LionSat Subsystem Detailed Requirements Listing

#8 LionSat should be able to withstand the depressurization and repressurization rates mentioned in the UN-0001, § 6.3.3.6 and a venting analysis should demonstrate a factor of safety of

LionSat Subsystem Detailed Requirements Listing

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Department of Electrical Engineering

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Table of Contents

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#3 The miniature RF ion thruster must demonstrate a minimum of one hour of continuous operation.

#4 The miniature RF ion thruster must demonstrate a measurable change in the satellite rotation rate

#5 LionSat should be able to withstand the critical thermal environments

encountered during launch and while in space

? Max/Min temp?

#6 LionSat should be designed to withstand the vibroacoustic environment of the shuttle without failure as described in the UN-0001, §6.3.3.5

#7 LionSat should be able to operate in the vacuum conditions encountered in space

? Vacuum capable?

#8 LionSat should be able to withstand the depressurization and repressurization rates mentioned in the UN-0001, § 6.3.3.6 and a venting analysis should

demonstrate a factor of safety of 2.0

#9 LionSat should meet the technical and safety requirements described in NSTS 1700.7B (i.e., failure tolerances and catastrophic hazards)

#10 The system should meet the margins of safety described in UN-0001 To meet this requirement, a stress analysis as described in the Stress Analysis

Guideline UN-SPEC-12311 must be performed

#11 LionSat will use inhibits to control hazardous functions as described in the NSTS 1700.7B

#3 Further the thermal design with a one-node analysis of the spacecraft (STK)

#4 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)

#5Provide temperature limits as defined by UN-0001 for each node in the

reduced thermal math model and for each subsystem

#6Determine all heat sources and their respective profiles during the various mission phases

UN-0001

#6The heat transfer must be quantified for the satellite to the external environment(ICU or space) UN-0001

#7 Determine temperature-critical components UN-0001

#8Determine all payload external surface properties including size, material/ process, absorptivity, and emissivity values UN-000

#9Component Operating Temperatures must stay within the limits of the worst case hot and cold temperatures found in the one node analysis

# 10 Component Operating Temperatures must conform to temperature limits set

by thermal computer analyses (SINDA)

• Thermal classification: Passive

– Control Types: Paints, Coatings – Thermal verification

– Temperature Sensors – Thermal Analysis

• • First-order, one-node, spherical-payload analysis

Temp Range:

Tmax = 38 oC steady state full sun Tavg = 10.5 oC

– Multi Node Analysis (SINDA)

Spacecraft Temperature Ranges

Component Operating

Temp (°C)

Storage Temp (°C)

RF Probes -40 to 85 -65 to

150GPS - 30 to 75 -55 to

90Onboard

Computer

0 to 70 -40 to

85Transmitter - 20 to 70 -55 to

100Command

Receiver

- 40 to 85 -40 to

150

Batteries (NiCad)

0 to 40 - 30 to 50

Power control Unit - 40 to 85 -40 to 85Magnetometer 0 to 70 TBDMagnetic

Mechanical Sub System: Structures

Technology Lead(s): Peter Cipollo

Structure and Mechanism Requirements

#1Mass must be no more than 30 kg and center of gravity (CG) location must lie

no further than 0.25” from ICU centerline in accordance with University NanoSat Document UN-0001, Rev-Issue Date 7/03, §6.1.2, §6.1.3, and §8.2

Get this data from pete

#2 Structure able to withstand maximum unidirectional test loadings of 23.8 G’s

in accordance with University NanoSat Document UN-SPEC-12311, Rev-, §2.1data FROM PETE

#3 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

#4 Structure designed to allow for ground handling and transportation as

approached by Spacecraft Structures and Mechanisms (Sarafin, 1995, pp 52-54)

and University NanoSat Document UN-0001, Rev-Issue Date 7/03, §6.3.3.4

#5 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, §6.3.3.8

#6 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, §6.1.2, §6.3.3.2, and §8.1.1

#7 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, §8.1.3

#8 Payload stiffness must exceed 50 Hz and a fixed base natural frequency greaterthan 100 Hz according to University NanoSat Document UN-0001 Rev-, §6.3.1

#9 Hardware should be qualified for the Space Shuttle vibroacoustic environment with regards to University NanoSat Document UN-0001 Rev-7/03, §8.1

#10 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-, §8.3

#11 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, §6.3.3.6

#12LionSat’s physical envelope should satisfy the requirements described in the AFRL Internal Cargo Unit User’s Guide UN-0001 rev July 2003

Subsystem

Estimated mass (kg)

Hybrid Probe 2Communications 0.37

Hybrid Probe ?Communications ?

