LionSat The Pennsylvania State University University Nanosat III PDR Presentation August 14th_15th, 2003 Logan, Utah

1
LionSatThe Pennsylvania State
UniversityUniversity Nanosat III PDR
PresentationAugust 14th_15th, 2003Logan, Utah
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Program Objectives
Mission Statement The LionSat mission will
investigate the local ambient and perturbed
plasma environments surrounding a small satellite
in the Earths ionosphere. LionSat will measure
the ambient plasma environment and the ram and
wake regions using a novel hybrid plasma probe
instrument. LionSat will test a miniature RF ion
thruster system that will augment the satellite
spin, which is necessary for mapping the plasma
environment surrounding the satellite.

• Technology Demonstration
• LionSat will use the Hybrid Plasma Probe as a
  measurement instrument and demonstrate its
  efficiency. It will also allow the testing in
  situ of the miniature RF Ion Thruster as a
  satellite spin control device.

• Mission Objectives
• Primary Objectives
• P1. To map the ram and wake plasma structure
  surrounding a small satellite
• P2. To collect data on ionospheric plasma in a
  variety of geophysically interesting locations in
  low Earth orbit
• P3. To test, on orbit, a miniature RF ion
  thruster
• Secondary Objective
• S1. To test IP communications for uplink and
  downlink to a spacecraft in low Earth orbit

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Program Objectives

• Mission Requirements
• The plasma environment mapping throughout the
  entire perturbed region (i.e., ram and wake) must
  be resolved to 305 degrees about the satellite
  spin axis.
• The electron density and temperature of both the
  unperturbed and perturbed regions around the
  nanosatellite must be collected and returned to
  the ground for a minimum of three separate
  geophysically meaningful campaigns.
• The miniature RF ion thruster must demonstrate a
  minimum of one hour of continuous operation.
• The miniature RF ion thruster must demonstrate a
  measurable change in the satellite rotation rate.

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Mission Timeline
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Schedule
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System Requirements
System Requirements Method Status
The mass of LionSat Shall not exceed 30 kg. Its center of gravity shall be no more than 0.25 from the Internal Cargo Unit (ICU) centerline and must not lie more than 12 above the Separtion Interface Plan (SIP). Reference the AFRL Internal Cargo Unit Users Guide UN-0001 rev July 2003. Design Test planed
The fundamental frequency of the LionSat should be above 100Hz according to the AFRL Internal Cargo Unit Users Guide UN-0001 rev July 2003. Design Analysis Test planed
LionSat s physical envelope should satisfy the requirements described in the AFRL Internal Cargo Unit Users Guide UN-0001 rev July 2003 Design Test planed
Clearance Design planed
LionSat will be mounted to the Internal Cargo Unit via a PSC Lightband system. The description of the mechanical and electrical interfaces are given in the following documents AFRL Internal Cargo Unit Users Guide UN-0001 rev July 2003 and PSC document 2.000.510 Design Test planed
LionSat should be able to withstand the critical thermal environments encountered during Launch and while in space. Design Analysis Test planed
LionSat should be designed to withstand the vibroacoustic environment of the shuttle without failure as precised in the UN-0001, 6.3.3.5. Design Analysis Test planed
LionSat should be able to operate in the vacuum conditions encountered in space. Design Analysis Test planed
LionSat pressure environment should be able to withstand the depressurization and repressurization rates mentioned in the UN-0001, 6.3.3.6. Design Analysis Test planed
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System Requirements
System Requirements Method Status
LionSat should meet the technical requirements described to meet the safety requirements described in NSTS 1700.7B Design Analysis planed
The system should meet the margins of safety described in UN-0001, helped by the Stress Analysis as described in the Stress Analysis Guideline UN-SPEC-12311 Analysis planed
Safety Requirements Critical and Catastrophic hazards NASA-STD-5003Fracture control requirements for Payloads using the Space Shuttle Design, Analysis, Test planed
LionSat wil use inhibits to control hazardous functions as described in the NSTS 1700.7B via inhibits. Design Control/Monitor Test planed
The University shall furnish a flight unit to AFRL prior the completion by AFRL of system level tests which are described in the UN-0001. 8.1 Design Analysis Building Test planed
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Experiment 1 Hybrid Plasma Probe
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Experiment 1 Hybrid Plasma Probe

• Purpose
• To collect data on ionospheric plasma in
  perturbed and unperturbed regions within
  geophysically interesting areas of low earth
  orbit
• To demonstrate combination of several plasma
  instruments is feasible, efficient, and powerful
• Success criteria
• Collect meaningful data from at least 3 of 5
  instrument modes in ram, wake, and unperturbed
  regions and transmit to ground
• Resolve plasma environment to 30º /-5º about
  spin axis for at least 3 regions of interest
• Collect data for a period spanning at least 2
  months
• Operational scenario
• T1 LionSat deployed, HPP booms in stowed
  configuration
• T4 LionSat powers up, HPP off, booms deployed
• T7 HPP powered on daily
• when have available power, and LionSat not
  transmitting
• mode determined by uploaded schedules or
  automatic run

? Satellite Motion
Monte Carlo particle impact simulation The
orbital plane coincides with the plane of the
image. Centers of booms rotate through ram and
wake while ends remain in ambient environment.
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Experiment 1 Hybrid Plasma Probe

