Microsatellite Design and Integration for INSPIRE · DWTS Design & Interface 15 draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum electrical

Microsatellite Design and Integration for INSPIRE: International Satellite Project in Research and Education

Loren Chang, Upper Air Dynamics Lab

Institute of Space Science, National Central University, Taiwan

2016-10-25 APSCO & ISSI-BJ Space Science School

Microsatellite Design and Integration for INSPIRE: International Satellite Project in Research and Education

Loren Chang, Upper Air Dynamics Lab

Institute of Space Science, National Central University, Taiwan

2016-10-25 APSCO & ISSI-BJ Space Science School

Nano

C.K. Chao, Jude Salinas, Ya-Chih Mao, Jack Chieh Wang, Jia-Yu Su, Duann Yi, Pei-Yun Chiu, Joe Hong, Yi-Chung Chiu Institute of Space Science, National Central University, Taiwan Steven C.R. Chen National Space Organization, Taiwan Amal Chandran, Michael McGrath Laboratory for Atmospheric and Space Physics, University of Colorado, USA Dave Fritts, Larry Gordley GATS, USA Priyadarshan, Kastubh Anand Indian Institute of Space Science and Technology, India Tarun Pant Space Physics Laboratory, ISRO, India

INSPIRESat team

Overview

• The INSPIRE Project

• Designing a Mission Concept: • Scientific Needs: The Middle and Upper Atmosphere

• Science Traceability: Doppler Wind and Temperature Sounder (DWTS)

• Can a satellite do this? Feasibility, System Definition, and Lessons Learned

Space Science and Capacity Building at NCU

INternational Satellite Program In Research and Education

• CubeSat constellation for Earth and Space Science Research (2 missions currently

planned)

• Global ground station network

• Academic program in mission & spacecraft design, operations, and data analysis.

University of

Colorado LASP

Indian Institute of Space

Science and Technology /

ISRO

Taiwan National Central

University / NSPO

INSPIRESat-1 Feasibility Study Subsystem NCU CU IIST

Payload

Science

STR

TCS

CDH

COM

EPS

ADCS

Launch

http://www.theonion.com/article/nasa-announces-plan-to-launch-700-million-into-spa-1950

Why is this mission needed?





















Mission Concept

Review (MCR)

To evaluate the feasibility of

the proposed mission

concept(s) and its fulfillment

of the program's needs and

objectives.

To determine whether the

maturity of the concept and

associated planning are

sufficient to begin Phase A.

http://nodis3.gsfc.nasa.gov/displayDir.cfm?Internal_ID=N_PR_7120_005E_&page_name=Chapter2



TROPOSPHERE

STRATOSPHERE

MESOSPHERE

THERMOSPHERE

You are here

Earth’s Atmosphere

~ 20 km

~ 60 km

~ 85 km

Drag is the largest orbital perturbation in LEO!

Orbital debris distribution

https://orbitaldebris.jsc.nasa.gov

GEO

MEO

LEO

TROPOSPHERE

STRATOSPHERE

MESOSPHERE

THERMOSPHERE

You are here

Satellite Drag

Reentry Dynamics

Ionospheric Variability

Communications

Stratospheric Polar

Vortex

Weather Prediction

Middle & Upper Atmosphere (>20 km)

~ 20 km

~ 60 km

~ 85 km

TROPOSPHERE

STRATOSPHERE

MESOSPHERE

THERMOSPHERE

You are here

~ 20 km

~ 60 km

~ 85 km

Limited Observational Ability

Gordley et al., SPIE 2013

Doppler Wind and Temperature Sounder (DWTS)

• New gas filter correlation radiometer measurements of wind and temperature (~25 – 250 km)

• Cross-track limb pointing IR camera with cryocooler.

• Designed and patented by GATS. Fabrication by University of Colorado LASP.

• Still requires operational heritage in space environment (TRL 4).

Limb scan of NO emissions

Doppler shift and width of

emission lines

Atmospheric winds and

temperature from 30 to 200

km

DWTS Design & Interface

15

draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum

electrical power draw of the cooler of 20W, thus we have significant margins to drive the cooler

harder. Thermal control is maintained through a combination of active and passive methods.

