power system introductory course
TRANSCRIPT
Power System
Introductory Course
Teerapat Charoenpru – Power Engineer
1
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
1
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
Role of Power System Engineer
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
Role of Power System Engineer
– 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.
Typical Power System Requirements
Mission requirement Ex2:
The spacecraft shall be power and thermally safe upon separation from the LV.
Derived requirement Ex2:
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.
Typical Power System Requirements
Mission requirement Ex3:
In the event of an anomaly that results in a loss of controlled attitude, the spacecraft shall be power and thermally safe.
Derived requirement Ex3:
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.
Typical Power System Requirements
Mission requirement Ex4:
Under nominal operating conditions, the spacecraft must permit payload operation at any in its orbit.
Derived requirement Ex4:
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
Teerapat_gistda.or.th