• Performance Characteristics
• Swept Bias Langmuir Probe operation yields
  electron and ion density, electron temperature,
  and spacecraft potential
• Fixed Bias Langmuir Probe yields relative
  electron or ion density
• Plasma Frequency Probe produces fast absolute
  electron density measurements
• Fast Temperature Probe operation yields fast,
  relative electron temperature measurement
• Boom Sensors Status
• Two aluminum booms of 42 cm length will extend
  from near center of LionSat into orbit plane
  using rigid telescoping method (model shown on
  right)
• Each boom will contain 2 sensors, one 12 cm from
  structure, one 32 cm away. Both will be
  surrounded by guard regions and consist of coated
  metal to minimize non-linear effects.
• Hardware Status
• Have initial electronics board fabricated and are
  currently populating and connecting to CDH
  processor

Initial Telescoping Boom deployment design.
Picture of HPP BOARD Here
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Experiment 2 Propulsion The RF Ion Thruster
Requirement Method Status
Pressure vessels and pressurized components as defined in NASA-STD-5003 are prohibited. Ref AFRL ICU Users Guide UN-0001 rev July 2003 Design completed
Fluid and gas containers or structural compartments that cannot be vented shall meet the definition of a sealed container as specified NASA-STD- 5003 section 3.39. Design completed
The sealed container must have a stored energy of less than 14,240 ft-lbs (19310 joules) and an internal pressure of less than 100 psia (689.5 kPa). Ref AFRL ICU Users Guide UN-0001 rev July 2003 Design Analysis Test underway
Each propellant delivery must contain a minimum of three mechanically independent flow control devices in series to prevent engine firing. These devices must prevent expulsion through the thrust chamber. Ref NSTS 1700.7B Section 202.2a Design Analysis Test underway
A minimum of one of the three flow control devices of the Xenon gas will be fail-safe, i.e., return to the closed condition in the absence of an opening signal. Ref NSTS 1700.7B Section 202.2a Design Analysis Test underway
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Experiment 2 Propulsion The RF Ion Thruster
Requirement Method Status
While the payload is closer to the Orbiter than the minimum safe distance for engine firing, there shall be at least three independent electrical inhibits that control the opening of the flow control devices. Ref NSTS 1700.7B Section 202.a (3) Design Analysis Test underway
A payload shall be two failure tolerant to prevent leakage of propellant into the Orbiter cargo bay past seals, seats, etc., if the leak has a flow path to the storage vessel. Ref NSTS 1700.7B Section 202.2d Design Test underway






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Experiment 2 Propulsion- The RF Ion Thruster

• Purpose The RF Ion Thruster will be designed to
  increase satellite plan.
• Type Total deltaV
• Functional Characteristics
• Thrust 0.6 mN
• Specific impulse 3800 s
• Total input power 15 W
• Frequency 13.56 MHz (industrial unregulated
  frequency)
• Weight 0.1 kg
• Voltage Drop across 2 grids 1kV
• Exhaust Velocity 120 m/s
• Physical characteristics
• The propellant used by the RF Ion thruster
  consists of a Xenon gas whose mass is TBD. The
  estimate peak power and temperature requirements
  are about respectively
• The propellant will be contained in a sealed
  container that meet the general requirements.
• Main components
• The RF Ion thruster is made up of a flow
  controller that monitors the Xenon gas flow rate
  coming from the reservoir to the main chamber. An
  RF Coil surrounding the chamber is connected to a
  matching network and fed by the RF source.
• Hardware status TBD

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Experiment 2 The RF Ion Thruster

• Purpose
• Test the RF Ion Thruster as a attitude control
  propulsion device
• Success criteria
• Operational scenario

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Structure and Mechanism Requirements
Requirement Method Status
Mass must be no more than 30 kg and center of gravity (CG) location must lie no further than 0.635 from ICU centerline in accordance with University NanoSat Document UN-0001, Rev-Issue Date 7/03, Section 6.1.2, Section 6.1.3, and Section 8.2. Design, Analysis Underway
Structure able to withstand maximum unidirectional test loadings of 23.8 Gs in accordance with University NanoSat Document UN-SPEC-12311, Rev-, Section 2.1. Test Underway
Following integration of ring and satellite, AFRL will verify that the mass/CG properties of the LionSat System fall within the constraints specified in the University NanoSat Document UN-0001, Rev-Issue Date 7/03. Test Underway
Structure designed to allow for ground handling and transportation as approached by Spacecraft Structures and Mechanisms (Sarafin 1995 pgs 52-54) and University NanoSat Document UN-0001, Rev-Issue Date 7/03, Section6.3.3.4. Design Underway
Structure must be free of pressurized components that do not meet the requirements defined in NASA-STD-5003 and University NanoSat Document UN-0001, Rev-Issue Date 7/03, Section 6.3.3.8. Design Complete
Design Factors of Safety (FOS) for the structure must be met and/or exceeded, as well as, Margins of Safety (MS) must be zero or greater for both yield and ultimate stress conditions as stated by University NanoSat Document UN-0001, Rev-Issue Date 7/03, Section 6.1.2, Section 6.3.3.2, and Section 8.1.1. Analysis Planning
AFRL will verify that the integrated ICU/LionSat System can exceed requirements for the following testing Sine Sweep, Sine Burst, and Random Vibration (University NanoSat Document UN-0001, Rev-Issue Date 7/03, Section 8.1.3. Test Planning
Payload stiffness must exceed 50 Hz and a fixed base natural frequency greater than 100 Hz according to University NanoSat Document UN-0001 Rev-, Section 6.3.1. Analysis, Test Underway
Hardware should be qualified for the Space Shuttle vibroacoustic environment with regards to University NanoSat Document UN-0001 Rev-7/03, Section 8.1. Design, Test Planning
Fracture control assessment should be done to remove all plausibility of catastrophic hazards to Shuttle Orbiter or Crew following NASA Document NASA-STD-5003, Rev- and University NanoSat Document UN-0001 Rev-,Section 8.3. Analysis Planning
Pressure Profile Analysis on all of NanoSat-3 hardware must be completed for pressurization and depressurization environments using values from University NanoSat Document UN-0001, Rev-7/03, Section 6.3.3.6. Analysis Planning
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Structures and Mechanisms

