacecraft bus, which is essentially built by merging two

GNB structures, as well as in the smaller areas such as enclosures for electronic components.

NEMO-AM

deployment from

XPOD Duo

Instrument Aperture

Main Solar Array(50 x 50 cm)

ZX Y

Z

X

Y


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For the purpose of maximizing the volume available for payload, the GNB layout was not merely

duplicated: as depicted in Figure 2.3,half of the NEMO-AM structure maintained the features

specific to GNB trays in order to accommodate bus electronics, while the other half was stripped

down to the four rails, which serve as both the primary load path to the XPOD Duo deployment

system, as well as the main structural support for the payload instrument.

Figure 2.3Tray architecture in the GNB (left) and the NEMO-AM (right)

A key design feature of the XPOD Duo, namely having two out of its four sides open, allows the

expansion of the spacecraft outside of the basic 20 x 20 x 40 cm interior envelope. The presence

of the predeployed solar array panel, featuring a surface of more than twice the lateral cross-

sectional area of the XPOD Duo interior volume, was made possible by this valuable

characteristic of the deployment system. As subsystem designs matured in the NEMO-AM

preliminary development stage, several changes were implemented in the satellite layout and

consequently in the structural subsystem. A clearer set of volume requirements for the optical

instrument and instrument computer boards was a major factor that determined an increase in the

overall size of the satellite bus. The presence of the second open side in the XPOD Duo, opposite

to the side that allows for the NEMO-AM solar array panel to protrude, was thus crucial for

enabling the expansion of the spacecraft overall dimensions. In order to accommodate three

Instrument On-Board Computers (IOBCs), the spacecraft bus volume increased its size in the X

direction. While still compatible with the XPOD Duo accommodation constraints, the NEMO-

AM volume settled at 20 x 30 x 40cm.Figure 2.4 andFigure 2.5 outline the external and internal

layouts of the satellite at the CDR stage (starting point of the authors structural design work

segment).

Payload Volume

18x20x22cm

Launch Rails

+X Tray

-X Tray

Payload

Volume

8x13x17cm

-Z Tray

+Z Tray

Z

X

Y

Z

X

Y


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Figure 2.4External layout of NEMO-AM at Critical Design Review

Figure 2.5Internal layout of NEMO-AM at Critical Design Review

Star TrackerMagnetometer Boom

-Y S-Band Antenna

Test Port

Solar Cell

Strings (8)

UHF Antenna (4)

GPS Antenna

+Y S-Band Antenna

Sun Sensor

Sun Sensor

Payload Lens

Solar CellStrings (4)

Sun Sensor

Sun Sensor

Payload Optical

Instrument

Payload Computers

Batteries

Reaction Wheels

Star Tracker

Bus Electronics

Bus Electronics


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Structura Subsystem Components2.3

The primary design features that define the spacecraft structural concept were outlined in the

previous section: a system of two trays joined by internal supporting structures and enclosed by

six body panels. A more focused view of each structural component is presented in this section.

+X Tray2.3.1

Driven in size by the two 434mm long satellite deployment rails spaced 180mm apart, the +X

tray is a major component of the NEMO-AM primary structure. Using the legacy GNB format

for its electronics accommodation section, the tray supports the majority of satellite power

subsystem components: a Battery Assembly of six cells, a Battery Charge/Discharge Regulator

(BCDR+), 2 DC-to-DC Convertors, the Array Connector Board (ACB) and the Payload Power

Board (PAYPOW).

Figure 2.6+X Tray Assembly

ACBPAYPOW

Battery

Assembly

BCDR+

DC-DC

Converters

+X Tray

Solar Array Brackets (4)

+X Sun Sensor

Enclosure

Solar Array Bracket

Stiffeners (6)

Z

X

Y


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TheY half of the tray is entirely dedicated to payload accommodation, with mounting features

provided on a reinforced section of the rails. As the rate of progress of the payload instrument

design was not as rapid as that of the satellite primary structure, the interface sections on the rails

were designed with flexibility in mind, in order to be adaptable to changes in the payload

mounting scheme.