The optics (including the gas cell) are maintained to < 240K. For the flight instrument, the heat

pulled by the cooler is transferred through heat pipes using non-toxic fluids to radiators on side

panels. Remaining thermal control is through heaters and passive conduction/radiation. The

envisioned constellation of “free-flyers” will allow much cooler optics and gas cells, which

reduces noise and narrows the cell line-widths improving the DWTS DIP signature.

Figure 13. Dark current density versus

temperature for InAsSb photodiode and nBn

detector, and HgCdTe photodiodes with a cut-off

wavelength of 5 μm. From: Performance

comparison of barrier detectors and HgCdTe

photodiodes (Opt. Eng. 2014;53(10):106105.

doi:10.1117/1.OE.53.10.106105)

OPTICS: The draft DWTS optical design is shown in Figure 14. This design supports the

SB204 detector form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal

length, 30 mm aperture, resulting in a fast F/1.4 system. The lens includes mixture of

Germanium and ZnSe lenses as required for stray light control and image quality. Both

spherical and aspherical elements are used, as the optics shall be diamond turned to the required

surface figure and finish.

Figure 14. DWTS Optical Design

ELECTRONICS: The camera electronics shall be 4-5 Watts from previously built boards.

The cryocooler controller electrical budget is included in the 10W given. Thus the power for the

instrument needed shall be 14-15 Watts (without reserves) or 23 Watts with 50% reserves. The

mass is estimated to be less than 4 Kg (1 kg for the Integrated Detector Cooler Assembly, .75

Kg for the mounted optics, 0.75 Kg for the electronics and cable hardnesses, and 1.5 Kg for the

structure and mounts. A radiation tolerant by software design approach shall be taken using

Figure 15. Single Channel DWTS Sensor Head

14

and IDL save sets. Useful mission documents, data user guide, published papers, and software

tools (usually in IDL) for reading and plotting the DWTS data will be provided on a DWTS

project web site on CU/LASP web server. The LISIRD databases are designed for large (TB)

data sets, so all of the DWTS data will be kept on-line, and the LISIRD backup system. We

intend to preserve public access of the data for the life of the LISIRD system. However, LISIRD

is not configured, nor funded, to be a long-duration archive center, so DWTS archive will be at

a NASA archive site.

2.2. DWTS Instrument Design

The DWTS instrument design leverages existing

technologies and commercial off-the-shelf (COTS) hardware.

The heart of the instrument is the Lockheed-Martin Santa

Barbara Focalplane SBF204 nBn Focal Plane Array with a

Ricor K527 cryocooler, integrated into a Brandywine micro-

cryocooled camera. Figure 12. Ricor K527 Split

Linear Cryocooler

DETECTOR: The detector format is 1296 x 1040 (8 pixel borders) with a 12-micron pitch

as shown in Table 2. This FPA is produced in large volumes (>50/month) for tactical programs,

with similar SBF nBn FPAs having passed radiation testing at the Air Force Research

Laboratory. nBn was chose over HgCdTe due to having COTS FPAs in stock, and having better

epitaxial growth properties (better pixel gain uniformity and higher operability.)

Parameter SBF204

Format: 1296 x 1040 (8 pixel borders)

Windowing: Windowing: 608 x 8 up to 1280 x 1024 in 1 x 4 pixel increments.

Pixel Size: 12 µm x 12 µm

Detectors: nBn with QE = 40% @ 5.25 um

Outputs: Selectable 1, 2, or 4 LVDS Data pairs, plus 1 LVDS clock pair

Integration Modes: 1: Snapshot, Integrate Then Read

2: Snapshot, Integrate While Read

Line Rate: < 125 kHz in four port mode

Frame Rate: < 99 Hz for 1280 x 1024 in four port & 14 bit mode

A/D Resolution Selectable, 14 or 13 bits

Signal Capacity: 50 Ke-

Total ROIC Noise: < 30 e-

Responsivity: 3 e-/LSB

Dark Current <10 nA/cm2 @110K (from graph for 5um cutoff)

Power: < 250 mW

Table 2. Lockheed Martin SBF204 Specifications

THERMAL CONTROL SYSTEM: Our operating requirements are a heat lift capacity of

400 mW, cold tip temperature of <90K while at a rejection temperature of -20K. The

manufacturer has confirmed that we can easily meet these requirements with an electrical power

Subsystem Notes

Camera Teledyne GEOSNAP FPA (in fabrication)

2000 x 2000 pixels

Ricor K527 cryocooler

Optics Includes NO gas cell. NO 5.3 𝜇m (MidIR)

absorption.