• The bus is octagonal in shape with a diameter of
  47 cm and a height of 46.5 cm
• Structural Components
• Side panels, end caps, mounting boxes
• The thicker of the two end caps will have a
  raised lip to act as an adaptor to the light band
  interface which will also allow for solar cells
  to be added to this face
• 6061 AL Alloy used for side panels and end caps
  which are closed isogrids
• 2 end caps, one being 0.75 thick (Interface
  adapter) and the second 0.125 thick
• Side panels are 0.45 thick with isogrid pockets
  that are 0.40 in depth
• Flanges added on end panels for attachment to end
  caps
• Surfaces to be treated with Iridite or Alodine
  and black and white paints may be used externally
  if there is room
• Trade studies conducted on internal metal
  temperatures by Thermal Subsystem
• Component Mounting boxes
• Eight 6061 AL Alloy boxes will house the
  subsystems on the end cap and other science
  packages which will be at various locations in
  the satellite (taking into account CG location)
• Lockwires will be used for fasteners for their
  reusability and back-out protection

End cap and two side panels closed isogrid
pattern is depicted
End cap showing internal mounting boxes
positioning
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Structures and Mechanisms

• Mechanisms
• Sealed Container Must meet requirement
  specified by NASA Document NASA-STD-5003 (stored
  energy lt 19,310 J and internal pressure lt 689.5
  kPa)
• Amount Xenon propellant must be specified
• Have discussed building sealed container with
  vendors using above specifications
• Booms
• Deployment There are two telescoping booms that
  will protrude
• from the bellyband. Deployment will be
  motorized.
• Attachment Housing will be designed to minimize
  area used
• and optimize attachment to side panels
• Solar Cells
• GaAs or Si cells will be mounted using a
  procedure designed by
• University of Michigan exact pattern will be
  finalized after exact cell type decided
• GPS, Antennas, and Thrusters
• Located on the bellyband of the satellite

Telescoping Boom Mock-up
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Structures and Mechanisms

• Mass Budget (see side notes)
• Using Space Mission Analysis and Design (Wertz
  and Larson 1999) and known hardware
• Preliminary Structural Analysis (during ascent)
• Cosmos Works
• Boundary conditions
• Loads and restraints
• Preliminary Results
• No Hardware has been constructed
• Mock-ups
• 1/5 scaled
• Full scaled
• Preliminary boom design

Structure 6kg Power 8kg Prop 1.8kg GNC 1.8kg S
cience Payload 4.50kg TTC 2.35kg Margin 5.55kg

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Structures and Mechanisms
Example of meshing in our structural Analysis
software, Cosmos Works mesh

• Structural Analysis using Cosmos Works
• Initial analysis underway for ascent portion of
  the mission because maximum loads occur during
  launch and burns
• Assumptions made
• Variation in pressure from 1 atm (ground) to 0
  atm
• Boundary Conditions applied
• Rigid constraint applied to thicker end cap to
  take into consideration the LBR interface with
  LionSat and ICU
• Loads applied
• Multiple load cases to show survivability of
  payload
• Expectations and Results
• Expected base natural frequency of 120 Hz
• Structure will withstand the limit and test
  loadings

Example of a stress plot in Cosmos Works
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Launch Vehicle Interface
Light band adapter flange

• Mechanical interface
• External flange (0.2)
• Allows for solar cell clearance on interface side
  of satellite
• Electrical interface
• 2 micro switches with inhibitors

LBR
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Electrical and Power Requirements
Requirement Method Status
Battery recharge currents Design Underway
Depth of discharge (DoD) Design Underway
Battery cell temperature monitor -prevent temperature related failures AFRL University Nanosat Internal Cargo Unit (ICU) Users Guide - 6.6, rev. July 2003 Design and Test Underway
Battery cell voltage monitor -prevent overcharge AFRL University Nanosat Internal Cargo Unit (ICU) Users Guide, rev. July 2003 Design and Test Underway
Safety -Batteries initially discharged until separation from the ICU AFRL University Nanosat Internal Cargo Unit (ICU) Users Guide - 6.5.1, rev. July 2003 -Interlock disconnect of solar array and battery pack until separation from the CAPE Design and Test Underway
Grounding/Bonding -Provide a continuous, electrically conductive path between each major structural component and the ICU AFRL University Nanosat Internal Cargo Unit (ICU) Users Guide - 6.5.3, rev. July 2003 Design and Test Underway
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Electrical Systems and Power

• Operational Power Scheme

• Operational modes Science Collection, Data
  Transmission, Attitude
  Correction, Microthruster Test, Idle

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Electrical Systems and Power