On the +X side of the tray, four solar array brackets are attached, along with a set of stiffening

panels spanning between them, in order to provide a firm interface for the satellite solar array

panel. Mounted to one of these brackets is a Sun Sensor enclosure, which protrudes through a

rectangular cutout in the solar array in order to achieve coverage of the satellite +X field of view.

X Tray2.3.2

Featuring the same layout concept and overall dimensions as the +X tray presented above, this

second major element in the NEMO-AM primary structure is designed for modular assembly and

a customizable arrangement of electronic components.

Figure 2.7-X Tray Assembly

IOBC

Boards

-X Tray

IOBC Tray

PCB Stack Plate

PCB Stack

UHF Receiver

S-Band TransmitterZ

X

Y


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Soon after CDR, a significant change in the NEMO-AM design was implemented in order to

improve mechanical alignment between the Payload Instrument and the Star Tracker. In effect,

the Star Tracker and the Instrument On-Board Computers (IOBCs) exchanged locations within

the satellite bus, in order to allow direct mounting of the Star Tracker to the Payload Instrument

structure. With respect to the design of the X tray, the modification becomes apparent when

Figure 2.7 is compared against the internal satellite configuration presented in Figure 2.5 (see

alsoFigure 2.16 in Subsection2.3.10).

Presently, the electronics accommodation segment of the tray consists of two removable

structural parts, the PCB stack plate and the IOBC tray. During the satellite assembly process,

these components are used independently of the tray to assemble several subsystem electronic

boards. The PCB stack plate accommodates a GPS Receiver, three On-Board Computers

(OBCs), namely the House-Keeping Computer (HKC), the Attitude Determination and Control

Computer (ADCC) and the Payload On-Board Computer (POBC), as well as a legacy GNB

Power Board. For the purpose of processing data from the payload instrument CCD sensors,

three Instrument On-Board Computer boards are mounted on the IOBC tray. This design concept

makes the satellite assembly procedure more practical by allowing pre-installation and tying

down of the wiring harness between the stacked components without the spatial constraints

present subsequently within theX tray.

Communication components are also present within this subassembly, attached to the +X side of

the PCB stack plate. The two enclosures are designed to provide electromagnetic shielding to the

PCBs inside, as well as to reinforce the underlying panel by having a robust mechanical interface

between them. On the Y half of the tray, payload mounting interfaces are present on both

launch rails, in a symmetric configuration with respect to the +X tray.

Reaction Wheel Brackets2.3.3

An assembly of four structural components spans the distance between the satellite trays in the

+Y half of the NEMO-AM bus. The box-type design concept provides stiff support for three

Reaction Wheels, as well as two S-Band filters. A pair of parallel brackets (+Y and Y reaction

wheel brackets) are connected by two smaller perpendicular panels, creating orthogonal

mounting interfaces for the three wheels.


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Figure 2.8Reaction Wheel Brackets Assembly

Instrument Mounting Brackets2.3.4

A major point in the NEMO-AM development timeline was the transition from an SFL-built

optical payload to one proposed and built by a subcontractor. The authors work involved

coordinated design with the payload provider, in order to ensure that the instrument

accommodation and wire routing scheme would be compatible with the satellite bus design. A

comparative look at Figure 2.5 and Figure 2.9 reveals some differences in the overall shape of

the optical payload, both on the front lens barrel and on the rear optics and CCD sensor

assembly.

-Y Reaction Wheel

Bracket

Y Reaction Wheel

+Y Reaction

Wheel Bracket

Z Reaction

Wheel Bracket

X Reaction

Wheel Bracket

X Reaction WheelZ Reaction Wheel

S-Band Cavity Filters

Z

X

Y


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From the standpoint of the satellite structural subsystem, new components were designed to

support the optical instrument and provide suitable load paths to the four satellite rails. Due to

the large size and complex geometry of the optics assembly inside the instrument back section,

accommodation within the dedicated volume of the satellite bus is asymmetrical. As a result, two

non-symmetrical payload brackets were designed, each presenting both mounting and alignment

features at their interfaces (using socket head screws and dowel pins).

