View
220
Download
0
Category
Preview:
Citation preview
Drag and Atmospheric Neutral Density Explorer(DANDE)
Colorado Space Grant Consortium and
CU Aerospace Engineering Sciences
University Nanosat 5 LASP Seminar
October 16th, 2008Boulder, Colorado
2
DANDE LASP Seminar
Overview
1. Introduction2. Science3. Systems Engineering4. Electronics5. Structure, Separation, Thermal6. Integration, Testing, Schedule
3I - Introduction
The University Nanosat Program
• University Nanosat – The National Championships of Spacecraft Design– 2 year program in its fifth iteration– 10 out of 30 university proposals selected based on Air
Force Relevance– $85k initial seed funding for hardware and student
support– In January 2009, one school wins additional $85k, I&T at
Kirtland, and flight to Orbit• CU Nanosat Entry
– Has involved a core team of graduate students and expanded into 40 graduate and undergraduate students
– Many aspects of the ASEN Graduate Projects but organized as independent research and MS research
– Has leveraged over $240k from University, Department, DoD, and COSGC Funds
4I - Introduction
CollaborationsUNP AFSPC/A9A AFOSR AFRL NOAA
Graduate Student TeamSubsystem Lead Engineers
Undergraduate StudentEngineers
STUDENTS
PI’sC. Koehler, S. Palo, J. Forbes
Gov. & Academic Support Industry Support
5I - Introduction
Operational Importance of Drag
The density of the atmosphere in this region varies greatly (300% to 800%*) due to space weather and not yet understood coupled processes.
* Forbes et. Al. “Thermosphere density response to the 20-21 November 2003 solar and geomagnetic storm from CHAMP and GRACE accelerometer data”, Journal of Geophysical Research, Vol. 111, June 2006
drag induced drift
Relative Orbit of Two Separating Spacecraft410 km
390 km
370 km
350 km
330 km
1999 2000 2001
reboost
decay
X5 flare7/14/2000
ISS drops 10kmin several days
E. Semones et.al. WRMISS 10 http://www.oma.be/WRMISS/workshops/tenth/pdf/ex02_semones.pdf
6I - Introduction
Scientific Importance of Drag
Satellite drag measurements suffer from errors caused by
• Unknown acceleration contribution from in-track winds
• coefficient of drag accuracy
CHAllenging Minisatellite Payload
Starshine I
7I - Introduction
DANDE will measure density, composition, and wind along its orbit
The DANDE Analogy
:
:
as
DANDE
8I - Introduction
Introduction to DANDE
Mission StatementExplore the spatial and temporal variability of the
neutral thermosphere at altitudes of 350 - 200 km, and investigate how wind and density variability translate to
drag forces on satellites.
DRAG and
ATMOSPHERIC
NEUTRAL
DENSITY
EXPLORER
9I - Introduction
Nanosat V Program at CU
DANDE will improve atmospheric models and calibrate near real-time models by measuring the following
•Deceleration
•Atmospheric composition
•Horizontal Winds
DANDE is spherically shaped to minimize biases resulting from estimation of the drag coefficient
1010
Science
Marcin Pilinski
11II - Science 11
How Measurements are Made
• Identifying all components of the constituents of the drag equation.• With a near-spherical shape, an a-priori physical drag coefficient may be calculated and a physical density can
be obtained from the measurements
atmosphere
ρ - densityV
A
FD
CD
VW
aMVVACF WDD 2
2
1
accelerometersWATS sensor
a priori knowledge
tracking
a priori knowledge
a priori knowledge comparison
solutionmeasured
a priori
solution
solved
12II - Science
Accelerometer Measurement System
ANALOG FILTERING
A/D CONVERSION
LEAST SQUARES
70 ng1x100 1x1021x10-31x10-5
1.6x10-10
4.0x10-15
Frequency [Hz]
PS
D [
g2 /
Hz]
1.0x10-12
1x100 1x1021x10-31x10-5
1.6x10-10
4.0x10-15
Frequency [Hz]
PS
D [
g2 /
Hz]
1.0x10-12
spin rate
Low frequency bias
13II - Science
R
T
Accelerometer Analysis
ACC-4
ACC-6
ACC-3 ACC-
2
ACC-
5
ACC-1
FD
PROCESS & AVERAGE
ω
AC
C-6
AC
C-2
AC
C-5
AC
C-1
AC
C-3
AC
C-4
ω = π/3 [rad/sec]
14II - Science
Accelerometer AnalysisLatitude [deg]
0∘ 82∘ -82∘0∘ 0∘ 82∘ -82∘0∘ 0∘ 82∘ -82∘0∘ 0∘ 82∘ -82∘0∘
15II - Science 15
Neutral Mass Spectrometer (NMS)INCOMING NEUTRAL
DISTRIBUTION
INSTRUMENT FOCAL POINT MCP
ANODES
ION DISTRIBUTION ELECTRONDISTRIBUTION
channel 8
channel 9
channel 10
.....
channel 11
channel 12
channel 7
channel 2
channel 1
16II - Science 16
Neutral Mass Spectrometer
1.Neutral particle (blue) enters the collimator. (Ions rejected)2.Neutral particle is ionized inside of a field free electron bombardment region3.Neutral particle enters the energy selector and undergoes acceleration towards the
exit 4.Outside the selector, the particle is accelerated abruptly by a -3kV potential towards
the Micro-Channel Plate (MCP)5.The impact on the MCP causes a cascade of electrons to travel towards one of the
anodes which measures the impact. Which anode is triggered depends on the angle at which the neutral particle entered the collimator.