• Batteries
• NiCad Sanyo N4000DRL, one 10-cell pack, 1.2 V /
  cell, 12 V / battery
• Charge control preflight EGSE control on-orbit
  control by the PCU
• Flight heritage Recommended by Nanosat 3
• Safety Thermistor temp. monitor, EGSE volt
  monitor, zener volt clamp if PCU fail
• Solar Cells
• GaAs/Ge, 18.3, body mounted 8-sides 2 end
  caps, 3 strings of 20 cells / side
• Organization responsible for lay-up/mounting
  PSU (ref. UM-procedure)
• Converters Interpoint MHE1205S Datel
  UNR-3.3/3-D12 PICO MRF5S, MRF15D, 12AV1000
• Sinks Zener overvoltage clamp TBD
• Interfaces
• Connector types 1 NLS0H14-35S for EGSE, 1
  NLS0H10-35S for SC
• Fusing/ Circuit protection 1 FM08A125V7A (50
  derated)
• Wiring- rating, material Need to choose
• Number and type of grounds One power-to-SC frame
  ground connection implemented by relay contact
  energized by recharged battery after deployment

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Electrical Systems and Power
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Communication Requirements
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Communication

• Requirements
• Downlink frequency 2.365 GHz
• Downlink data rate 200 K symbol /sec
• Uplink data rate 9.6 kbit/sec
• Radio
• Xmittr L-3 Comm. T-155
• 0.5 W output, FM modified for PM, 85
  grams, 1.3 in3
• 3 W Power Amp, NE6510179A zz
  grams, 0.75 in3
• Receiver Honeywell HRF-ROC09325 yy
  grams, 3.3 in3
• Front end, WJ mixer, RF amp, etc. Zz
  grams, TBD
• Uplink/Downlink antennas
• Slot antennas under development (radial
  wave guides)
• Mass, dimensions, TBD
• Located on bellyband, four sides for uplink, four
  sides for downlink
• Omnidirectional pattern with nulls
• GPS antennas Microstrip patches (TOKO
  DAX1575MS63T) on eight sides

T-155 transmitter
Honeywell receiver on a chip

• Hardware status T-155 in stock
  Honeywell evaluation board used on another
  project

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Communication

• Link requirements
• Six sequential overpasses/day
• Reacquisition needed for each pass
• Maximum path length 1470 1840 km? maximum
  tdelay 4.9 6.1 ms
• Data generated on satellite, asymmetric link
• LionSat initializes handshake based on GPS and
  ephemeris data
• If no response, LionSat broadcasts in the blind
• Protocol options
• Lionsat will use IP communications for return of
  science data
• After mission criteria met, can be used as
  testbed for testing/verifying relative
  performance of various protocols
• Status of frequency allocation current satellite
  transmitter unit (T-155) is designed for 2.365GHz
  (military band), waiting for Nanosat 3 guidance

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Science Data
Communication Requirements

• Roll rate of 10 rpm? 14,400 rolls/day
• 12 samples per roll? 4 sensor heads? 691,200
  samples/day

Functional Objective Swept Plas. FP Swept Bias LP Tracking Plas. FP Fast Temp. P Fixed Bias LP Portion of day MB/day
1 5 10 40 10 40 15 11.0
2 100 0 0 0 0 1.5 10.6
3 0 20 40 10 40 15 11.0
4 0 2.5 0 0 97.5 100 10.2
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Communication Requirements

• Data Per Sensor Subunit

MB/day
Science Functional Objective dependent 11.0
Magnetometer (for science and attitude) 6 B/sample 172,800 samples/day 1.04
GPS (for orbit determination and time) 20 B/sample 0.1 samples/second 0.17
Housekeeping 6 temp., 40 voltage, 10 current, 2 tank pressure, and 3 horizon sensors 2 B/sample/ch 61 chs 0.1 samples/second 1.05

• Total 13.3 MB/day to download

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Transmission Rate
Communication Requirements

• Minimum elevation angle and overhead dependent
• Baseline design 200 ksymbols/sec downlink
  9.6kb/sec uplink

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Orbital Parameters
Communication Requirements

• Launched into LEO orbit from Shuttle
• 51 inclination
• 400 km altitude
• 6 mo ? 1 yr lifetime

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LionSat Communications Timeline View
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LionSat Communications Layer View
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ADCS Requirements
Status
Method
Requirement
Done
Simulation
Need to spin to collect the experimental data
through extended probes
Done
Simulation
Maintain spin axis always normal to the orbit
plane
Capability to control spin rate with magnetic
torque rod
Done
Simulation
Underway
Attitude information using only a
magnetometer Then, attitude update with a Sun
sensor
Analysis Simulation
Underway
Test Design
Spinner has to be oblate spinner,
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ADCS

• Performance requirements
• Attitude Determination
• Attitude estimation 15 deg
• No gyro sensor
• Attitude determination with a 3-D magnetometer
  only (angular velocities and integration to find
  Euler angles or Quaternion)
• Attitude update with a Sun sensor (TRIAD/QUEST)
• Attitude Control
• Maintain Orientation within 5 degrees
• No Nutation
• Successful correction of regression of RAAN
• Nutation damping with passive damper

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ADCS

• Physical Characteristics
• Type of ADCS
• Spin stabilization using geomagnetic field and a
  magnetometer only attitude determination
• Passive nutation damping
• Magnetic torque rods
• Basically coils around highly permeable core
• Mass depends on size of the rods
• For 1015 magnetic moment rods
  approx. 0.5 kg
• Mechanisms
• Controller decides polarity of the rod to change
• Two magnetic torque rods will be used to maintain
  desired attitude