Figure 2.9Optical Instrument Assembly

Based on the design decision previously described in the X tray subsection, the Star Tracker

was relocated on an angled bracket that is mounted directly onto the Instrument main structure.

In order to improve the accuracy of the assembly, dowel pins were used on the two bracket

interfaces, to set the optical axes of the Instrument and Star Tracker as close to the nominal

design angle as possible.

+Z Payload

Bracket

Star Tracker

-Z Payload

Bracket

Star Tracker

Bracket

Instrument

Baffle

InstrumentBody

Z

X

Y


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Communication Equipment Panel2.3.5

Another post-CDR addition to the NEMO-AM structural subsystem is a panel that supports

frequency-conversion equipment for the communication uplink segment. As a consequence of

the asymmetry in the payload accommodation within the satellite bus, a suitable volume for theadded radio components was identified on the Z side of the instrument lens barrel (can be seen

in Figure 2.16 under Subsection 2.3.10). Spanning the distance between the trays, the

communications equipment panel is a 2mm thick plate reinforced by two cross-ribs, featuring

mounting points on the -Z facing side for an S-Band Down-Converter, an S-Band filter and a

Co-axial two-way splitter. Additionally, the surface available on the +Z side is used by the

Down-Converter Power Board and the Instrument Shutter Driver Board, the most recent

additions to the system architecture at the time of this writing.

Figure 2.10Communication Equipment Panel

+X Solar Array Panel2.3.6

A major feature of the NEMO-AM bus is the +X solar array, which consists structurally of a

580mm x 625mm honeycomb composite panel of 20mm thickness. The panel is attached to the

satellite +X tray through four solar array brackets (Subsection 2.3.1), using 12 M4 screws.

Thermal considerations related to the conductivity of the panel resulted in the selection of CFRP

Instrument Shutter

Driver Board

S-Band Down-Converter

Power Board

S-Band

Cavity Filter

S-Band

Down-Converter

Co-axial

two-way Splitter

Z

X

YZ

X

Y


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(Carbon Fiber Reinforced Polymer) as the panel facesheet material, while the core is a standard

Aluminum alloy 3003 honeycomb. The panel serves primarily as a substrate for 70 photovoltaic

cells, distributed in strings of seven, with eight strings on the +X side (which faces the Sun

during a satellite imaging campaign for maximum power generation), and two strings on theX side.

Due to the solar array surface extending beyond the satellite bus dimensions, the communication

subsystem requirement of omnidirectional coverage [12]drove the placement of most antennas

on the solar array panel (the single exception being one bus-mounted S-Band antenna). While the

satellite CDR design envisioned four UHF antennas and one S-Band antenna placed at different

angles along the edges of the solar array panel, subsequent analyses and design revisions led to

an exclusively S-Band communications link with four antennas mounted orthogonally on the +X

face of the solar array panel. In addition to these components, a GPS antenna that is part of the

attitude determination and control subsystem is also present near the +Y edge of the solar array panel.

Figure 2.11+X Solar Array Panel (+X side)

+/-Y Body Panels2.3.7

These two structural panels of identical overall dimensions are designed with a baseline

thickness of 2mm and a series of reinforcing ribs, in order to minimize mass while providing

suitable stiffness characteristics. Each panel is populated on the outside surface with seven

photovoltaic cells, in different arrangements according to the particular layout restrictions

applying in each case.

GPS Antenna

Z

X

Y

+Y S-Band

Downlink Antenna

-Y S-Band

Downlink Antenna

-Y S-Band

Uplink Antenna

+Y S-Band

Uplink Antenna

Solar Cell

Strings (8)

+X Sun Sensor

Cutout


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The +Y panel is located near the electronics accommodation section of the trays, and it supports

on its interior surface a Sun Sensor and a Magnetorquer.

