d rag and a tmospheric n eutral d ensity e xplorer (dande) colorado space grant consortium and cu...

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


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


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


Evolution of DANDE


Manufacturing

Advised by Tim Flaherty and Mathew Rhodes


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


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


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


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


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


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


Completed Formal Testing

TESTING


Integration & Testing Schedule

Today Competition Review


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


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


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


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%)


• 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


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


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


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


Verification

1.SYS76 The spacecraft shall spend no more than 300 days above the minimum perigee requirement of 350 km altitude.


• 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


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


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


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


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


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


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

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