17II - Science 17
NMS Science Data Product Analysis
wind angle
N2 wind mag.
O wind mag.
O temp. N2 temp.
18II - Science
Density Error – Drag and Wind Data
normally distributed zero-mean wind measurementsRMS= ±30 m/s
19II - Science
NMS Status and Future Work
• First round of testing scheduled for November 17th-21st at NASA, Goddard
• NMS structure built at LASP
2020
Systems Engineering
Mike Grusin
21III - Systems Engineering
DANDE Overview
ESPA Ring DANDE Sphere
Lightband Adapter Bracket (LAB)
Baseline Configuration
18”
22III - Systems Engineering 22
Mission Timeline• Phase 1: LV Separation and
commissioning1. Launch Mode - time delay –
Safe Mode2. Full charge and checkout
[18 – 30 hours]3. Lightband jettison
• Phase 2: Attitude Acquisition1. Spin Up [24 h]2. Spin-Axis Alignment [120h]3. Reserve time [24h]
LV SEPARATION AND COMMISSIONING PHASE
Day 2Day 1
Wind
Composition
Acceleration
Tracking
Tracking
SCIENCE PHASE
DATA ACQUISITION1 orbit SCIENCE1 orbit STANDBY
DOWNLINK/UPLINK~2x in 24 hours
ATTITUDE ADJUST~1 orbit per day
RE-ENTRY DYNAMICS~LAST WEEK OF ORBIT
200 km – 100 km
Day 9 Day 100
• Phase 3: Science [~90 days]1. Science Mode2. Standby Mode3. Comm. Pass4. Attitude Adjust5. Repeat
23III - Systems Engineering
Ligh
tban
d as
sy.
Functional Block DiagramFunctional Block DiagramWiring Harness
LV e
lect
rical
in
terfa
ce
FOV 32° x 1.8°
NMS
WATS instrumentCSGC / GSFC
ControlATmega128
Atmel
DataAcquisition HV sources
CDH SFT
I2C busRTC
Linux OSModeManager
COM proc
ADC proc
ACC proc
NMS proc
RS-232
CPUAVR32
120MHzAtmel
32 MB SRAM
64 MBCF CARD(DATA)
8 MB NOR FLASH (SFT)
EPS
ControlATmega128
FOV360°
Photovoltaics 30W
BatteryB
12V 4AH
Regulation
Inhibit x4
Inhibit x4
Battery A12V 4AH
Charge
Sate
llite
Sep
Plan
e (S
SP)
COM
TNCSymek
38.4kbpsmodemAm
pSy
mek
7W
Switched power
Critical power
Tx AntRx Ant
FOV360°
TransmitterSymek
70cm (436MHz) 38.4kbps
ReceiverHamtronics
2m (150MHz)9.6kbps
9.6kbpsmodem
FOV+/-80°
off nadir
LightbandAdapterBracketassy.
ACC
ControlATmega128
Atmel
AccAcc
Acc
Acc
AccAcc
Analog
SEPLSRM1SpaceDev
LSRM2SpaceDev Re
leas
e se
nsor
sHOP1
SpaceDev
HOP2SpaceDev
ACC
ControlATmega128
Atmel
AccAcc
Acc
Acc
AccAcc
Analog
NMS
WATS instrumentCoSGC / GSFC
ControlATmega128
Atmel
DataAcquisition HV sources
CDH SFT
I2C busRTC
Linux OSModeManager
COM proc
ADC proc
ACC proc
NMS proc
RS-232
CPUAVR32
120MHzAtmel
32 MB SRAM
64 MBSD CARD
(DATA)
8 MB NOR FLASH (SFT)
SEPLSRM1SpaceDev
LSRM2SpaceDev Re
leas
e se
nsor
sHOP1
SpaceDev
HOP2SpaceDev
COM
TNCSymek
38.4kbpsmodem
Pwr A
mps
Sym
ek 7
Wx2
Switched power
Critical power
Tx AntsRx Ant
FOV360°
TransmitterSymek
70cm (436MHz) 38.4kbps
ReceiverSpaceQuest2m (150MHz)
9.6kbps
9.6kbpsmodem
FOV+/-80°
off nadir
EPS
ControlATmega128
FOV360°
Photovoltaics 30W
BatteryB
12V 4AH
Regulation
Inhibit x4
Inhibit x4
Battery A12V 4AH
Charge
2323
ADC
ControlATmega128
Atmel
X torque rod
Y torque rod
MagnetometerHoneywell HMR2300
Passive nutation damper
FOV 2° FOV 2°90°
Horizon Crossing
Indicator AServo
Horizo
n
Crossin
g
Indica
tor B
Servo
THM Coatings, InsulationSensors
THMSensors
Coatings, Insulation
24III - Systems Engineering
Inside of DANDE
battery box (x2)
mass trim system (x8)
3-axis magnetometer
wind sensor
EGSE connector
Stiffeners (x4)
lightband adapter bracket
ball & tube nutation dampener (x2)
EMI box (x3)
horizon crossing indicator (x2)