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ADCS

• Performance Characteristics
• Magnetic Torque Rods
• Power consumption 1 Watt per axis
• Operational modes
• Orientation Control
• Spin Rate Control
• Stand-by (resting mode)
• Torque Generated at 400km altitude
• Start-up/activation
• At the separation from the launch vehicle approx.
  spin rate of 5 rpm will be provided
• Then, orientation change to desired spin axis
  orientation
• Finally, spin rate control, if desired
• In normal operation, most correction will be
  orientation control

Goodrich TR series Magnetic Torque Rods
Example Microcosm MT series Magnetic Torque Rods
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ADCS

• Performance Characteristics(cont.)

Actualized pictureErikas hand
Goodrich Sun sensor Model 13-515
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ADCS

• Performance Characteristics (cont.)
• Stowed configuration
• Hardware status
• Collection data on available hardware
• To buy space-rated hardware from companies
• Or some possibilities to produce our own magnetic
  torque rods with Air Force assistance

Magnetometer
Mag. Torquer
Passive Damper
ADCS
Sun sensor
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Thermal Requirements
Status
Method
Requirement
Underway
Design, Analysis, Test
Provide adequate thermal protection for the
spacecraft during all mission phases along with a
complete thermal analysis. AFRL Users Guide
UN-0001
Analysis Completed Test planed
Analysis, Test
Complete a preliminary analysis of the thermal
subsystem, determining the worst case hot/ cold
temperatures for the spacecraft during the
mission.
Analysis Completed Test planed
Analysis, Test
Further the thermal design with a one-node
analysis of the spacecraft (STK).
Underway
Analysis, Test
According to document UN-0001, thermal models
must be supplied to AFRL, at a minimum including
a simplified model which includes nodes for each
of the temperature-critical components (SINDA).
Analysis Completed Test planed
Design, Analysis, Testing
Provide temperature limits as defined by UN-0001
for each node in the reduced thermal math model
and for each subsystem.
Planed
Analysis, Test
Determine all heat sources and their respective
profiles during the various mission
phases. UN-0001
Planed
Analysis, Test
The heat transfer must be quantified for the
satellite to the external environment (ICU or
space). UN-0001
Planed
Analysis, Test
Determine temperature-critical components. UN-0001
Underway
Design, Analysis
Determine all payload external surface properties
including size, material/ process, absorptivity,
and emissivity values. UN-0001
41
Thermal
Spacecraft Temperature Ranges
Component Operating Temp (C) Storage Temp (C)
RF Probes -40 to 85 -65 to 150
GPS - 30 to 75 -55 to 90
Onboard Computer 0 to 70 -40 to 85
Transmitter - 20 to 70 -55 to 100
Command Receiver - 40 to 85 -40 to 150
Solar Cells - 105 to 110 TBD
Batteries (NiCad) 5 to 40 - 30 to 50
Power control Unit - 40 to 85 -40 to 85
Magnetometer 0 to 70 TBD
Magnetic Torquer TBD TBD
Sun Sensor - 40 to 85 -40 to 85
RF Ion Thruster TBD TBD
Boom Motor TBD TBD
GPS Antenna - 40 to 105 -40 to 105

• Thermal classification Passive
• Control Types Paints, Coatings
• Thermal verification
• Temperature Sensors
• Thermal Analysis
• Spherical Analysis
• Temp. Range -76 to 54 ºC
• One Node Analysis (STK)
• Temp. Range
• Multi Node Analysis (SINDA)
• Temp. Range TBD

42
Thermal

• The objective of thermal control is to maintain
  all components of the satellite within their
  allowed temperature limits during all the mission
  phases utilizing a passive control system.
• The Software packages STK, TRASYS, SINDA

• LV Phase Thermal Margins
• Pre-launch Shuttle 18 to 21 ºC
• Minotaur 18 to 28 ºC
• Launch Shuttle -20 to 60 ºC
• Minotaur 22 to 93 ºC

• Analysis Cases
• Spherical Analysis
• One Node Analysis
• Multi Node Analysis (SINDA)
• Analysis Assumptions
• (SA) - Steady-state, Uniform energy
  dissipation, no electrical generation on the
  surface
• - Satellite Altitude 400 km (400,000 m)
• - Internal Dissipation values are estimated to
  be between 10 - 30 W
• - For upper limits, high side values of Albedo
  (,- 30), Solar constant Gs
• (1358 ,- 5 W/m²), and Earth IR
  Emission ( 237 ,- 21 W/m²) were used for lower
  limits,
• the lower values.
• (1N) -
• -
• (MN) - Assume properties of absorptivity
  and emissivity as 0.805 and 0.825, respectively.
• - Assume a cylindrical spacecraft of radius 47 cm

43
CDH Requirements
Requirement Method Status












44
CDH Requirements

• Structures
• Deploy probes
• Verify probe position
• Monitor strain gauges during flight testing
• Power Systems
• Monitor solar panel temperature
• Monitor solar panel health
• Monitor battery temperature/health
• Control subsystem operation (on/off)
• Guidance, Navigation and Control
• Attitude management
• Horizon sensors
• Sun sensors
• Magnetometer
• Magnetic torque units
• GPS