Figure 2.12+Y Panel

As a result of the optical payload being positioned in the Y half of the bus, the Y panel is

required to provide an appropriately sized opening for the instrument aperture. A Sun Sensor is

mounted to the inside surface of the panel.

Figure 2.13-Y Panel

Z

X

YZ

X

Y

Solar Cells

+Y Sun Sensor

Cutout+Y Sun Sensor

Y Magnetorquer

Wiring Path

Sun Sensor

Wiring Path

Z

X

YZ

X

Y

Solar Cells

-Y Sun Sensor

Cutout

Instrument Aperture

Cutout

-Y Sun Sensor

Solar Cell

Wiring Path


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X Body Panel2.3.9

The design of this panel is less restrictive than those of all other satellite body panels due to the

absence of solar cells on its exterior surface power generation in a X Sun facing attitude is

achieved by the two solar cell strings present on the X side of the solar array. A rectangularcutout in the panel allows the angled Star Tracker to image a portion of the sky while

appropriately oriented away from the Instrument pointing direction to enable fine attitude

sensing during an imaging campaign.

As detailed in the following section, a design decision was made after CDR to relocate the

Magnetometer sensor board from the boom enclosure shown inFigure 2.4 to a different location

within the satellite bus. Benefitting from the absence of solar cell power wires that would induce

electromagnetic disturbances, theX panel was provided with mounting bosses to accommodate

the Magnetometer PCB on its interior surface. In addition to this component, the panel also

supports a Magnetorquer and a Sun Sensor.

Figure 2.15-X Panel

Z

X

Y

Z

X

Y

X Magnetorquer-X Sun SensorWiring PathMagnetometer

Star Tracker

Baffle Cutout -X Sun Sensor

Cutout


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Toerance Stacking Anaysis for Magnetometer Mounting2.4

The magnetometer is one of the three on-board attitude determination hardware components on

NEMO-AM. Its mounting scheme throughout preliminary and detailed design was based on an

approach which has heritage in previous spacecraft flown by SFL: the PCB containing three

orthogonal magneto-inductive sensors is mounted inside a non-magnetic box-shaped enclosure

that is supported outside the satellite body at a distance of 80mm through a cylindrical tube

section the structure comprising of the enclosure and the support is referred to as the

magnetometer boom (Figure 2.18). The layout of the entire assembly containing the PCB,

wiring paths, thermal control spacers and the magnetometer boom required no major design

work a measure of risk reduction and optimal use of resources by employing a previously

demonstrated solution.

Despite the advantages of this implementation, as the design of NEMO-AM evolved towards

finalization, two points became apparent:

- the satellite mass was exceeding the design budget target of 15 kg or less;

- the larger overall dimensions of the NEMO-AM bus tend to eliminate the need for an

externally-mounted magnetometer. The primary design driver for the magnetometer boom on

the smaller GNB bus was to distance the three sensors from major sources of electro-magnetic disturbances, like power wires and reaction wheels comparable distances can be

achieved within the confines of the NEMO-AM bus.

With these considerations in mind, a simple panel-mounted PCB approach was proposed,

minimizing mass expenditure and choosing a location that would provide enough distance from

sources of significant electro-magnetic disturbance. Relocating an attitude determination

component requires an estimation of the errors introduced by the new mounting scheme.

Consequently, an investigation was done to determine and quantify the tolerance chain that

results in the overall mechanical misalignment between the magnetometer PCB and the closest

primary structural component that it interfaces tothe satelliteX body panel.


