Loading presentation...

Present Remotely

Send the link below via email or IM


Present to your audience

Start remote presentation

  • Invited audience members will follow you as you navigate and present
  • People invited to a presentation do not need a Prezi account
  • This link expires 10 minutes after you close the presentation
  • A maximum of 30 users can follow your presentation
  • Learn more about this feature in our knowledge base article

Do you really want to delete this prezi?

Neither you, nor the coeditors you shared it with will be able to recover it again.


Make your likes visible on Facebook?

Connect your Facebook account to Prezi and let your likes appear on your timeline.
You can change this under Settings & Account at any time.

No, thanks

Aero 401 MCR/SRR

Combined Mission Concept Review and Systems Requirement Review

Shivani Ghai

on 5 December 2012

Comments (0)

Please log in to add your comment.

Report abuse

Transcript of Aero 401 MCR/SRR

Ultima Finium Inc. AggieSat X Mission Statement “The purpose of the Mission 3: A R & D Demonstration Project is to define multiple solutions for a spacecraft that is capable of performing the third mission in a sequence of four missions designed to demonstrate autonomous rendezvous and docking between two spacecraft, one built in the AggieSat Lab at Texas A&M University and one built at the University of Texas” History Mission 1 Mission 2 The first mission in this sequence –
consisting of AggieSat2 and the UT
spacecraft, Bevo-1 – was launched
on STS-127 on July 15, 2009. The
two picosats, both identical in size
(5”X5”X5”), were ejected from the
Space Shuttle Endeavor on
July 30, 2009. The mission was not
completed as planned. Mission 2 is more complex than the first
mission. Once again it is a TAMU/UT cooperative mission. This time, however, the two spacecraft
are not identical. AggieSat4 is approximately four cubic feet with a footprint of two square feet. The UT spacecraft, Bevo-2 - which fits inside AggieSat4 - consists of three 10cmX10cmX10cm CubeSats connected together. The expected launch date
is January 2013. Organizational Strategy AggieSatX Simulation Structures Tyler & Scott Power Andrew CDH Shivani & Kyle GN&C Olga & Mark Communication Kyle Gantt Chart Refinements Prior Work Refinement Structures Structures Requirements Launch Delta IV Load Factors for Dynamic Envelope Requirements ESPA Ring Max 400 lbs (181.4 kg), CG must be within 30 in (76 cm).
For our case 65 kg (well within mass
envelope of ESPA requirements)
Max envelope of 24”x24”x38”
(69.1 cm x 69.1 cm x 96.5 cm) including lightband Lightband Mark II 18 lightband
Planetary Systems Corporation
Cost: ~ $125k
18”(46 cm) diameter with 8 separation springs Materials & Design Current Solidworks Model Docking Mechanisms Electromagnet Probe and Drogue Easier to capture and release
Can switch magnet on/off and
reverse polarity for separation
Softer impact upon successful docking Easier to design
More reliable incase of electrical system failure
Does experience higher impact loads GNC GN&C Requirements Actuators Calculations Propulsion Power System EPS Architecture Energy Storage Calculations Communication Communication Requirements Communication and C&DH Conclusion Mass Allocation Expenses A shout out to the man behind it all! Thank you! Questions? Communication Architecture Store and Forward
With Crosslink Communications
Tracking, Telemetry and Command (TT&C)
Point to Point Data Dissemination
Bandwidth needs
Uplink commands
Downlink photos Link Budget Transmitter/Receiver ISIS TRXUV VHF/UHF
Uplink – 400 b/s
Downlink – 9600 b/s
HISPCO Transmitter
Downlink – 1 mb/s
UFH Uplink Receiver
Uplink – 1200 b/s Con-Ops Reconfigured
Thrusters and Placement
Satellite Design and Packing
Range and Bearing
Mission Flow Chart
Communications and CDH Link Budgets Establish Requirements
Material Analysis
Preliminary Design
Orbital Mechanics and Propellant Selection
Component Selection
Cost and Mass Analysis
Mission Viability
Risk Analysis Central Processor Andrew Model 160 High Performance Flight Computer
2 GB of Flash
Supports GPS daughter card
Two parallel digital camera interfaces Concept of Operations WBS Simulations, Testing and Manufacturing HOMER
Plume Testing
Battery Energy Simulations
GNC simulations
Testing Model Designed
Manufactured in-house Prior Work Thermal Control Conclusion Risk Analysis Looking Ahead Model plumbing and wiring in SolidWorks
Possibility of shrinking volume
Perform updated SolidWorks simulations
Potential design and modeling of a docking mechanism
Model internal components for thermal analysis Plume Testing Sensors Simulations Cameras Argus 1000 Infrared Spectrometer NanoCam C1U Infrared band
Actively Cooled
Remote Sensing Application 3MP sensor 10 bit color
Capable of data processing and storage
JPEG compression Minimum recommended frequencies:
Axial- 27 Hz
Lateral- 10 Hz
Max Acoustics:
130 dB over 150-500 Hz Budget: $5,000,000 Total Mass = 60.04 kg Solidworks Thermal Analysis 60 Minutes Exposed to Sun Solidworks Thermal Analysis 30 Minutes In Earth’s Shadow Nitrogen Gas Cylinder Airgas NI 40 Latch Valve Latch Valve Model 51-212 Fill Valve Flow Control Valve Model 51E186 Reduction Valve Regulator Model 50X713 Pressure Transducer Space Rated Series ST13000 Maneuvers Simulations Simulation – Position Simulated nonlinear point mass EOM in orbital frame
Input: Mass, Forces.
Output: dr, r, dthetaθ, theta θ Attitude Simulation and Control Linearized attitude equations of motion
Primary axis frame
Contains constant and periodic disturbances Maximum allowable reaction wheel momentum Preliminary Thermal Calculations Spacecraft Thermal Energy Balance
Qin = Qout
Qsun +Qer + Qi = Qss + Qse
Qsun=alpha_αs*As*Isun (solar input to spacecraft)
Qer=a*alpha_αs*Fs,se*As*Isun (earth reflected solar input)
a = earth albedo (0.07-0.85)
sigma = Stefan-Boltzmann constant
αalpha_s=spacecraft surface absorptivity
epsilon_Ɛs=spacecraft surface emissivity
Qi=internally generated power
Qse=σsigma*As*Fs,e(T^4sc – T^4e) net power radiated to earth
Qss= σsigma*As*Fs,s(T^4sc – T^4space) net power radiated to space
Fs,s= fraction of radiant energy leaving SC that is intercepted by space
Fs,e= fraction of radiant energy leaving SC that is intercepted by earth Assume:
Tspace ~ 0 and Fs,s + Fs,e = 1 Thermal Energy Balance becomes:
Ɛepsilon_s*sigmaσ*As*T^4s = epsilon_Ɛsσ*sigma*AsFs,eT^4e + Qsun +Qer + Qi Simulation – Attitude disturbance due to effectsof gravity gradient and magnetic field LVLH frame
Attitude states were fed back to convert the Earth’s magnetic field to the body frame and to feedback states into the controller Solar Arrays 3 large faces (56.8cm x 69.1 cm)
119 solar cells per large side
5 strings of ~25 cells per
One small (56.8cm x 56.8cm) face
98 solar cells per small side
4 strings of ~25 cells per
Maximum power: 223W
Average power: 151W
T = 94.6 minutes
Te = 0.38T = 36.0 minutes Orbital Mechanics 500km circular orbit
T = 94.6 minutes