patch antenna (x3)
Accelerometers (x6)
separation mechanisms (x2)
kinematic mounts (x4)
25I - Objectives and Requirements 25
–Spin stabilization about orbit normal• 40°/sec (10 rpm)• Only two maneuvers:
spin-up and axis alignment
–Sensors• Magnetometer for spin-up• Horizon Crossing Indicators for
spin axis alignment
–Actuators• 2x Torque rods: one along spin axis
and one transverse• Passive nutation damper
DANDE Attitude
2626
Electronics
Brandon Gilles
27IV - Electronics
Communications Overview
Ground Station
CDHSYMEK
TNC31S
RS-232MODEM
38,400 TX
9,600 RX
AUDIOSYMEK
TX
SPLITTER
(tbd)
S’GART
POWER
AMP
S’GART
POWER
AMP
70cmPATCH
ANTENNA
70cmPATCH
ANTENNA
LOWPOWER
RF
LOWPOWER
RF
HIGHPOWER
RF
2mPATCH
ANTENNA
YAESU
FT847
SYMEK
ICD MODSYMEK
TNC3
MODEM
38,400 RX
9,600 TX
RS-232PC
70cmYAGI
2mYAGI
Spacecraft
AUDIO
HIGHPOWER
RF
LOWPOWER
RF
SPACE-QUEST
RX
28IV - Electronics
Communications Testing
• Partnership with First RF Corporation• Concept:
– ¼-wave patch antenna – Spacecraft acts as ground plane
• Transmit Antenna Pair– TX: 1 x 3.72 inch – Gain: -3.5dBi– Beamwidth: 240º
• Receive Antenna (estimated)– RX: 1” x 13”– Gain: -20dBi– Beamwidth: 240º
Patch Antenna
FIRST RF Corporation
28
29IV - Electronics
Ground Station Communications
• Antennas– 70cm: 38-element Yagi– 2m: 22-element Yagi
• Transceiver– Yeasu FT-847– RF Power output variable up to 50W– IF-Amplifier/Demodulator (IFD)
Upgrade for bandwidth• TNC
– Symek TNC31S– (2) FSK9601 Modems
• 1 x 9600 baud• 1 x 38400 baud• Wired for 9600 up and 38400 down
29
30IV - Electronics 3030
Command & Data Handling Architecture
31IV - Electronics 3131
Software Architecture
32IV - Electronics
Power System Overview
+
_
Battery B12V
4000mAh
+
_
Photovoltaic arrays24 watts, 15 volts
+/-15V
+5VX 3
X 1
Battery A12V
4000mAh
X 3
X 1
Solid-State Relays
Ground lines to subsystems
Single-pointChassis ground
Main Bus, +12VUnregulated output
Regulated outputsLatchingrelays
Latchingrelays
DC-DCConverters
InhibitsInhibitsBatteryCharge control
Miniature PV strings. 15 V for direct-energy transfer
3333
Structure, Separation, Thermal
Andrew Tomchek
34V - Structure, Separation, Thermal
Separation System
3D: A sphere resting inside a cup• Restricts translational motion in the A, B & C
directions at one point.2D: A sphere resting inside a trench• Restricts translational motion in two directions
at one point.• When combined with 3D mount, restricts
rotational motion in the A & C axis. 1D: A sphere on a flat plane• Restricts translational motion in the C direction• When two are combined with the 3D & 2D
mounts, restricts rotational motion in the B axis.
Material Selections:• Male: Al 7050-T745 PER AMS 4050 • Female: S15500 Stainless Steel• Male material is softer than female, reducing
stiction.• Materials have flight history with kinematic
mounts
3D
2D
1D 1D
35V - Structure, Separation, Thermal
Evolution of DANDE
36V - Structure, Separation, Thermal
Manufacturing
Advised by Tim Flaherty and Mathew Rhodes
37V - Structure, Separation, Thermal
Structural Analysis and Testing
•Y-Direction-Shows stress through internal X
•Vibe test–Demonstrated Structural Integrity–Low natural frequency (89 HZ)
–Trade studies to increase stiffness
38V - Structure, Separation, Thermal
Thermal
• Currently working on trade studies to improve hot and cold cases.