43.3 mm effective focal length

30 mm aperture.

7° x 20° FOV

Cooler & Heater Gas cell temperature control.

Processor Xiophos Inc. Q6 FPGA

Data aggregation.

0.5 Hz sample return to bus: FPA data,

cooler and heater settings.

Data storage and power provided by bus.

Test Objectives

• Mission shall demonstrate ability of DWTS to

measure winds and temperatures from 30 –

200 km.

• DWTS shall be demonstrated in a space

environment.

• Demonstrate that stray light effects can be

eliminated using optical or digital means,

including lunar scanning to demonstrate stray

light calibration.

• The satellite shall resolve planetary-scale

waves and tides with zonal wave numbers 1

to 4 from the lower to the upper atmosphere

from 60S to 60N.

Mission Objectives (Initial)

• The mission shall survive the space environment.

• The mission life-time shall be at least 3 months.

• The data rate shall be 225 megabits/day.

• The communication system for the mission shall be VHF and UHF radio frequency bands.

Customer Top Level Requirements

• The mission shall have four ground-stations- Colorado, Taiwan, Singapore and India.

• The mission shall be launched by ISRO as a secondary payload.

• The spacecraft shall be disposed of at the end of life.

• The satellite bus shall be a CubeSat with 6U dimensions.

Will a 6U CubeSat be able to support the DWTS

payload and mission?

Determine payload requirements.

Set subsystem requirements that

satisfy payload requirements.

Identify available COTS options that

satisfy payload requirements.

*Feasibility Study funded by National Space

Organization of Taiwan for SY 2015-2016

System

Requirements

Review (SRR)

To evaluate whether the

functional and performance

requirements defined for the

system are responsive to

the program's requirements

on the project and represent

achievable capabilities.




DWTS Requirements & Specifications

Mass 4 kg

Volume 29 x 10 x 9 cm (3U)

Power Standby and Operational: 7 W

Safehold: 1 W

Off: 0 W

Pointing

Knowledge

±0.5 arcmin, all axes

Pointing Stability < 6 arcsec/sec

Attitude Control 1° all axes

Field of View Sun most not appear in field of

view.

Spacecraft Velocity

Knowledge

±1 m/s

Data Rate Downlink 200 Mbits/day

Thermal Stability FPA: ±0.1 K/minute

Gas Cell: ±1 K/minute

Thermal

Requirement

Anti-Sunward Side: -10 – 0°C

Observed Emission NO 5.3 𝜇m (MidIR)

15















doi:10.1117/1.OE.53.10.106105)















15















doi:10.1117/1.OE.53.10.106105)















SUB-

SYSTEM

MASS

PAYLOAD 4.000 kg

ORB/NAV 0.012 kg (pqNAV)

EPS 0.580 kg (GOMSpace

20 SA panels)

CDH 0.088 kg (Pumpkin)

COMM 0.078 kg (Astrodev)

0.100 kg (Antenna)

ADCS 0.850 kg (XACT)

STR 1.640 kg (6U Chassis)

TOTAL 7.348 kg

Margin 0.652 kg

Mass Budget

ORBIT, GUIDANCE & NAVIGATION (ORB/NAV) ATTITUDE DETERMINATION AND CONTROL SUBSYSTEM (ADCS) ORB/NAV Manager: Jack Chieh Wang

ADCS Manager: Cornelius Csar Jude H. Salinas

ORB/NAV Requirement: The mission orbit shall be at least

350 km in altitude and be at near-circular orbit.

NASA Debris

Assessment

Software

650 km: 37.7 years

800 km: >100 years

Debris Mitigation

Deorbit Guidelines:

25 years

ORB/NAV Requirement: The mission orbit shall be at least

60N in inclination.

At 60 N, local time precession

rate is:

39.7097 days at 650 km altitude

42.0544 days at 800 km altitude

ORB/NAV Requirement: The mission shall monitor the

orbit via ground-tracking and GPS.

COTS Option for Navigation

Best Option

pqNAV-L1/FM

(SkyFox LabsTM )

SGR-05U

(Surrey)

GPS Receiver

(SSBV)

GPSRM 1 GPS Receiver Module

(CubeSat Kit)

Feature

GPS L1 C/A

signal, 15

channels

GPS L1 C/A

signal,

12 channels

GPS L1 C/A

signal,

12 channels

GPS L1/L2/L2C signal,

GLONASS L1/L2 signal

Price 4900 EUR =~5360

USD ? ? ?