• Data/Communications
• Run OS and subprograms
• Data processing
• Data storage
• Data forwarding
• Transmitter/power amp. operation

45
CDH

• Physical Characteristics
• Main components
• CPU - Intel SA1110 _at_ 206MHz
• SA1110 Memory
• Up to 768 MBytes static memory such as SRAM,
  FLASH, SMROM
• Up to 512 MBytes dynamic memory
• SA1110 subsystems interfaces
• 28 GPIO lines
• Multiple serial systems (SPI, UART, USB)
• Ground interface through TCP/IP
• Performance Characteristics
• Power consumption
• lt400mW _at_ 206MHz (excluding external memory)
• Software
• Running RTLinux to handle time critical and
  background applications

• Hardware Status
• Purchased two SA1110 evaluation systems from SSV
  for software development
• Documents and example code in hand for running
  Linux on SA1110

46
System Block Diagram
CDH
47
Science Data
Communication Requirements

• Roll rate of 10 rpm? 14,400 rolls/day
• 12 samples per roll? 4 sensor heads? 691,200
  samples/day

Functional Objective Swept Plas. FP Swept Bias LP Tracking Plas. FP Fast Temp. P Fixed Bias LP Portion of day MB/day
1 5 10 40 10 40 15 11.0
2 100 0 0 0 0 1.5 10.6
3 0 20 40 10 40 15 11.0
4 0 2.5 0 0 97.5 100 10.2
48
Software Requirements
Status
Method
Requirement
Complete
Design
Must support IP communication protocols (TCP,
UDP, etc.) and subsystems (ECC) if needed
Planned
Design
Must support real-time system clock for tagging
of science data
Planned
Design
Needs to have threading or other prioritization
functionality
Complete
Design
Must have essential operating files secured to
prevent alterations by third parties
49
Software

• Language Support supports C/C code (with
  compiler and assembler in-hand through GNU
  tools), also supports Arm assembly code
• Operating system currently running with embedded
  Linux kernel, but will install Real Time Linux
  kernel to obtain threading capability
• Architecture hierarchical through threading
  (e.g. transmitter will have higher priority than
  hybrid plasma probe). Also, GPS or other method
  will keep system clock accurate, so scheduled
  tasks will be used.
• Software Data Management and Test Plan
• system setup data only altered through secured
  means, science and monitor data will be stored in
  memory until successfully downloaded to ground
  station
• plan to have basic software components
  written/installed within 8 months to allow
  inter-component testing of hierarchy and
  scheduling before presenting CDR.

50
Ground Station

• Operations
• Operated by undegraduate students
• Located on Penn State Campus
  Latitude xx, Longitude yy
• Receiving and commanding
• Tracking verification (NORAD and
  onboard GPS)
• Hardware
• 3.6 m surplus dish
• Downlink Microdyne frontend
• Uplink TBD
• Tuscon Amateur Packet Radio Trakbox
  microcontroller
• Discussions with Embry Riddle for shared tracking
  for / by Eagle Eye
• Software
• STK Toolkit from Analytical Graphics, Inc.
  Penn State site license includes HPOP, High
  Precision Orbital Propagator
• Trakbox firmware
• MATLAB-based software radio development

51
Program/Subsystem Risk Assessment
CDH Power Structures Launch Vehicle GNC Thermal Communications Propulsion Scientific Inst. Overall Program Assessment
Performance
Schedule
Cost
Safety
Testing
Manpower
Facilities
Overall Subsystem Assessment
52
BACK-UP SLIDES

• Required Slides
• Organizational chart
• K-12 Involvement
• Detailed schedule
• Exploded view of satellite
• Detailed requirements for all subsystems
• Solar cell mounting
• EPS Inhibit schematic
• EPS Battery box design
• COMM Link Budget
• IT Testing overview
• IT flow

Optional Slides The remainder of the back-up
slides are examples types of slides that might
also be included. There may be additional slides
not shown that are beneficial to include.
53
Organizational Chart
Dr. Sven Bilen, PI sbilen_at_psu.edu 814-965-2859

• Yellow spacecraft subsystem
• Purple science payloads
• Green program subsystem

Valerie Mistoco (G,EE), PM qecia_at_psu.edu 814-865-0
188
Structure
CM Safety
Propulsion
Outreach/Public Relations
Christopher Barella (UG, AE) Rachel Larson (G, AE)
Marc Hoffman (UG, AE)
NST (G, EE)
Erika Mendoza (UG,AE) Rachel Larson
GSE
Dr. Sven Bilen Dr Micci
(G, AE)
NST (UG, ME)
EPS
(UG, EE)
Daniel Dobrin
Plasma Probes
I T
Business / Financial (UG, BA)
Robert Siegel (G, EE)
Erika Mendoza (UG, AE)
C DH/ COMM
Brendan Surrusco (G, EE)
Dr. Sven Bilen
Dr Robert Melton Dr Deborah Levin

• Red advisory roles
• UG undergraduate
• G graduate student
• NST no summer team

AODC
Phil Hur (G, AE)
Pr Charles Croskey Dr David Spencer Dr Robert
Melton
54
K-12 Outreach