Operating Allowable
(°C)
Hot Case
(°C)
Cold Case
(°C)
ACC -40 to + 80 53 -18
CDH 0 to +70 44 -17
COM 0 to +50 58 -16
Solar Cells -85 to 90 40 -35
Mechanism -53 to +71 41 -20
Batteries +5 to +45 48.5 -13
EQ Plate -- 42 -24
NMS -55 to +75 45 -19
Advised by Jenny Young
3939
Integration, Testing, Schedule
Bruce Davis
40VI - Integration, Testing, Schedule
State of the DANDE Program: Hardware
HARDWARE & MANUFACTURING
Engineering Design UnitLessonsLearned Competition Review Hardware
41VI - Integration, Testing, Schedule
Integration Planning – Wiring Harness
COM
RC ANT
TX ANT
TX ANT
HEM+X
HEM-X
ACC CDH
TOR
EGSE
SYS241
BATT37-pin
37-pin
50-pin37-pin
15-pin
25-pin 25-pin
NMS25-pin
9-pin25-pin
25-pin
LVI
15-pin
SEP
SYS242
SYS243
SYS244
SY
S2
45
SYS246
SYS247
SYS248
SY
S2
49
SYS251
SYS250
SYS252
SYS254SYS253
PA 1
PA 2
Spacequest Receiver
Symec TNC
AVR
9-pin
15-pin
SYS255
HCI
LV ABSSYS256
SYS257
50-pin37-pin
SYS258
9-pin
9-pin
W2-pin,W4-pin
TOR
W4-pin?
W4 pin
EPS(Inhibit, Regulator, Control
Boards)
50-pin
25-pin
25-pin
50-pin37-pin50-pin
50-pin
50-pin
W4 pin
BATT
Symec Transmitter
Splitter
W2x2-pinW3x4-pin
SEP
MAG9-pinW4-pin?
W2-pin,W4-pin
257 258
17
241
250
247
248
248247
243
4
243
250
15
42VI - Integration, Testing, Schedule
I&T Planning – Box Integration
• Key Box Design Features– All connectors interface to one surface– PCB boards stack removed easily– Wiring harness flexibility with the
fabrication of a new interface board
Interface board routs electrical lines (contains no circuitry)
PC-104 connectors, placed adjacent to hex standoffs
PCB Stack connected to lid
Standoffs free-fit into box surface, improves rigidity
43VI - Integration, Testing, Schedule
I&T Planning – Hemispheres
• Unique Design Problem, Unique Challenges– Risk to the launch vehicle– Optimization of Assembly Time (1300 cells)– Ease of Integration / De-Integration– Utilizing Existing Facilities– Cleanliness / Quality Control– Protective Ground Support– Life Testing
• Solutions– Build a full hemisphere for practice / testing
by Oct 31st
– Use crimp-pins to allow for ease of removal– Extensive trade study for optimal cell-to-
PCB attachment– Backup plan established with traditional
methods
44VI - Integration, Testing, Schedule
I&T – Upcoming Formal Tests
• Roughly 600 formal requirements– Verified through subsystem testing on the
flight hardware– 20% currently verified– 50% excepted to be verified on hardware by
January ‘09• 30 planned tests throughout the fall
semester– Documented with as-run test procedures
Sample of Upcoming Tests:ACC707 - Accelerometer Flight Board Acpt TestADC710 - Torque Rod Functionality TestCOM709 - Long Range Antenna TestEPS705 - Power Board Inhibit Functionality TestNMS702 - NMS Calibration Test (at Goddard)STR710 – MGSE Proof Loading Test
As-Run Test Procedure Example
Test Report Example
45VI - Integration, Testing, Schedule
Completed Formal Testing
TESTING
46VI - Integration, Testing, Schedule
Integration & Testing Schedule
Today Competition Review
47VI - Integration, Testing, Schedule
Tall Poles
• Resolved Risks (selected)– Personnel Turnover DANDE recently doubled in size from other COSGC Projects– Solar Cell Effects on Drag Detailed study shows surfaces can be characterized– Center of Gravity Management Analysis & reserved mass allocated for CG authority
• Current Risks (selected)– Solar Cell Attachment Mitigated by testing, practice and alternative fabrication
options– Link Budget / Antenna Testing Mitigated by meeting with industry advisors,
re-definition of requirements, de-scope options identified.– Orbital Envelope Mitigated by re-definition of requirements allowing for elliptical
orbits. Currently matches 32 of 68 previous military launches (with transfer orbit). Drag chute senior design project ’07-’08 trade study.
– Thermal Extremes Mitigated by meeting with industry advisors, identifying trade studies, re-definition of orbital requirements, possible addition of heaters.
48I - Objectives and Requirements
LASP Contributions to DANDE
LASP has contributed to DANDE for the last two years
– Advisement to students– Heritage design references (SNOE)– Sample of procedures / operations
LASP has recently become a formal supporter– Established cost effective machining
opportunities for precision parts– Supplied materials to stock our cleanroom
Thank You to those at LASP who have made this possible:
– Tim Flaherty– Caroline Himes– Mike McGrath– Norm Perish– Ed Wullschleger– Jenny Young Thank You
49
Questions
dande.colorado.edu
5050
Backup Slides
5151V - Analyses 51
Radiation Analysis
COTS Limit*
Inputs into CREME96• 100 mils of aluminum shielding• 90 degree inclination (TID worst case)• Near solar max conditions (likely for
nominal orbit)
Results• TID as a function of orbit altitude for
400-day orbit– Max circular orbit of 1080 km – Part life as function of orbit altitude– Enough margin for a 350x1200km
elliptical orbit
*NASA Public Lessons Learned Entry: 0824
52
System Driving Requirements I
Ref. Description Parent Ref.