>Spacecraft structure octagonal in shape with diameter of 18.50 in (47 cm) and height of 18.31 in (46.5 cm)

Top End Cap

 mounting point/fasteners for xenon tank, valves, and pipe work

 at least 6 holes for solar cell wiring

 magnetometer mount/fastener—hole for wiring

 cable braces

 sun sensor mount/fastener—hole for wiring

 24 (3 per side) #6 tapped holes on flange for attaching side panels

Bottom End Cap

 at least six holes for solar cell wiring

 light band connection—24 ¼ “ holes for mounting satellite

 sun sensor mount/fastener—hole for wiring

 light band insulator

 cable braces

 mounting points/fasteners for mounting boxes

 2 holes for light band power inhibitor switch connection ports

 24 (3 per side) #6 tapped holes on flange for attaching side panels

Side Panels

 mounting points/fasteners for whip antennas

 2-holes for boom deployment (1 hole in two opposite side panels)

 sun sensor mount/fastener—hole for wiring

 at least ten holes per side panel for solar cell wiring

 cable braces

 mounting points/ holes/ fasteners for thruster nozzles (in two opposing side panels)

 mounting points/fasteners for thruster piping/ wiring

 torque rod mounting points/fasteners

 14 holes for connection to neighboring side panels and end caps

o 3 #6 flat head holes for top and bottom of panel

o 4 #4 flat head holes per side

Boom System

Construct booms and deployment system.

 Boom design/construction

 Boom deployment system

1 Flight Computer SA1110 &Memory

-need connection points for

o Power/GND

o SSP Data Bus

o I/O Lines (chip selects)

o RS-232 Port

o Light Band Interface (power/GSE)

o House Keeping Sensors

o Downlink Data

o Uplink Data

2 Analog/Digital Sensor Board

-need connection points for

o SSP Bus, chip select

o Torquer control (SSP Bus, chip select)

o Wiring to x-axis magnetic torquer

o Wiring to y-axis magnetic torquer

o Wiring to z-axis magnetic torquer

5 Main Power Controller

-need connection points for

o Ten solar array inputs & solar cell return

o Power/GND from voltage/current sensor

o Power control (SSP Bus, chip select)

o DC/DC converter (may need up to seven holes

TBD absorbent material installed in void spaces to minimize free volume

Two filtered vents

Two thermistors for temperature sensing

o Band pass filter

o 3 for communication array

-need connection points for

o Power/GND

o Uplink Data

o Band pass filter

11 Band Pass Filter

-need connection points for

13 Miniature RF Ion Thruster

-need connection points for

o Wires to thrusters

o Pipe work to thrusters

o RF splitter

o Connection to xenon tank

o Gas flow control valve Flow Control Unit (FCU)

o Pressure sensor (analog inputs)

o Power Supply and Control Unit (PSCU): Beam supply, Accel electrode supply,

RF generator supply, Neutralizer supply, Beam current controller for thrust control, current limiter to the main bus, telemetry and remote controlinterface

Propulsion – Miniature RF Ion Thruster

#1 Pressure vessels and pressurized components as defined in NASA-STD-5003 are prohibited Ref: AFRL ICU User’s Guide UN-0001 rev July 2003

#2 Fluid and gas containers or structural compartments that cannot be vented shallmeet the definition of a sealed container as specified NASA-STD- 5003 §3.39.

#3 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 User’s Guide UN-0001 rev July 2003

#4 Each propellant delivery must contain a minimum of three mechanically independent flow control devices in series to prevent engine firing These devicesmust prevent expulsion through the thrust chamber Ref: NSTS 1700.7B §202.2a

#5 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 §202.2a

#6 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 §202.a (3)Minimum safe distance is Approxmintly 7.39x10^-7”

#7 The sealed container must be pressurized to 1.5 atmospheres or less and be compliant to the requirements 4.2.2.4.2.3a (1) and (2) from NASA-STD-5003 If not, an analysis shall show the safety factor is 2.5 or greater or that the container shall be proof-tested to a minimum of 1.5 times the Maximum Design Pressure

#8 A payload shall be two failure tolerant to prevent leakage of propellant into theOrbiter cargo bay past seals, seats, etc., if the leak has a flow path to the storage vessel Ref: NSTS 1700.7B §202.2d

• Functional characteristics

– Thrust: 0.6 mN (calculated)– Specific impulse: 3800 s– Exhaust velocity: ~38 km/s

• Physical characteristics

– Total input power: 15 W– Excitation frequency: 13.56 MHz (industrial unregulated frequency)– Mass: ~1.1 kg (total system mass)

