power system introductory course

Power System

Introductory Course

Teerapat Charoenpru – Power Engineer

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Definitions and Abbreviations

COTS Commercial/consumer off-The-Shelf

CONOPs Concept of Operations

DDVP Design Development and Verification Plan

FDIR Fault Detection, Isolation, and Recovery

ICD Interface Control Document

PDM Power Distribution Module

BCM Battery Charge Module

PDR Preliminary Design Review

CDR Critical Design Review

MRR Module Readiness Review

TRR Test Readiness Review

ARR AIT Readiness Review

PSR Pre-shipment Review

TVM Test Verification Metrix

DR Discrepancy Report

OBC Onboard Computer

QSL Qualification Status List

Introduction

No power = Nothing

“One of the critical components of any spacecraft is the power system that allows the operation of its many systems.”

https://wmleader.com/technology/1046/were-entombing-the-earth-in-an-impenetrable-shell-of-dead-satellites/

Introduction

Standard 2U CubeSat diagram1

1https://www.researchgate.net/figure/Standard-2U-CubeSat-diagram-CubeSats-are-small-scale-satellites-composed-of-several_fig1_331595385

2https://directory.eoportal.org/web/eoportal/satellite-missions/u/unisat

Illustration of the COTS camera2 Illustration of the COTS GPS reciver2

https://ocw.mit.edu/courses/aeronautics-and-astronautics/16-851-satellite-engineering-fall-2003/lecture-notes/l3_scpowersys_dm_done2.pdf

Power source applicability

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2

3

4

Introduction

Introduction

1https://www.popularmechanics.com/space/satellites/news/a28496/how-sputnik-worked/

Sputnik 1 contained:• radio transmitter• remote switch• thermal control system• barometric switch

• silver-zinc batteries

The satellite sent out radio beeps for 22 days before the silver-zinc batteries ran out.

Two of these powered the radio, while the third one was responsible for temperature control.

Ref: https://www.upsbatterycenter.com/blog/first-batteries-space-silver-zinc/

Sputnik 11

International Space Station1

1 https://www.esa.int/Science_Exploration/Human_and_Robotic_Exploration/International_Space_Station/

SSTL Carbonite 2 Satellite2

2 https://www.sstl.co.uk/what-we-do/earth-observation-spacecraft

Introduction

Introduction

https://www.youtube.com/watch?v=4qkvoVRdoNg

Multi-mission Radioisotope Thermoelectric Generator

Role of Power System Engineer

– Delivers the power sub-system for all our spacecraft

– Power sub-system responsible for the generation/conversion, storage and distribution of power across the spacecraft

– Includes the following ‘hardware’:

• Solar Panels

• Batteries

• Power Distribution Modules

• Battery Conditioning Modules

• Solar Array Regulator Modules

• Thermal Knife / Pyro Driver Modules

• Activation Switches

• Safe Arm, Battery Arm, Dump Resistors


Battery Charge Module (BCM)

Redundant BCR

Array Plugs

Nominal BCRs1-6

Li-ionBatteryPack

BCM

Power distribution module

Dump resistor

2xPower switches

60xPower

switches

Battery Arm PlugOverVoltageClamp

Activation Switches

Solar Arrays

5v Dc to Dc

CAN

5v Dc to Dc

CAN

2 battery bus voltage switches

60 battery bus voltage switches

Battery Arm Plug


– Spacecraft grounding system

• Subsystem requirements

• Grounding scheme

• Testing in AIT

– Provides electrical support for other teams that require power/analogue electronics expertise:

• Reaction Wheels Driver

Role of Power System EngineerPDR

•Derive power

system

requirements

•Heritage baseline

• Initial solar panels

and battery sizing

•Support system

engineer (power

budget)

CDR

•Detail design

(Freeze)

•Qualification status

list

•TVM

•Risks

•Support derived

requirements

•Schematic and

layout design

MRR

•Prepare for build

•Schematic and

PCB layout

reviewed

•De-rating

•Parts assessment

•Build procedures

reviewed

•Test procedures

draft

TRR

•Prepare for testing

•Test procedures

reviewed

•Test equipment

and facilities check

•Test result

template

ARR

•Prepare for AIT

• Integration

procedure

reviewed

•AIT test procedure

reviewed

•AIT test result

template

PSR

•Modules status

update

•TVM update

•Closed all DRs

•Closed all Risks

•Closed all peer

reviews and

actions

•Release all

documents

General Power System Design

•Power switch purpose is to protect the battery bus; not the loads

•In the case of OBC failure, the Power System sets all loads to a ‘safe’ state

•Shuts down all loads in the event of the battery reaching 100% Depth of discharge DoD (except loads on fused lines,

such as the Rx)