Verification
1.SYS1 The system shall measure densities, compositions, and winds in an altitude range of 350 km +20 km -250 km and over latitudes of 54° or higher as verified by the system budgets and thermal and radiation analysis.
0.SYS1, 0.SYS2
Thermal and radiation analysis as well as system budget analysis. Functional testing. Attitude system analysis and testing.
• Defines science data products.• Constrains perigee of orbit to the region of scientific interest• Constrains inclination
Ref. Description Parent Ref.
Verification
1.SYS77 The system shall measure densities, compositions, and winds in an altitude range defined in 1.SYS1 over a horizontal range of no less than 3000 km.
G1 Analysis and testing of child requirements
• Defines minimum along-track distance for orbit to intersect region of interest• Driven by interest in large scale disturbances (Q1, Q3, and Q6)
53
System Driving Requirements II
Ref. Description Parent Ref.
Verification
1.SYS4 The system shall make density, in-track wind and cross-track wind, and composition measurements during five 4 ± 1 hour periods following each observed SSC and during four 4 ± 2 hour quiet geomagnetic condition periods. (goal: 100 hours of data collection during both geomagnetically quiet and stormy conditions in the altitude range of 350 to 150 km).
0.SYS1 Functional testing and budget analysis.
• Constrains minimum science-mission lifetime and minimum operating periods
Ref. Description Parent Ref.
Verification
1.SYS5 The system shall make density, in-track wind and cross-track wind, and composition measurements of the fidelity described in 1.SYS2 and 1.SYS3, 1.SYS7, 1.SYS21, 1.SYS57, 1.SYS56, 1.SYS57 at 60 ± 4 second intervals (goal: 40 second intervals) for at least 50% (one orbit on, one orbit off) of the measurement cycle described in 1.SYS4 (goal of 80%)
0.SYS1 Functional testing and budget analysis.
• Defines minimum cadence for atmospheric sampling (motivated by model-grids)• Defines minimum success science-mode duty cycle
More on these in a later slide
54
System Driving Requirements III
Ref. Description Parent Ref.
Verification
1.SYS10 The design orbit lifetime of the density measuring spacecraft shall not be less than 90 days from beginning of science measurements until re-entry.
0.SYS1, 0.SYS3
Area and mass values of the spacecraft are used in an orbital simulation to determine the orbit lifetime.
• Helps define range of apogees given assumed densities and historical storm conditions• Calls out minimum mission lifetime
Ref. Description Parent Ref.
Verification
1.SYS38 All sub-systems shall conform to the System Modes and Design Reference Mission (DRM, also known as the Concept of operations, OPS108)
0.SYS5 Verified by inspection and day-in-life testing
• All subsystems must design to the same mission• Defines operational modes and subsystem duty cycles (see design section)
55
System Driving Requirements IV
Ref. Description Parent Ref.
Verification
1.SYS53 The system shall measure coefficient of drag to a fidelity defined in 1.SYS52 and 1.SYS72 over every 20 ± 5 km change in altitude over an altitude span of 350 ± 20 km to 250 ± 20 km. Goal: the system should measure coefficient of drag to a fidelity defined in 1.SYS52 and 1.SYS72 over every 10 ± 3 km change in altitude over an altitude span of 350 ± 20 km to 100 ± 20 km.
0.SYS2, PO4
Verified by orbit determination analysis
• Defines minimum and goal ranges for coefficient of drag capture• Capturing drag data down to ~100km is a goal, not a minimum mission requirement
Ref. Description Parent Ref.
Verification
1.SYS76 The spacecraft shall spend no more than 300 days above the minimum perigee requirement of 350 km altitude.
0.SYS5 Functional testing and budget analysis.
• May wait up to 300 days before desired altitude range is achieved• Low cost drives (0.SYS5) use of COTS components, this caps the amount of time spent at high
altitudes and limits the TID received• Drives design of optional aerobraking system
56
Refs Requirement Precision(1-sigma)
Accuracy
absolute percent* absolute percent*
1.SYS2, 1.SYS52 Density 2x10-13 kg/m3 2% 1.0x10-12 kg/m3 10%
1.SYS6, 1.SYS7 In-Track Wind 100 m/s 20%** 100 m/s 20%
1.SYS56,1.SYS57 Cross-Track W. 100 m/s 10% 100 m/s 10%
1.SYS21
Composition Densities (O & N2) 7x1013 m-3 2% 5.3x1014 m-3 15%
1.SYS52,1.SYS72 Coefficient of Drag 0.1 5% 0.2 10%
*percent value based on average conditions during solar maximum, vernal equinox
**assuming a wind velocity of 1 km/s, storm conditions
Key Measurement Requirements
1.SYS26 , composition measurements with resolution of 1.5 m/Δm. Driven by 0.SYS1 and 1.SYS21(where m/Δm = half peak width at mass m)
horizontal resolution of 500km (~64s)
57
Derived Orbital Requirements
• The orbit segment used for science passing at or below 350 km altitude shall be at least 3000 km long.