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Figure 2.18Magnetometer Boom Mounting

Magnetometer board tolerances

Three sensors and additional electronic components are mounted on a 22 x 42 mm PCB

manufactured by Sierra Circuits Inc. Figure 2.19 shows the board design and dimensions (in

millimetres):

Figure 2.19Magnetometer PCB dimensions (left) and reference system (right)

All angular tolerances used in the error estimation are measured as rotations in a Cartesian

reference system having the centre of the board as the origin, the X and Y axes in the plane of

the PCB, and the Z axis normal to the PCB, as shown inFigure 2.19 - right. Numerical values for

these tolerances are determined based on trigonometric calculations using linear dimensional

Magnetometer PCB

Magnetometer

boom

Delrin spacers

Bus panel

X

Z

Y


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PCB mounting interface tolerances

The magnetometer PCB is designed to be interfaced to a structural component through four

screws (#2-56 thread, made of brassa paramagnetic metal), which will pass through clearance

holes in the corners of the board and thread into the structure. The reported tolerances arederived from part drawings and mounting hardware dimensions. Due to interface similarity, the

results apply both to boom mounting and panel mounting scenarios:

Table 2.3Magnetometer PCB mounting interface tolerances

Mechanical positioning

element

Tolerance

[mm]

Combined

tolerance*[mm]

Distance to

rotation axis[mm]

Angular

misalignment[]

Around

axis

d. Screw play ( clearance

hole minus screw)

0.11 0.155 20.4 0.435 Z

e. Threaded hole position on

structure surface0.10 0.141 20.4 0.396 Z

f. Structure surface height

(boss height)0.10 -

16 0.358 Y

6 0.955 X*tolerances which appply to two orthogonal positioning dimensions on the same feature (e.g. hole vertical and

horizontal distance from board edges) are combined to their root sum square value.

Figure 2.21Magnetometer PCB mounting interface tolerances

Magnetometer boom tolerances

The boom-mounted magnetometer assembly is depicted in Figure 2.18 (cutaway view): it

consists of a PCB enclosure held at 80mm distance from the spacecraft bus mounting surface by

a tubular section. In addition to the PCB mounting interface tolerances listed in the previous

paragraph, supplementary angular misalignments are introduced by the machining precision of

the thermal spacers and the boom enclosure and tubular section (based on dimensional tolerances

prescribed in the manufacturing drawings).

e. Threaded hole

position tolerance

d. Screw to clearance

hole play

f. Surface height

tolerance X

Z

Y


41/107


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sight) to be coplanar with two outside faces of the PCB enclosure (indicated in green in Figure

2.23), thus constraining all three rotations of the assembly.

Figure 2.23Magnetometer boom calibration reference features (green)

Through this procedure, if the assembly is preserved all the way to satellite integration (i.e. the

PCB is not removed from the boom and subsequently re-installed), angular misalignments a-g

associated with the PCB and with the mounting interface are calibrated out. This is because the

readings from the magnetometer sensors are linked directly to the orientation of the boom

enclosure walls that were used for alignment, essentially bypassing the contribution of the

tolerance chain between the magnetometer board and the calibration reference features.

Panel mounting tolerances

In the situation where the magnetometer PCB is mounted directly on the satellite -X body panel,

the previously indicated interface tolerances d, e and f contribute to magnetometer angular

misalignments (Figure 2.21).

Calibration

Calibration is done similarly to the process described in the previous paragraph, by using PCB

board edges instead of enclosure faces as alignment reference features. In this situation, angular

misalignments a-f apply (PCB and interface tolerances). Furthermore, due to the need to mount

the PCB on two different interfaces, the calibration plate and the satellite panel, angular

misalignments d-f (interface tolerances) should apply twice. In this scenario, only the board

manufacturing tolerances aand cwill be calibrated out.

Crosshair

Line of sight 1

Line of sight 2


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Worst-case tolerance stacking

Also called arithmetic tolerance stacking, this method is the most conservative way of

estimating the final dimensional tolerance of an assembly of parts. The primary assumption made

by this approach is that all dimensional imprecisions in the parts are maximal, i.e. equal to themechanical drawing tolerances, and that they are all pointed in the same direction. The use of

this assumption gives a highly unlikely estimation, but which nevertheless covers any possible

combination of part imprecisions in the assembly [14][15]. As the name suggests, the method

relies on identifying and arithmetically summing the absolute values of all the tolerances present

along the same direction within an assembly of parts (2.1). Table 2.5 contains the derived

angular worst-case tolerance stack values for the two scenarios being investigated: the

magnetometer boom assembly and the panel-mounted magnetometer. Because the purpose of

this analysis is to estimate the contribution of mechanical tolerances to the attitude determination

error budget, supplementary misalignments resulting from the calibration process are included.

|| || || || , (2.1)

where T1Tnare tolerances present in the analysed stack-up.