Mass 12 gram 40 gram < 30 gram ?

Dimensions 42*42*15 mm 70x45x10mm 50x20x5mm ?

Power 120 mW

(3.3V@25°C) 800 mW(5V) < 1W (5V) ?

Temporal

resolution 1 Hz 2x106 Hz > 1 Hz ?

Operation

Temperatur

e

-40~+85°C -20~+50°C -10~+50°C ?

Altitude Up to 3600 km ? (Already flown) ? (Already flown) ?

Velocity Up to 9 km/s ? (Already flown) ? (Already flown) ?

ADCS

REQUIREMENTS

ADCS CAPABILITY

BC XACT

• Payload shall always point

cross-track and away from

the sun.

• Attitude knowledge: +/- 0.5

arcminutes

• Attitude control: +/- 1

degree for all direction

angles.

• Pointing stability: <6.0

arcsecs/sec.

• Attitude knowledge: +/-

0.06 arcminutes for all

direction angles

• Attitude control:

+/- 0.003 degree (1-sigma)

for 2 axes

+/- 0.007 degree (1-

sigma) for 3rd axis

• Pointing stability:

1 arcsec/sec for all axes

XACT: On orbit tests on MinXSS

http://digitalcommons.usu.edu/smallsat/2016/S5GuidCont/2/

THERMAL CONTROL SUBSYSTEM (TCS) TCS Manager: Sunny Duann-Yi

Components Operational Survival

EPS -40 ~ 85 ℃ -50 ~ 90 ℃

Batteries 5 ~ 20 ℃ 0~ 25 ℃

Solar Panels -150 ~ 110 ℃ -200 ~ 130 ℃

DWTS Payload -30 ~ 20 ℃ -35 ~ 25 ℃

GPS -40 ~ 85 ℃ -55 ~ 95 ℃

Reaction Wheels (ADCS) -10 ~ 40 ℃ -20 ~ 50 ℃

Motherboard -40 ~ 85 ℃ -45 ~ 90 ℃

Antennas -100 ~ 100 ℃ -120 ~ 120 ℃

Worst Case Estimates: Radiative Equilibrium

-69.4 – 50.1 C at 600 km

-72.2 – 48.8 C at 800 km

Lesson Learned: TCS

Payload requirements should be frozen before a

preliminary design for the spacecraft.

14

The DWTS optical design is shown in Figure 13. This design supports the GEOSNAP detector

form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal length, 30 mm

aperture, resulting in a fast F/1.4 system. The lens includes mixture of Germanium and ZnSe

lenses as required for stray light control and image quality. Both spherical and aspherical

elements are used, as the optics shall be diamond turned to the required surface figure and

finish. The optics will have low-emissivity anti-reflection coatings to keep internal emissions

low. The inside surface of the lens barrel and the sensor housing will be coated black to reduce

reflected stray light.

Thermal Design: The thermal design has two actively-controlled thermal zones, the detector,

which must be kept at 95 K (110K is acceptable) to reduce dark current to our goal of

approximately 100 e-/pixel/s, and the cryocooler radiator which rejects heat at 23 C. In

addition, the cryocooler will use passive cooling to keep the optics 210 K to reduce thermal

emission. A Ricor K527 Stirling split-linear cryocooler will be used to cool the detector. The

K527 has a cooling capacity of 0.7W for an active

load at 95K and requires 20W input power. A

flexible thermal strap will be used to connect the cryo

cold head to the detector, providing vibration

isolation from the cryocooler pistons. A large

deployable radiator will be used to reject the heat

generated by the cryocooler. Precise control of the

temperature in both zones will be accomplished via

heaters placed on the detector housing and on the

radiator. The detector will be mounted to the lens

assembly using thermally-isolating mounts, and then

both are enclosed inside a double walled thermal

enclosure. The lens assembly will be mounted to the

enclosure also using thermal isolating mounts. To maintain radiative isolation, the outside

surface of the lens barrel and the inside surface of the enclosure walls will be gold coated.