• Design and Implementation of a K-12 packet
• This packet includes summaries of the LionSat
  project, a slide presentation, and experiment
  packets aimed to several audiences. These
  packets are meant to help pull more diverse
  students towards the sciences and help them feel
  comfortable in a TEAM motivated environment.
• LionSat Group Involvement
• Space Day (400 children in attendance) The
  LionSat team led children to build small
  "nano-satellites" out of cans of soda and art
  supplies. Here, children learned about the
  various uses of a satellite and its components.
  Older children were able to learn about the
  ionosphere and what the "ram wake" of a
  satellite was.
• Engineering Open House High school students
  learned about their own particular strengths
  interests and how they can be used in building a
  spacecraft. At this event, LionSat won 1st place
  for Best Technical Display
• MTM Engineering Camp for Girls Small
  introduction to LionSat backed by three tours to
  our local Sailplane Lab, Water Channel, and to a
  facility housing the SWIFT Project as well as,
  implementation of a Thermal Protection System
  Experiment that was put together by NASA
• MEP Student Panel Introduction to LionSat and
  question answer
• Girl Scout Saturdays (Grades K-12) Introduction
  to LionSat and student involved activity
• High School 6 week and 3 week courses Interact
  with students to promote industry awareness, team
  activities, and presentations
• Home field High School Visits Students
  involved in LionSat are encouraged to visit their
  former high school in order to spread knowledge
  about LionSat and other space-related projects
  and opportunities.

55
Other Outreach Activities PSU students

• April 28-30, 2003
• Review Process of Semester Project work
  Electrical Engineering.
• The LionSat team was invited to participate to
  the review as a real project manager, to give
  their appreciation and talk to students about the
  LionSat experience.
• April 30, 2003
• Windows to Space. Talk given by Michael Wyland
• May 2, 2003
• Senior Design Project Exhibit
• The LionSat Probe team (Michael Wyland) won
  the 1st price for the best exhibit, poster
• LionSAt Involvement in the Classroom
• 3 LionSat projects via EE4033 project teams in
  Aero 402 3 teams doing small sat technology
  stuff in 4921 team in EE402 doing antenna
  design1 indep. Study on antenna design

56
Detailed Schedule
57
Spacecraft Overview Exploded View
end cap
Magnetometer
side panel
Pressure Vessel
Propulsion
Torquers
Motorized Telescoping Booms
Closed Isogrid
Damper
Probe Holes
ADCS
Receiving Antennas
Power
Sun Sensor
Miniature Ion Thruster
Transmitting Antennas
Solar Cells
GPS
Communications
CDH
58
Solar Cell Mounting

• Solar Cell Mounting Technique
• Nusil 2568 will be used to attach the
  cells to the panels
  and then a cover
  glass will be used to protect
  the cells
• Solar Cell Arrangement
• Cells will either be Si (20mm by 40 mm)
  or GaAs (22 mm by
  40 mm)
• end caps
• Top will hold approximately 60-72 cells
  depending on type of solar cell used GaAs (60)
  and Si (72)
• Bottom will have an lip extended to allow space
• for solar cells while still taking into
  consideration
• the LBR (60-72 cells as well)
• side panels
• Each panel will hold 60-72 cells
• Belly band ensures space for motorized booms
  and TM slot antennas

Figure depicts solar cell mounting arrangement
on LionSat
59
Structures and Mechanisms
Side (hidden lines removed)
Trimetric (solid rendering)
Trimetric (showing hidden lines)
Top (hidden lines removed)
60
Structures MechanismsComposite Panel
Specifications
Joint Attachments

• Locally manufactured Aluminum 6061 Alloy closed
  isogrid
• 0.45 thick plate with 0.40 milled sections and
  0.050 skin
• Solar cells mounted using technique designed by
  University of Michigan (Dave Morris)
• Side panels can accommodate approximately 80
  cells
• Joint connection points w/ ¼ holes
• Belly band
• GPS and TM slot antenna mounting
• Boom extension
• Edges
• side panels have chamfered corners on
  octagonal directional change
• fillets used where possible

24 - ¼ Attachment Locations
Solar Cells
Antenna, GPS, Boom/Probe, And Thruster Attachment
Points
61
Subsystem IonThruster Detailed Requirements
62
Subsystem Plasma Probe Detailed Requirements
63
Subsystem Power Detailed Requirements
64
Electrical Systems and Power
65
Electrical Systems and Power Battery Box Design

• Battery
• 10 Sanyo NiCd Type N-4000DRL cells
• 12V, 4 A-hr Battery
• Internal fusing
• Battery box
• 6061-T651 Al, Internal coatings Anodized
  Uralane 5750. Exterior Iridite
• Stainless steel egg crate TBD thermal epoxy
  provide structural support and thermal path for
  cells
• TBD absorbent material installed in void spaces
  to minimize free volume.
• Two filtered vents
• Two thermistors for temperature sensing
• Battery Testing
• Flight units provided by Nanosat 3
• System level thermal vibration testing followed
  by battery servicing
• Temperature and voltage monitoring during thermal
  testing

Representative layout
66
COMM Link budget
67
Integration and Testing
Structural Tests Structural Tests Structural Tests Structural Tests
Test Component Spacecraft Margins
Strength Sine Burst X Test conducted at 1.2 times the limit loads
Stiffness Sine Sweep X Overall payload stiffness gt 50 Hz fixed base natural frequency gt 100 Hz
Thermal Tests Thermal Tests Thermal Tests Thermal Tests
Thermal Vacuum X X Predicted thermal model limits with 5 degree Celsius margin minimum addition
EMC Tests EMC Tests EMC Tests EMC Tests
Shuttle Ku-Band tolerance X Design verified by scaled mockup in RF anechoic chamber
Functional Tests Functional Tests Functional Tests Functional Tests
Electrical RF probe, sensors, RF thruster Mechanical Boom deployment X X X Performance to spec verified Performed at 1 G
68
Integration and Testing
Integration and Test Flow
EPS Assembly
EPS functional Test
RF probe Assembly
Thermal tests
Ion Thruster Assembly
COMM Assembly
Structure Assembly
Assembly and Tests location
PennState
GSFC (under negotiations)
AFRL
69
(No Transcript)
70
Preliminary Structural Analysis