Statistical tolerance stacking

It is generally admitted that the worst-case tolerance stacking method produces

overly-conservative values for assembly dimensional imprecision, and makes unlikely

assumptions about the characteristics of the analysed tolerance chain. The more popular and

practical approach known as statistical tolerance staking was developed in order to obtain

narrower tolerance envelopes that have a certain degree of likelihood (commonly 1, 2, 3

or 6) of covering the dimensional variation of the assembly. The most widely used of all

statistical methods is the RSS method (root sum square method), which is based on two

assumptions: that for a number of identical parts there is a centered normal distribution of actualdimensions around the nominal value, and that the part dimensional variations within an

assembly are independent of each-other (thus highly unlikely to all add up in the same direction)

[14] [15]. The method consists of identifying all tolerances involved in a tolerance chain and

applying a root sum square to determine the likely assembly tolerance envelope (2.2). A

statistical reliability of 3 (99.73%) is associated with this method.Table 2.5 contains the RSS


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values of the derived angular tolerances for both the magnetometer boom assembly and the

panel-mounted magnetometer.

, (2.2)

where T1Tnare tolerances present in the analysed stack-up.

Table 2.5Magnetometer angular errors due to mechanical tolerances

Angular misalignment source

Magnetometer onboom []

Magnetometer onpanel []

X Y Z X Y Z

a. PCB hole edge - - 0.452 - - 0.452

b. PCB board edge - - - - - -

c. PCB thickness at mounting surface 1.050 0.394 - 1.050 0.394 -

d. PCB clearance hole diam. andscrew diam. (screw assembly play)

- - 0.435 - - 0.435

e. Threaded hole position on structuresurface

- - 0.396 - - 0.396

f. Structure surface height(boss face height)

0.955 0.358 - 0.955 0.358 -

g. Delrin spacer height 0.955 0.358 - - - -

h. PCB enclosure top face height - 0.637 0.217 - - -

i. PCB enclosure side face width 0.637 - 0.217 - - -

j. Boom base clearance hole diam. and

screw diam. (screw assembly play)0.916 - - - - -

k. Boom base clearance hole position 0.456 - - - - -

l. Panel threaded hole position 0.456 - - - - -

m. Panel surface height - 0.573 0.169 - - -

Subtotal*Arithmetic 5.425 2.320 1.886 2.005 0.752 1.283

RSS 2.142 1.070 0.820 1.419 0.532 0.742

Calibrationadded - - - f f b, d, e

calibrated out c, f, g c, f, g a, d, e c c a

TotalArithmetic 2.465 1.210 0.603 1.910 0.716 2.339

RSS 1.289 0.857 0.350 1.351 0.506 1.066

*the Subtotal row indicates the stacked angular misalignments in the assembly without any calibration

considerations.

Conclusion

By first comparing the pre-calibration angular tolerance stack-ups (Subtotal in Table 2.5), an

expected condition is confirmed: the more simple panel-mounted approach introduces a smaller


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angular error between the magnetometer and the satellite bus than the magnetometer boom

implementation.

In examining the final results from Table 2.5, the importance of the calibration procedure

becomes apparent. By removing the magnetometer PCB from the calibration plate interface and

mounting it on the satellite panel interface, as it is envisioned in the panel-mounted scheme, a

second set of interface misalignments (b, d, eandf) can potentially be introduced, outweighing

the advantages inherent in the simplicity of this mounting solution.