• The altitude shall not vary by more than ±40 km during that segment
• Apogee no higher than 1200 km altitude
• Initial apogee and perigee altitudes such that the orbit lifetime is no shorter than 100 days
• Goal: Initial apogee and perigee altitudes such that the orbit lifetime is no longer than 400 days
58
Derived Orbital Requirements
Parameterizing the orbital requirement results in the following acceptable orbital envelope
Required inclination range of 54º - 126º
Earth
0
200
400
600
800
1000
1200
1400
0 50 100 150 200 250 300 350 400
Perigee Altitude [km]
Ap
og
ee A
ltit
ud
e [k
m]
(D) orbital lifetime (C) Radiation Envioronment (B) 3000 km horizontal observibility (D) 400 day max lifetime goal
Goal Envelope: observation of altitudes <250km within COTS components lifetime:Minimum Required Envelope:
59
System Driving Requirements V
Ref. Description Parent Ref.
Verification
1.SYS14 The physical coefficients of drag (CDP) of the spacecraft shall not vary during tracking or acceleration measurement cycles by more than ±3% from the pre-determined (1.SYS11) values. (goal: 1%)
0.SYS1, 0.SYS3, 1.SYS11
Inspection of the spacecraft shape and analysis.
1.SYS16 The cross sectional area of the spacecraft shall not vary during tracking or acceleration measurements by more than ±1% from the pre-determined (SYS13) values. (goal 0.1%)
0.SYS1, 0.SYS3, 1.SYS13
Verification will consist of measurements of the spacecraft dimensions.
• Together these drive the system to be a spheroid during measurements• High Accuracy Drag Model (HASDM) evaluation also drives spheroidal shape and is driven by
0.SYS2
60
System Driving Requirements VI
• Position and velocity accuracy as driven by the science and ADC requirements• Can be accomplished with tracking
Ref. Description Parent Ref.
Verification
1.SYS18 The spacecraft velocity shall be determined to ±30 m/s in magnitude and ±0.5 degrees in direction with respect to RTN to one-sigma precision.
0.SYS1, 0.SYS3
Error analysis and orbit determination analysis given the expected tracking errors.
1.SYS19 The position of the density measurements shall be determined to 10 km (in ECI) in all axes to one-sigma precision (goal 1 km in each ECI axis).
3.ADC7, 0.SYS1
Error analysis and orbit determination analysis given the tracking errors expected.
61
System Driving Requirements VII
• Define overall knowledge and pointing requirements• Driven by the science error budget
Ref. Description Parent Ref.
Verification
1.SYS25 The spacecraft body axes shall be determined to ±2º precision (1-sigma) with respect to the ECI reference frame during science measurements.
0.SYS1, 0.SYS3
Worst case attitude determination analysis. ADC hardware testing.
1.SYS44 The spacecraft body axes shall have an offset from their nominal ECI vector of no more than [±5º] during measurements. (this is a pointing requirement)
0.SYS1, 1.SYS1, 1.SYS6, 1SYS21
Verified my mass property measurement and attitude simulation with inputs from attitude system performance tests.
62
System Driving Requirements VIII
• Timing as well as time transfer requirements• Note that these do not require the data to be simultaneous but it defines the relative drifts between
time stamps• Flows down to operations and command and data handling
Ref. Description Parent Ref.
Verification
1.SYS60 The time of the space sector density, composition, and flux measurements shall coincide within ±5 millisecond precision of the system clock-time. (precision of time associated with data)
0.SYS1, 1.SYS1
Functional Testing
1.SYS63 The system clock-time shall be accurate to within ±1 second of UT. (accuracy of time associated with data)
0.SYS1, 1.SYS1
functional testing
63
ACC: Primary Design Problem
63III - System and Subsystem Design Overview
Problem: The raw accelerometer output is not directly useful because it is masked by noise
Goal: To isolate and amplify a signal from a significantly (>3 orders of magnitude) larger noise environment
64
Design Updates
solar substrate design finalizedantenna design:
substrate changed to Ultemsubstrate flat
2 receive and 1 transmit patch
RX
TX
TX
65
Multi-Instrument Analysis
Present satellite drag measurements result in densities with large errors at higher latitudes, this is primarily due to unmeasured winds.