•Can provide a Low-Level Command Link (LLCL) to the receiver

Power system protects the mission

•Graceful degradation - Solar regulation function

•Dual redundant - CAN nodes and Activation system

•Redundant loads on separate switches

•BCM switches

Designed to avoid credible SPFs

•Maximum Power Point Tracking

•Efficiency/Reliability trade-off

•Bus voltages

Architecture

Mission Specific Power System DesignPower Generation

•Required Orbit Average Power

•Body mounted or sun tracking

•Solar Cells degrade over time so mission lifetime important

•Number of BCRs and BCMs required

Power Storage

•Power required in eclipse

•Peak discharge current

•Required Battery Bus voltage

•Cells degrade over time so mission lifetime important

Power Distribution

•Number of loads and number of power distribution switches required

•Specific needs such as series switches, timed switches, etc

•Telemetry for monitoring and failure detection

Spacecraft Activation

•Activation Switches acceptable for peak current, or Relay Module required

Typical Power System Requirements

Mission requirement Ex1:

The satellite design shall be such that no credible single point failure can lead to loss of mission.

Derived requirement Ex1:

Power system interfaces shall be single point free.

Power system design also considers the failure modes within battery, partial redundancy is valid with BCRs, Solar panels.



The spacecraft shall be power and thermally safe upon separation from the LV.


Power system shall ensure the spacecraft is unpowered on the launch vehicle.

The power system shall autonomously activate.

On separation from LV the power system shall only activate specified loads.



In the event of an anomaly that results in a loss of controlled attitude, the spacecraft shall be power and thermally safe.


In term of power consideration, the design ensures to meet the maximum rating(i.e., solar array output power output). The power analysis for the tumbling mode is carried on by system level power analysis.



Under nominal operating conditions, the spacecraft must permit payload operation at any in its orbit.


The solar panels and battery shall be sized to accommodate the required mission power budget.

Partial failures (e.g., loss of a battery string) may result in degraded performance.

Solar Arrays design

•Convert solar energy into electrical energy to meet the power requirements of the mission over the mission’s

lifetime.

Primary Function

•Efficiency, Qualification Status, Heritage > Cell type and size

•Maximum Open Circuit Voltage at BOL > Maximum string length

•Minimum MPP Voltage at EOL > Minimum string length

•Maximum Power at BOL > Number of strings per section

•Minimum Power at EOL > Number of strings

•Redundancy > Number of strings

•Environmental Factors:

•Radiation Degradation > Mission orbit and lifetime

•UV, Micrometeorite Degradation > Mission orbit and lifetime

•Temperature Effects > Expected panel operating temperature range

•Available Surface Area > Fixed / Deployed / Sun Tracking

•Substrate > Cell type and size, laydown

Solar Array Design Drivers

Solar Arrays design• Design Process

• Power System Engineer works with System Engineer to define power requirements and mechanical constraints

• Electrical Design

• Design is usually for end of life (EOL)

• Cells are interconnected in series to give a string of cells at thecorrect voltage at EOL

• Number of strings required is determined by the power requirementat EOL

• Strings separated into sections based on BCR maximum powerthroughputs

• Calculation for size is determined by BOL and EOL loss factors

• Mechanical Design

• Body-mounted or deployable? > Depends on power requirement

• What substrate type? > Depends on cell type requirement (largearea cells need CFRP because of the low thermal expansion)

• What size of panel ? > Usually determined by spacecraft size

• Deployment mechanism ? > Heritage and reliability

Solar Arrays design

1https://iaeimagazine.org/columns/photovoltaic/back-to-basics-pv-volts-currents-and-the-nec/

Photovoltaic I-V curve1

BOL EOL

Hot Hot

Normal Normal

Cold Cold

Best case Worse case

Solar Arrays design

1Mitsuru Imaizumi Space Solar Cells

• Build Process– Solar Cells procured

– Solar Cells processed into Solar Cell Assemblies• Inspected• Interconnects and cover glass fitted• Measured and sorted into current classes

– Welding into strings (Ultra sonic welding)

– Strings laid down (glued) onto panels

– Strings wired into sections on panel (Parallel gap welding)

Complete solar cell (space)1

Bare solar cell (space)2

Electronics Modules (BCM & PDM)Primary Functions

•Regulate power from the solar arrays to the spacecraft loads and battery

•Enable activation/deactivation of spacecraft loads (by the OBC)