The results of this analysis were used to update the attitude determination error budget. While the

panel-mounted solution was eventually implemented in the satellite design, measures were taken

to reduce the calibration-related errors:

- of the four mounting screws for the magnetometer PCB, two diagonally opposing ones

will have countersunk heads which will eliminate the play in the clearance holes;

- each screw used on the calibration plate will be kept and used for the same PCB

mounting hole when attaching to the satellite panel.

Wiring Harness Design2.5

As a result of the combination between the distinctively compact design of the NEMO-AM

spacecraft and the complexity of its system architecture, special attention had to be given to the

development of a sound wiring harness concept. While there are numerous subsystem-specific

and component-specific factors that need to be taken into account throughout the wiring harness

design process, some considerations apply overall:

o Wire path selection must, above all, consider the satellite assembly plan. This means that

the designer has awareness of what structure or temporary auxiliary equipment is present

and possibly obstructing access to otherwise advantageous routing areas, or what

structural components are absent and thus not valid choices for wire tie-down points;

o The wiring harness design must be clean. Wire bundles from multiple connectors can and

should use the same paths and tie-down points on the structure if they share the same

routing direction along a certain segment of their length. This process avoids turning the

harness into a webthat precludes access to components after satellite integration;


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o Provided that the previous points are not affected, the wiring harness should minimize

wire lengths as much as possible, especially for connections which have stringent

requirements on signal power loss along the wires. Mass budget savings are also a

desirable outcome from this practice;

o Wiring harness design should be optimized through local structural adaptations, provided

the capability of the structure to carry design loads is not diminished. This close

relationship between the design of the satellite structure and the layout of the wiring

harness is the primary justification for assigning the structural designer with the task of

developing the wiring harness;

o The final specifications for the wiring harness components need to account for minor

variations in the assembly and routing process. Contingency lengths should be added to

the wires in order to allow some slack in the final assembly this is done based on the

designers judgment and on one or several mock routings.

The overall stage of development of NEMO-AM had to be considered when a design approach

was devised for the satellite wiring harness. Because functional testing of various subsystem

electronics and software components is an extensive process, it must be started as soon as

possible, and must be set up in a flight-like arrangement this requires the availability of the

satellite wiring harness components. At the time that the authors work to design the wiring

harness began, there were two major issues to overcome:

- no satellite structural components had been fabricated no physical model was available

to use for wire routing and sizing;

- the spacecraft configuration was not finalized specifically, not all communication

uplink components were selected yet, and the instrument design was not complete.

Under these circumstances, the development of the wiring harness was divided into two steps,

one consisting of a virtual design phase using the satellite 3D CAD model, and the second

involving building a satellite structural mock-up and validating the harness CAD design by

assembling and integrating it into the mock-up.

Before design work commenced, the author compiled the NEMO-AM Wiring Harness

Specifications document. The fundamental resource for this document and the subsequent


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wiring harness design work was the NEMO-AM Interconnect Diagram, a schematic generated

by the power subsystem engineer, showing all satellite electronic components and the pin-by-pin

wiring between them. The wiring harness specifications compiled by the author used the

interconnect diagram to determine each wiring harness component: a connector is selected as a

starting point and each wire is tracked to its destination connector, then the process is repeated

with the newly found connectors until a segment of the harness is isolated, containing all

connectors that have wire runs between each other. The identified wiring harness components are

then registered in the harness specifications sheet as top level groups and each one is populated

with appropriate connector types. Every wire bundle is then listed under its parent connector and

several defining parameters are recorded: connector type, wire type (data or power), pin

designator on end connectors, wire color, wire gauge, twisting of wire pairs (to mitigate

electromagnetic field generation through current loops), length, and other notes. Figure 2.24

exemplifies this process on the harness segment that connects one of the DC-DC Converters to

the Payload Power Board and to a bulkhead Micro-D connector.