66I - Introduction
DANDE Research Involvement
• University: Dr. Scott Palo, Dr. Jeff Forbes,
• AF: Bruce Bowman, Dr. Chin Lin, Dr. Frank Marcos
• ONR: Andrew Nicholas, Dr. Robert McCoy
• NASA: Dr. Fred Herrero
• NOAA: Dr. Tim Fuller-Rowell, Dr. Mikhail Codrescu
6767
NMS Science Data Product Analysis
Total number densities across all spectra as the satellite spins
-180
-155
-130
-105-80
-55
-30-520457095120
145
170
050
010
0015
0020
0025
00
WACC-4
ACC-6
ACC-3 ACC-
2
ACC-
5
ACC-1
ω
Angular position about the satellite spin axis, degrees
# of
par
ticle
s im
pact
ing
dete
ctor
Peak count of vertical distribution
~2000 counts
R
T
68
Coefficient of Drag Analysis
Multiple reflections are prominent inside the concave geometry and at lower incidence angles
The effect of facets including multiple reflections on CD is on the order of -2% for DANDE
Reducing the α proportionally to the reduction of flux of O causes a change in CD of +7% for Starshine I (Simulation not yet run for DANDE)
Locations of multiply reflected particles on a faceted geometry
69
Problem Description: Satellite Drag
wV
a
TV
ρ – density
Asc – projected area
Msc – s/c mass
CD – drag coefficient
scV
wV
a
TV
ρ – density
Asc – projected area
Msc – s/c mass
CD – drag coefficient
scV
7070V - Analyses
NMS: Expected Errors in the Determination of W, n, T
• Will meet science requirements
•The error depends on the number of particles registered.•Determined for a true wind velocity magnitude W of 10 m/s:
Parameter % error*
Noise(Peak cts)
W T N(O) N(N2)
3%(1000)
13% 0.40% 0.55% 0.38%
1%(10,000)
5% 0.13% 0.18% 0.13%
*from Herrero et. al. unpublished work
DANDE
7171
NMS: Functional Description
CO
LLIM
AT
OR
ION
SO
UR
CE
DE
TE
CT
OR
AN
AL
YZ
ER
NMS302 POWER BOARD
NM
S303
DE
TE
CT
OR
BO
AR
D
-167 V
-100 V
e-
-3000V
T
NMS304 CONTROLLER BOARDNMS304 CONTROLLER BOARDNMS304 CONTROLLER BOARD
0V – 5V SCAN
5V
BU
S
I2C DATA BUS
-100 V -101 V
FARADAYCUP
IONIZER
TEMPERATURESENSOR
GROUND DEFLECTOR PLATE
HOT DEFLECTOR PLATE
BA
FF
LE (IO
N-D
EF
LEC
TO
R)
DIGITAL DATA ANALOG DATA
DE
TE
CT
OR
AN
OD
ES
15V BUS
Incoming Neutrals
72
Objective Requirements
Measure in-situ density and composition (O:N2 ratio) during at least 5 sudden geomagnetic storms and 4 periods of quiet geomagnetic conditions in an altitude of at most 350 km and covering a minimum latitude of at least 54 degrees
Calibration and validation of models. Goal: also estimate the coefficient of drag in orbit at 350 - 100 km altitude.
Measure neutral winds at an altitude of up to 250 km and below and at latitudes of at least 54 degrees during 5 sudden geomagnetic storms and 4 periods of quiet geomagnetic conditions. Provide the wind data with a spatial resolution of at least 500 km (goal: 100 km).
Measure large-scale horizontal variations with in-situ density data over the course of at least 5 geomagnetic storms and 4 periods of quiet geomagnetic conditions
Develop a low-cost system to make in-situ measurements of the neutral atmosphere and adhere to Nanosat Program Requirements. Finish the proto-qualification unit on time and on budget.