Figure 12 shows the instrument concept thermal model. End of Life (EOL) material properties

were used for hot case analysis and Beginning of Life (BOL) properties were used for cold case

analysis. In the worst case hot orbit scenario, the cryocooler active load is 0.54 W, which has

30% margin to the K527 capacity of 0.7W. Additional margin could potentially be gained

through detailed design of the isolating mounts and increased detector flex cable length. The

maximum predicted heater power required to maintain the detector and radiator at the design

temperatures are 0.25W and 16W respectively. There are also maneuvers that can be

performed, such as 180 degree role about the line of sight (effectively flipping the FOV, which

does not affect the measurements), that will dramatically improve radiator and solar array

efficiency. This strategy was not fully evaluated before the proposal, but will conservatively

add 30% to our cooling margins.

Figure 12. Thermal Model of Sensor and Platform.

• Later payload component level thermal analysis by

LASP engineers found that a radiator is needed for

sufficient heat dissipation from DWTS cryocooler.

• Cryocooler introduces 60 Hz disturbance torque that

must be characterized by duty cycling cryocooler

during payload operation.

• TCS and ADCS requirements require revision to

accommodate.

COMMUNICATIONS (COM)

COM Managers: Ken Lai, Pei-Yun Chiu, Kitty Jia-Yu Su

HOUSEKEEPING DATA: 1. Orbital position and

velocity (ORB) 2. Battery Condition (EPS) 3. Power Consumption

(EPS) 4. Sub-system and

payload current temperature (TCS)

5. Attitude (ADCS) 6. Angular Velocity (ADCS) 7. Universal time 8. Error Feedback

RAW MISSION DATA: 2D Radiance in Pixel Values

FINAL MISSION DATA: Temperature and

Horizontal Wind Vertical Profiles

GROUND RECEIVING STATION 1. Mission Control

2. Ground-based processing

USERS: Researchers, companies and mission center

new commands

COMMUNCATIONS ARCHITECTURE:

The space-craft will utilize a centralized communication system. All data-processing will be done on the ground.

STK Analysis at altitude 800 km and inclination 65N

Total access time/day in

worst case

Colorado, U.S 6,250 Seconds

SARC,

Singapore 3,286 Seconds

ISRO, India 3,444 Seconds

NCU, Taiwan 2,400 Seconds

Total 15,380 Seconds

With the 225 megabits per day requirement, it is required that the

communication module onboard the spacecraft must achieve a minimum

data rate of 14.629 kbps.

COM

REQUIREMENTS

COM DESIGN

• COMM data rate: 225

megabits/day.

• Transmit to four ground

stations: Colorado, Taiwan,

Singapore and India.

• Data rate: 590 megabits per day

• Ground stations: Colorado,

Taiwan, Singapore and India

• COMM Module data rate:

> 38.4 kbps

• VHF and UHF radio frequency

bands COMM system.

• AX.25 packet protocol

• GMSK Modulation downlink.

• COMM Module Frequency

bands: VHF and UHF

• Protocol: AX.25 Packet Protocol

• Modulation: GMSK Modulation

AstroDev Helium 100 Radio

38.4 kbs

(120 – 150, 400 – 450 MHz)

ISIS Deployable Antenna

NCU Ground Station • VHF/UHF Amateur

radio bands.

• HyGain 218SAT antennas with dual bandpass filters.

• ICOM 9100 transceiver.

• Kamtronics KAM-XL terminal node connector.

Lesson Learned: COM

Remember to account for reduced tracking speeds during high

elevation angle passes, and losses at low elevation angles.

Typhoon Magi

NCU Weather Station

Vmax = 54 km h-1

Lesson Learned: COM

Exercise caution in defining ground operations SOP and

be strict on implementation!

ELECTRICAL POWER SUBSYSTEM (EPS) EPS Manager: Cornelius Csar Jude H. Salinas

TLR/PAYLOAD

REQUIREMENTS

EPS REQUIREMENTS EPS DESIGN

1. Science goal TLR

2. Engineering goal TLR

3. 3-month Mission Lifetime

4. DWTS power requirement is 7W.

5. DWTS has four modes: Off (0W),

Safehold (1W), Standby (7W) and

Operational (7W)

• Adequate power for at least 3

months.

• Total power: 17.13 W (0.8565 W

margin)

• Solar panel power : 44 W

• Battery capacity: 20 Wh

• Protection against: Bus Voltage,

Bus Faults, Under Voltage, Over

Voltage, Over Current and Single

Event Effects.

• Switching mechanism such that

the switching is on the positive

side.