•Protect the battery bus from failures associated with spacecraft loads

Other Functions

•Provide overvoltage protection to protect the battery from overcharging

•Activate/deactivate spacecraft loads in the event of an OBC failure

Design Drivers

•Solar Section Characteristics (Power/Voltage)

•Efficiency

•Battery Voltage

•Maximum Power Point Method

•Thermal Dissipation

•Number of Spacecraft Loads

•Required Current of Spacecraft Loads (nominal, maximum, in-rush)

•Redundancy

•Switch Trip Point & Current Telemetry Setting & Accuracy

Electronics Modules (BCM & PDM)

• Design Process

Typically based around ‘heritage’ power modules

– Battery Conditioning Modules

• Power requirements define number of BCRs required and hence number of BCMs required

• Typically, 1 BCM (6 BCRs) for 100/150kg satellite and 2 BCMs (7-12 BCRs) for 300kg satellite

• Component values defined for EoC. based on battery

• Power Component fit/no-fits defined based on number of BCRs required

• Mission specific firmware generated (e.g., LLCL encryption)

• Often last-minute system level changes so firmware can be reprogrammed using In circuit programming interface

Electronics Modules (BCM & PDM)

• Design Process

Typically based around ‘heritage’ power module

– Power Distribution Modules

• Number of loads and load power requirements define number of PDMs required

• Switches allocated to loads based on maximum current requirements, redundancy (to avoid XOR SPFs), series switch requirements

• Often long process as not all loads defined until late on in the system design process

• Switches set up for their maximum rated current and current TLM scaling set to the switches maximum de-rated current output; always last-minute system level changes to switch allocation so only changes to harnessing and no hardware changes required to power system.

• Often last-minute system level changes so firmware can be reprogrammed using In circuit programming interface

Electronics Modules (BCM & PDM)• Build Process

– Reflow > Majority of parts

– Hand Assembled parts > Inductors

– Hand Fit > Inductors, chassis mounted components, connectors,

– Select on Design Components > EoC set-point, Charge Current limit

– Assemble into module housing

Electronics Modules (BCM & PDM)• Test Process

– Module Level Electrical Testing• Check functionality & performance• PDM – Trip Set points, switch telemetry, watchdog, mask, LLCL, mission specific commands• BCM – BCR Efficiencies, MPPT, EoC, overvoltage clamp, array telemetry

– Module Level Thermal Cycling• Check workmanship; 3 cycles, +50degC to -20degC @ 120degC/min

– Module Level Thermal Testing• Check functionality & performance at temperature (+50degC and -20degC)• Subset of module level electrical testing

– Module Level Burn-in• 120 hours prior to AIT, 168 hours prior to FRR• Check for infant mortality of parts

– Module Delivered to AIT

Battery

•Store power for use by the spacecraft loads when there is insufficient power available from solar arrays

Primary Function

•Nominal Bus Voltage > String length

•Minimum Operating Voltage > Expected DoD

•Minimum Power at EOL > Number of strings / cell capacity

•Redundancy > Number of strings / cell capacity

•Environmental Factors:

•Degradation > Lifetime and Cycling

•Temperature Effects > Expected operating temperature range

•Vibration > Cell type

Design Drivers

•Run as a subcontract

•Battery configuration determined by mission power requirements

•Mission power profile provided by Mission System Engineer

•Typically, offerings from suppliers

•Down selection process

Design Process

Battery

Build process

•Contract raised

•Manufacturer makes/procures cells and assembles into battery pack(s)

•Typically, EM (workhorse) and Vib Pack also procured

•Typical battery build time is 24 weeks for EM and 52 weeks for FM from contract KO

Test Process

•Visual, Energy Capacity and Internal Resistance check on receipt

•EM used for AIT testing

•FM used for TVT testing > kept on a separate hotplate to maintain battery temperature

•Stored at 5degC, 80% DoD

•Transported at 80% DoD

•Charged and maintained at 100% SoC at launch site

Activation switches and Relay modules

Primary Function

•Activate the spacecraft power distribution bus (Powerbus) once the spacecraft has separated from the launch

vehicle

Design Drivers

•Maximum Powerbus Current

•Number of Operations

•Powerbus Voltage

•Voltage Drop

•Thermal Dissipation

•Activation Method (Plunger, breakaway connector, etc)

Design Process

Build Process

•Activation switches as part of Spacecraft Harness

•RM as per Electrical Modules

Test Process

•Activation switches as part of Power System AIT tests

•RM as per Electrical Modules

Q&A

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power system introductory course

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