Figure 2.24DC-DC Converter wiring harness specification

Satellite interconnect diagram

Wiring harness specifications

Wiring harness between DC-DC Converter and Power Board

Payload power board (PAYPOW)

DC-DC

Converter

15 pin bulkhead

connector


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NEMO-AM Wiring Harness Modeling2.5.1

The task of modeling the wiring harness started by taking the most up-to-date satellite 3D model

available at the time and verifying the positions and labels of connectors on every PCB. The

Solid Edge application module Harness Design was then used to create wiring paths betweenPCB connectors according to the wiring harness specifications document. Each path is a spline

having end points on connectors and several control points along the structure where tie-down

locations were selected. After the path is created, a conductor (wire) is defined on the path. A

sample of the resulting harness model can be observed inFigure 2.25.Because the NEMO-AM

wiring harness contains more than 550 individual wires, adding up to a length of more than 90m,

the sensible approach was to have each modeled conductor represent a bundle of wires sharing

the same path; the physical parameters of the conductor, such as gauge and bending radius, are

defined such that the wire bundle properties are approximated.

A measure of good design was to account not only for wire bending radii, but also for harness

connector dimensions as they protrude outside the PCB-mounted mating connectors. Connector

3D models were easily obtained from the manufacturer website, and they were subsequently

integrated into the satellite 3D model in order to verify that enough clearance was provided

between PCBs and structural components in the vicinity. Structural changes implemented in

order to optimize the wiring harness design and the satellite assembly procedure include:

- introducing cut-outs to allow shorter, more direct routes;

- providing screw-down points for wire zip-tie mounts and P-clips;

- introducing bulkhead connector cut-outs to divide complex harness segments and

facilitate the assembly of theX PCB stack.


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Figure 2.253D model of the wiring harness between the NEMO-AM trays

Wiring harness routing paths were established in this design phase, and wire lengths were

determined by inquiring the properties of each conductor created along the splines describing the

routing paths. Length entries in the wiring harness specifications received an added margin of

5% to allow some flexibility (slack) in the physical wiring harness, along with an extra 5mm to

account for the wire length used on each end to crimp pins.

Satellite Mock-up for Wiring Harness Design Validation2.5.2

Upon completion of the integrated 3D CAD model of the wiring harness and characterization of

each component in order to provide preliminary manufacturing specifications, work commenced

IOBC Stack

-X Tray

PCB Stack

Array Connector

Board

Payload Power

Board

Batteries, BCDR+,

DC Converters

Bulkhead

Connectors

Z

X

Y


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Wiring harness components were also mocked-up by using accurate wire lengths and gauges

based on the specifications determined in the first design stage. Connectors were not included as

the benefits associated with their presence were not significant, especially when weighed against

the added wiring harness manufacturing time and cost expenditure. The resulting wiring harness

components were then integrated with the satellite mock-up and tie-down points, routing paths

and lengths were verified (Figure 2.27). The process demonstrated that the approach of 3D

modeling the harness as a preliminary design step was effective; the final tally showed that less

than 10% of all wire bundles between connectors needed length adjustments of not more than

1-2cm as a result of the verification procedure.

Figure 2.27Wiring harness mock-up on the PCB stack (left) and between the satellite

trays (right)

Sateite Mass Dummy2.6

Because the NEMO-AM bus is a new design that differs considerably from previous spacecraft

built and flown by SFL, the structural subsystem is required to undergo a qualification vibration

test campaign that verifies the subsystems strength and stiffness against design loads (flight

limit loads augmented by a factor of safety of typically 1.25). The proto-flight approach that was

chosen for the NEMO-AM structural qualification means that, provided the post-test inspection

reveals no mechanical failure, the structural components are subsequently used as flight parts.

This reduces the overall time frame and cost of development. Following the structural

qualification, an acceptance vibration test is performed on the final satellite assembly to clear the

system for LV (launch vehicle) integration.


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Since components of every other subsystem are qualified during the last stages of their individual

design, the structural qualification test will make use of mass dummies to substitute these

components within the satellite assembly.

Design of Mass Dummies2.6.1

The major mechanical loading effects that a

Mechanical Aspects of Design, Analysis and Testing of the Nanosatellite for Earth Monitoring and...

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