• Protection Mechanisms: Under

voltage, over voltage (3.3 V, 5 V

and 12 V) and over current

protection. Temperature sensor.

EPS micro-controller unit.

SUB-

SYSTEM

COTS OPTION WITH HIGHEST POWER

REQUIREMENT

PAYLOAD 7 W (DWTS)

ORB/NAV 1.3 W (GPSRM)

EPS -

CDH 0.003 W

COMM 6 W (Astrodev Helium Radio)

ADCS 2.83 W (XACT)

TOTAL 17.133 W(maximum)

EPS Requirement: The EPS shall provide for a required

power consumption of 17.133 W (0.8565 W margin).

InSPIRESat-1

MISSION ORBITAL PERIOD 100 minutes

ECLIPSE PERIOD

(SPAT Analysis)

1.9758 – 35.0728 minutes

TOTAL POWER

CONSUMPTION

17.13 W

REQUIRED SOLAR

PANEL POWER

(assuming 30%

efficiency power loss for

all transfers)

25.1709 W to 43.1479 W

EPS Requirement: The solar panel shall provide 44 W of

power for the spacecraft.

InSPIRESat-1

MISSION ORBITAL PERIOD 100 minutes

ECLIPSE PERIOD

(SPAT Analysis)

1.9758 – 35.0728 minutes

TOTAL POWER

CONSUMPTION

17.13 W

REQUIRED BATTERY

CAPACITY (assuming

30% efficiency and 80%

DOD)

17.8809 Wh

EPS Requirement: The EPS battery shall have a required

capacity of at least 20 Wh.

Are there COTS solar panels that can provide 44 W?

GOMSpace NanoPower P110 Series solar panels (http://gomspace.com/?p=products-p110).

Clyde Space 3U Solar Panel (https://www.clyde.space/products/25-3u-solar-panel).

Pumpkin Space Systems 3U Solar Panel (http://www.cubesatkit.com/docs/press/P

umpkin_CSDWLU_2010-2.pdf).

OPTION 2:

3U Size at

7S1P

Configuration

7W at BOL

OPTION 3:

3U Size at

8S1P

Configuration

8W at BOL

OPTION 1:

60.36 cm2 Size

1 Solar Panel

2.3 W at EOL

ESTIMATED NO. OF

PANELS AND AREA:

20 panels. 1207.2 cm2.

Are there COTS batteries that can provide 20 Wh?

Clyde Space 20 Whr stand-alone battery

(https://www.clyde.space/products/22-20whr-

cubesat-battery).



cubesat-battery).



cubesat-battery).

GOMSpace battery pack 38.9 Wh (http://gomspace.com/document

s/ds/gs-ds-batteries.pdf).

OPTION 1:

20 Wh capacity

OPTION 2:

30 Wh capacity

OPTION 3:

40 Wh capacity

OPTION 4:

40 Wh capacity

Are there COTS power-distribution modules with

built-in voltage and current protector?

GOMSpace P31us power distribution module (http://gomspace.com/index.php?p=products-p31us).

Clyde Space FlexU power distribution module (https://www.clyde.space/products/19-3rd-generation-flexu-eps).

Blue Canyon Tech XEPS electrical power sub-system module (http://bluecanyontech.com/wp-content/uploads/2015/05/EPS-Data-Sheet_1.0.pdf).

OPTION 1:

3 Solar Panel Inputs for

maximum of 30 W.

Voltage: 3.3 V and 5 V.

OPTION 2:

3U to 12U deployable solar

panels

Voltage: 3.3 V, 5 V and 12 V

OPTION 3:

6 Solar Panel Inputs for maximum of 60 W.

Voltage: 3.3 V, 5 V and 12 V. 20Wh/40Wh battery pack.

COTS Options for Solar Panel Configuration

ESTIMATED NO. OF

PANELS AND AREA

FOR 44 W:

20 panels. 1207.2 cm2.

Solar Panel Configuration Constraints:

1. DWTS shall point cross-

track.

2. DWTS shall avoid pointing

directly at the sun.

PROBLEM: Zero flux at 0 degree beta-angle.



track.



SOLUTION 1: Solar panel drives.



track.



SOLUTION 2: Additional orthogonal solar panels.



track.



SOLUTION 3: Periodic charging mode. Reduced payload duty cycle

(off during eclipse).

SOLUTION 1 SOLUTION 2 SOLUTION 3 1. Added mass.