• Assumptions
• Modeled iso-grid as solid aluminum plates with
  smeared properties per Isogrid Design Handbook
  (NASA CR-124075)
• Modeled lightband as a set of thin shell elements
  with mass and size constraints per Lightband ICD,
  Rev C.
• Analysis performed on stack in cantilever
  position under 29.8gs
• FEM
• Models created using I-DEAS
• 3CS Stack modeled as three hexagonal volumes and
  a ring(LB).
• S/C bulkheads modeled as triangular thin shell
  elements
• S/C side panels Lightband modeled as
  quadrilateral thin shell elements
• FEA
• Static and dynamic Analysis conducted using I-DEAS

71
Ground Support and Equipment

• Power servicing requirements/activities
• Battery charge/discharge
• Inhibit verification
• Power on/off overrides
• Mechanical servicing requirements/activities
• Hoist rings and transportation box
• Propellant refueling/verification
• Gas reservoir pressure gage read out through GSC
  (No verification needed mechanically)
• Thermal servicing requirements/activities
• Monitoring temperature ranges (everything else
  autonomous)
• Functional activities
• Timeline updates
• Operational mode commanding
• RF probe stimulus
• Functional verification upon delivery or after
  testing at AFRL
• Such as
• Battery Conditions
• Sealed Container Testing
• RF Probe Activation (electronics)

72
Ground Support and Equipment
Transportation Box from vendor

• Electrical
• Control suitcase
• Umbi through CAPE adapter
• Umbi for direct connection to LionSat
• Laptop with Ethernet connection
• Safety inhibit verification
• Power on/off overrides
• Battery charging
• Volt / current / temperature monitors
• Mechanical
• 4 holes drilled (17/64 dia.) in satellite
  for hoist rings (used for lifting
  and moving after initial transportation)
• Transportation box Prelim. Design includes
• Simulated LBR to secure satellite May be
  lowered and connected directly to box
• Airtight to protect hardware
• Will look at vendors for box that can be combined
  rigidly with simulated LBR for transportation

Simulated LBR for rigid transportation
73
Detailed Risk Assessment / Mitigation
Risk element Description Proposed Mitigation
Structural Failure Critical Material Fractures Stress Corrosion Structural Stiffness Sealed Compartments Rupture Ultimate factors of safety must be equal/ greater than 1.4 Rate materials for resistance to stress corrosion according to MSFC-HDBK-527/ JSC 09604 and MSFC-STD-3029 Abide to a minimum natural frequency of 100 Hz, as described in UN-0001 Must withstand the decompression and repressurization environments associated with ascent/ descent stages.
CDH Failure Payload commanding Radiation contamination Electrical Faulting of circuits A function whose inadvertent operation could result in a critical hazard (2) or a catastrophic hazard (3) must be controlled by the said number of independent inhibits. Radiation Hardening Design Electrical Power Distribution circuitry to include circuit protection
Power Failure Overcharging systems Battery Hazards Distruption of electromagnetic field (lightning) Design Electrical Power Distribution circuitry to include circuit protection Mitigations for battery hazards are outlined in detail in JSC 20793 Manned Space Vehicle Battery Safety Handbook Harden the circuit against the environment, or by adding relays for control
Thermal Control Incapacity of the nanosatellite to maintain and control temperatures Monitor all component temperatures while in-orbit Provide passive thermal control on sensitive components with down linked temperatures
74
Detailed Risk Assessment / Mitigation (cont.)
Risk element Description Proposed Mitigation
Attitude/ Control Failure Magnetic Torquer failure Premature reentry of spacecraft GPS Failure The two Magnetic Torquers being used are non-coupled with one controlling the orientation of the spacecraft and the other controlling the spin rate. Failure of the spin rate torquer is backed by the RF thruster. A second torquer in the orientation axis could act as a backup system, however costs are a concern. Complete as many scientific mission objectives during the earliest stages of the mission as possible. Assistance from other ground tracking stations (NORAD), followed by an attempt to uplink with LionSat
Propulsion chemical Leakage Xenon propellant leakage All materials/containers exposed to hazardous chemical materials must conform to the guidelines outlined in document NASA-STD-6001
Comm. Failure Loss of communication/ control medium and charge pump overheating, burnout Backup communication systems The pump has a manual reset from ground over temperature function
Launch Vehicle Failure to deploy from ICU Ground tests of the system prior to launch
75
Detailed Risk Assessment / Mitigation (cont.)
FAULT TREE ANALYSIS (FTA)
Objective Fault Trees are used to identify and
prevent failures prior to their occurrence, but
are more frequently used to analyze accidents or
investigative tools to pinpoint failures. When an
accident or failure occurs, the root cause of the
negative event can be identified.
76
Management Documentation People

• Database to store our different documents.
• Development of different sets of documents
• Link between different teams changes reports,
  one main storage location for hard copies.
• Link between teams report of technical progress
  and notes