7373
Accelerometer Design
AVR µController
AVR µController
16 Bit ADC6 Inputs
16 Bit ADC6 Inputs
Temperature Sensor
Temperature Sensor
ACC1QA-2000
ACC2QA-2000
ACC3QA-2000
ACC4QA-2000
ACC5QA-2000
ACC6QA-2000
TemperatureAccelerationDigital I/O
Legend:
6 6
66
Accelerometers
4th Order Band Pass Filters
1st Order Low Pass Filters
Control
SPI Data
CommandsData
Analog Board
Digital Board
System Bus (I2C)
Cost ~$3,000Precision ng*Bandwidth 6 μHz – 10 KHz****must be able to reject the larger noise outside of 6 μHz - 1 Hz to achieve 79 ng
QA-2000 accelerometer
x 6
74
Accelerometer Testing
1/3 Hz drag signal
10-1
10-3
10-5
0.1 10Frequency [Hz]
Measured Frequency Domain
Filt
er
Res
pons
e
Measured signal to noise is one decade
better than that assumed in analysis
above
7575V - Analyses
ADC: Alignment Maneuver
• Requirement: align spin axis to within 5° of the orbit normal direction in 120 hrs
• Concept: open loop command the alignment maneuver using data collected over the past orbit
• Status: – Performance specified,
conservative compared to SNOE performance
– Coding algorithm for testing
Determine in ECI frame
Downlink HCI data since previous ground pass
Find change in angular momentum ( ) from current to desired attitude L
Predict local B-field in the upcoming orbits
With known and , use . to find necessary
Generate torque rod commands and timestamps
Upload to spacecraft
Bm
m
Col
lect
HC
I da
ta
betw
een
grou
nd
pass
es
B
76
COM: Antenna Analysis Background (2)
• Recommended Alternative: ¼ wave shortened patch antenna
– Length tunable to frequency– Feed point adjustable to 50Ω
• Simple Design: Can meet VSWR requirements without matching network
FIRST RF Corporation
Feed pinthrough substrate
PatchAntenna
Short
Connection toground plane
76V - Analyses
77
COM: Model Geometry Representation
• Antenna– ¾” wide, variable– 4.35” long, variable– ½” feed point, variable
• Hemispheres– 1/16” thick– 9” radius
• Equatorial Plate– 1” thick– 8.25” radius
• Substrate ring– 1.5” wide– 3/4” thick, variable
• Metal Bridge – Ensures conductivity between hemispheres
and equatorial plate for ground plane• Antenna Short
– Short from antenna to equatorial plate
77V - Analyses
78III - Systems Engineering 78
DANDE Attitude Components
• Magnetometer– Honeywell model HMR2300 identified– Requirements to be validated by
system magnetic environment test• Horizon Crossing Indicators
– Component specified– Specifications approved by vendor
• Torque Rods– EDU design frozen– EDU Details, Assy, PCB designed
• Nutation Damper– EDU design frozen– Details and Assy drawings released
79I - Objectives and Requirements
Power budget
1. Safe mode acquisition and recovery mode (beaconing and ADC sensors)
2. Standby mode operational low-power mode to recharge batteries3. LAB sep high-current draw during HOP activation (5 min)4. Spinup ADC mode for attaining proper spin rate5. Science active science sampling6. Alignment ADC mode for trimming spin orientation
Power deficits are handled operationally by alternating between science mode and standby mode. DANDE can gather science for 22 hours before recharging for 4 hours (storm events are on the order of 10 hours)
Safe Standby LAB Sep. Spinup Science Alignmentsystem power draw 5,301 mW 3,251 mW 33,428 mW 7,000 mW 12,374 mW 7,115 mWPV output per orbit 20.47 WH 20.47 WH 20.47 WH 20.47 WH 20.47 WH 20.47 WHBattery balance at end of orbit 6.63 WH 9.01 WH -26.07 WH 4.66 WH -1.59 WH 4.52 WHTime can sustain to 25% DOD continuous continuous 1.38 hours continuous 22.62 hours continuousTime to recharge from 25% DOD 3.99 hours 3.99 hours 3.99 hours 3.99 hours 3.99 hours 3.99 hours
80I - Objectives and Requirements 80II - State of the Program
State of the DANDE Program: Budgets
LinkUplink• + 8 dB margin• Requirement met for minimum 18dB Eb/N0• Margin with most recent antenna testing data, which shows low
receive antenna gain.Downlink• + 1 dB margin• Requirement met for minimum 18dB Eb/N0• Ongoing analysis to increase positive margin.
81II - State of the Program
State of the DANDE Program: Budgets
Mass
DANDE subsystem masses by milestone
0
10
20
30
40
50
60
PDR CDR PQR
Review
Mas
s w
ith
co
nti
ng
ency
(K
g)
SYS
Cabling
NMS
ACC
CDH
COM
EPS
ADC
THM
SEP
STR
Estimate Contingency
46.0 kg 7.5 kg
82II - State of the Program
State of the DANDE Program: Budgets
Personnel Skills
Hardware Budget
Purchases Pending
SUMMER 2008 FALL 2008
Total Team Size 10 27
Programming 5 9
Electrical 3 12
Mechanical 3 5
COM 2 4
ADC 1 3
I&T 6 23
Estimated Cost Amount Spent Total Funds Percent Remaining
$134,400* $55,400 $136,350 60%
Precision Machining
Flight Rev.
Transmitter
Flight
Hemispheres
$10,000 $8,000 $10,000
*12% contingency added
83IV - Electronics
Power System Overview
• Miniature PV strings– High (15V) voltage for direct-energy transfer– All 8 cells in a string are coincident - no intra-string cosine losses– Small strings minimize faceting and canyoning improving science – Miniature cells surface-mount soldered directly to PCB substrates,
eases student manufacturing process– PCB substrates are backed with machined Delrin to conform to
spherical surface– Two types of arrays (square, triangular) allow high coverage, low
manufacturing complexity
+
_
Battery B12V
4000mAh
+
_
Photovoltaic arrays24 watts, 15 volts
+/-15V
+5VX 3
X 1
Battery A12V
4000mAh
X 3
X 1
Solid-State Relays
Ground lines to subsystems
Single-pointChassis ground
Main Bus, +12VUnregulated output
Regulated outputsLatchingrelays
Latchingrelays
DC-DCConverters
InhibitsInhibitsBatteryCharge control
Recommended