2. Additional costs.

3. Additional attitude

control.

4. Additional electronics.

5. Driver malfunctioning.

1. Added mass.

2. Additional costs.

3. Additional attitude

control.

4. Deployment risk.

1. Frequent slewing.

2. Reduced data

coverage.

Test Objectives

• Mission shall demonstrate ability of DWTS to

measure winds and temperatures from 30 –

200 km.

• DWTS shall be demonstrated in a space

environment.

• Demonstrate that stray light effects can be

eliminated using optical or digital means,

including lunar scanning to demonstrate stray

light calibration.

• The satellite shall resolve planetary-scale

waves and tides with zonal wave numbers 1

to 4 from the lower to the upper atmosphere

from 60S to 60N.

Mission Objectives (Revised)

Will encounter 0 solar flux every 90° of orbital

precession.

Total power budget, solar panel requirement and battery capacity

requirements are estimated. COTS options for space-grade and

prototype solar panel, battery and power distribution modules are

identified.

Options on solar panel configurations that can support the

operational test of a single-channel DWTS are identified.

Mission objectives must be revised such that they can be

achieved at the expense of reduced data coverage.

Will a 6U CubeSat be able to support the DWTS

payload and mission?

Risks Identified

Subsystem EPS ADCS Thermal

Risk High power requirements

No Sun at certain beta angles

Attitude knowledge and control

requirements

Expected temperature range exceeds

component survival range.

Mitigation

Options

• Charging mode with reduced

payload duty cycle

• Steerable solar panels

• Perpendicular fixed arrays

• Blue Canyon XACT ADCS

module

• DWTS cryocooler duty cycling

to characterize disturbance

torque smearing.

• DWTS radiator pointed at deep space.

• Component level thermal modelling.

• Passive thermal control components

for non-payload subsystems.

Steady State Temperature Range (600

km):

-69.4C ~ 50.1C

Components Operational

EPS -40 ~ 85C

Batteries 5 ~ 20C

Solar Panels -150 ~ 110C

DWTS -30 ~ 20C

GPS -40 ~ 85C

ADCS -10 ~ 40C

Motherboard -40 ~ 85C

Antennas -100 ~ 100C

DWTS Radiator -20 ~ 0C

�

� = � �°

� = �°

�

Pointing

Knowledge ±0.5 arcmin,

all axes

Pointing

Stability < 6

arcsec/sec

Attitude

Control 1° all axes

Feasibility & Iteration Revised mission objectives can be

achieved with a 6U CubeSat at the

expense of reduced data coverage.

Payload design must be revised, with

laboratory testing of in-orbit calibration

methods to reduce risk.

Additional details:

Chang, L.C. et al. (2016), A Preliminary Design for the

INSPIRESat-1 Mission and Satellite Bus: Exploring the

Middle and Upper Atmosphere with CubeSats,

Proceedings of the AIAA/USU Conference on Small

Satellites, LEO Missions, SSC16-WK-02.

http://digitalcommons.usu.edu/smallsat/2016/S4LEOMis

/5/.

14

The DWTS optical design is shown in Figure 13. This design supports the GEOSNAP detector

form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal length, 30 mm

aperture, resulting in a fast F/1.4 system. The lens includes mixture of Germanium and ZnSe

lenses as required for stray light control and image quality. Both spherical and aspherical

elements are used, as the optics shall be diamond turned to the required surface figure and

finish. The optics will have low-emissivity anti-reflection coatings to keep internal emissions

low. The inside surface of the lens barrel and the sensor housing will be coated black to reduce

reflected stray light.

Thermal Design: The thermal design has two actively-controlled thermal zones, the detector,

which must be kept at 95 K (110K is acceptable) to reduce dark current to our goal of

approximately 100 e-/pixel/s, and the cryocooler radiator which rejects heat at 23 C. In

addition, the cryocooler will use passive cooling to keep the optics 210 K to reduce thermal

emission. A Ricor K527 Stirling split-linear cryocooler will be used to cool the detector. The

K527 has a cooling capacity of 0.7W for an active

load at 95K and requires 20W input power. A

flexible thermal strap will be used to connect the cryo

cold head to the detector, providing vibration

isolation from the cryocooler pistons. A large

deployable radiator will be used to reject the heat

generated by the cryocooler. Precise control of the

temperature in both zones will be accomplished via

heaters placed on the detector housing and on the

radiator. The detector will be mounted to the lens

assembly using thermally-isolating mounts, and then

both are enclosed inside a double walled thermal

enclosure. The lens assembly will be mounted to the

enclosure also using thermal isolating mounts. To maintain radiative isolation, the outside

surface of the lens barrel and the inside surface of the enclosure walls will be gold coated.