– Acceleration (grid) voltage: ~1 kV

• Propellant

– Xenon gas (total stored Xe mass TBD)– Propellant contained in sealed container with pressure <100 psia (per requirements

Hybrid Plasma Probe Experiments

• Purpose

– To collect data on ionospheric plasma in perturbed and unperturbed regions within geophysically interesting areas of low earth orbit– To demonstrate that combination of several plasma diagnostics is feasible, efficient, and powerful

• Success criteria

– Resolve plasma environment to 30±5° about spin axis for at least 3 regions

of interest

• Operational scenario

– LionSat off, HPP booms in stowed configuration during launch

– LionSat powers up, HPP off, booms deployed for flight duration

– LionSat powered up, booms deployed, HPP on

– Electronics board prototyped, working C&DH interface

– Boom sensor to be protoyped Fall 2003

#1 Booms must be deployed after safely away from ICU/Launch Vehicle (NanosatProgram Requirement)

#2 Capacitance between sensors and surrounding guard must be minimized to mitigate fringing E-field effects

#3 Any HPP data must be time stamped to be correlated with orbital position(LionSat Mission Requirement)

#4 While HPP system is running the three inactive probes must be grounded (LionSat Mission requirement)

#5 Use of cables, lines, cords, plastic parts, or other non-metallic “soft good” in the primary structure as means of retaining deployable mechanisms, or in

instances where a failure of that part could be result in a hazardous situation Cf User Guide UN-0001Table 6-7 9(Nanosat Program requirement

#6 Probes must be cleaned Cleaning procedure prior to launch (LionSat Mission Requirement)

#7 Probe mass, volume and power usage

#8 Probe special needs and or regiments

Requirements Summary Table

Sub System: Power

Technology Lead(s): Yashar Fakhari, Jeff Wagner

SBLP Plasma Density, ne 109– 1012 m-3 108–1013 m-3 (est.)

SBLP Electron Temperature,

Te

development)FBLP High Speed Plasma

Solar cells:

#1 Solar Cells

GaAs, 18.3%, body mounted 8-sides + 2 end caps, 3 strings of 20 cells / sideOrganization responsible for lay-up/mounting: PSU (ref: UM-procedure)

1 Solar cell simulation for peak power tracker determination

2 Catalog all solar cells

a Determine total number of usable cells and separate them during testing

i Partly completed: 2306 total, 1950 sealed, 359 unsealed as of 4/04 (BSS)

b Perform visual inspections and record serial numbers

i Started 4/04

c Perform basic voltage/current test using test unit to verify functionality

i Started 4/04 using terrestrial simulator through Prof Wronski crw3@psu.edu

d Follow Michigan provided test plan

3 Develop LionSat specific lay-up plan using Michigan provided plan as a guide(communicate with structures group)

a Consider information found in LionSat SDR travel report

b Choose materials and suppliers for lay-up materials

4 String construction

a 20 cells per string

b Schottky barrier diodes on each panel to prevent current draw when not in sunlight

5 Power tracker construction? (TBD)

Number of solar cells

Mass of cells

Voltage and othe power things

Solar Cell Mounting

#1 Solar cell mounting technique

Nusil 2568 will be used to attach cells to panels and then cover glass will be used

to protect cells

#2 Solar cell arrangement

Cells will either be GaAs (baseline: 22 mm by 40 mm) or Si (alternate: 20mm by

40 mm)

End capsTop 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)

#3 Side panels

Each panel will hold 60–72 cells

“Belly band” ensures space for motorized booms and TM slot antennas

Battery

1 10 Sanyo N-4000DLR type 1.2V cells stacked to make 12V

2 **Each cell must have leads coming off of it so that each cell voltage can be individually tested after they are packaged

3 *Inhibit on battery return through light band

4 **Thermistors on each cell separate from computer monitor

5 Overall voltage and temperature sensor

6 Construction of battery box - Hazleton?

Charger:

1 Low power consumption for electronics

2 Charge battery from dead

3 Maintain battery charge with excess current from solar cells

4 Shut down when batteries are sourced for extra power

ii Minimum satellite draw in a low power mode

1 Recovery time from depth of dischargeiii Worst case satellite power draw

1 Recovery time from depth of discharge

iv Depth of Discharge for each operational mode (experiment combinations)

DC/DC Converter (on each subsystem)

1 Watch dog to power on computer when battery supply is sufficient to run computer on dark side (may put this on computer)

2 Switches for each subsystem to power on/off (none for computer)

a Read Sync Serial Port + chip select for subsystem on/off info

3 Outputs

a 1000V Ion-Thruster

b +5V Magnetometer/GPS/Computer/HPP