Figure 12 shows the instrument concept thermal model. End of Life (EOL) material properties

were used for hot case analysis and Beginning of Life (BOL) properties were used for cold case

analysis. In the worst case hot orbit scenario, the cryocooler active load is 0.54 W, which has

30% margin to the K527 capacity of 0.7W. Additional margin could potentially be gained

through detailed design of the isolating mounts and increased detector flex cable length. The

maximum predicted heater power required to maintain the detector and radiator at the design

temperatures are 0.25W and 16W respectively. There are also maneuvers that can be

performed, such as 180 degree role about the line of sight (effectively flipping the FOV, which

does not affect the measurements), that will dramatically improve radiator and solar array

efficiency. This strategy was not fully evaluated before the proposal, but will conservatively

add 30% to our cooling margins.

Figure 12. Thermal Model of Sensor and Platform.

Revised Mission Timeline

19

3.2. Project Schedule

The DWTS top-level schedule is shown in Figure 16. The project will start on January 21,

2017 on receipt of funds. The project will be part of the graduate course in Spacecraft Design in

Spring 2017. The schedule balances design and review efforts with the need to order critical,

long lead, items early enough to preserve a robust system I&T effort. The Preliminary Design

Review (PDR) will be completed, seven months after program start on August 20, 2017 and

The Critical Design Review (CDR) is scheduled five months after PDR. Since many subsystems

are commercial off the shelf (COTS) parts and the XB1 spacecraft bus comes in a defined

package, the procurement of long lead parts (such as solar arrays, XB1, detector from Teledyne)

begins after PDR.

Figure 16. DWTS Top-Level Schedule

Three months are allocated to mechanically integrate system, make any necessary

mechanical adjustments, and perform initial telemetry

checks and verify subsystem handshaking. Following

system integration, two months are allocated to vet

operational test scripts, perform a CPT, and perform an

end-to-end test from the MOC to the spacecraft through the

ground test equipment. Environment tests proceed after

baseline compatibility and performance testing is complete

on the integrated flight hardware with two

3.3. Risk Reduction Plan

Because of their limited on-board resources, modest

development budgets, and commitment to involving

relatively inexperienced student teams in their design and

implementation, CubeSat projects present a higher level of

risk than would be acceptable for most space flight

missions. However, the overall approach to continuous risk

management used by larger programs applies equally to

small ones: identification of risks, categorization by

likelihood and consequences, planning of mitigation strategies, and tracking of progress

Color Code. Green=acceptable w/o

further mitigation; Yello=

mitigation necessary, acceptable at

launch; Red=risk drivers, aggressive

mitigation need, mitigate before launch.

Additional Lessons Learned (So far)

• For a payload defined mission, freezing payload requirements is essential before the spacecraft bus can proceed to PDR levels.

• Project management and standard operating procedures are essential and should be strictly defined, especially for international collaborations.

• Good communication and design iteration is essential between the science and engineering teams.

• Never underestimate the value of good PR for securing support and team building!

INSPIRESat-2 / Compact Ionosphere Probe

• Schedule: • Payload and bus PDR: 12/2016

• Payload and bus CDR: 8/2017

• Funded by Taiwan MOST, University of Colorado.

STR

Requirements

● All the structures shell not cause Interference.

● Mass budget shell be less than 4kg, and mass margin is +2kg.

Assembly and Interior layout

3U CubeSat, size is 474 x 1113 x 337mm with deployed antennas. Total mass is

2.598kg without harnesses, screws and GPS antenna. It is not possible to reduce

the size from 3U to 2U because the mass will overload for 2U structure. CubesatAIP

has total 2855 parts.

CDH

CDH Requirements

18

INSPIRESat-2: Subsystems Breakdown Subsystem NCU CU IIST

Payload

STR

TCS

CDH

COM

EPS

ADCS

Software

Launch

Thank you. Questions?

Documents

Microsatellite Design and Integration for INSPIRE · DWTS Design & Interface 15 draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum electrical