Inclination Angle of 45 degrees
Te = 0.38T = 36.0 minutes Battery Sizing Battery Factors considered:
DOD: dependent on Mission Duration
Voltage larger than demand of components 8s2p configuration
Voltage is 32.8V
10A-hr PCDU Power Control and Distribution Unit Three Main Elements:
Battery Charge Regulator (BCR)
Power Conditioning Module (PCM)
Power Distribution Module (PDM) PCDU Diagram Wiring Gage of wire is based on the current demands for each component
Wiring is insulated with a polyimide coating
15.24 m (50 feet) 24 AWG wiring was approximated for our purposes Powering Components ESPA requirements met by current design

Lightband model: Mark II 18

Material of bus: Aluminum 6061-T6

Preliminary thermal calculations showed that components can remain in survivable ranges Stabilization with periodic and constant disturbances using reaction wheels
Example of maximum allowable pointing error Typical Performance Specifications

Outlet Pressure……......2 MPa

Weight…………..…………..0.30 kg

Filtration…………………….60 micron abs

Thermal Capability…..…−20°C to 70°C Typical Performance Specifications

Proof Pressure……………………..103 Mpa

Operating Voltage Range…….. 5 vdc

Weight………………………………… 0.23 kg

Thermal Capability………………−20°C to 120°C Typical Performance Specifications

Operating Pressure………………15 MPa

Operating Voltage Range…….. 28±4 vdc

Weight………………………………...0.2 kg

Thermal Capability…..………....0°C to 69°C Typical Performance Specifications

Operating Pressure……...……..0 to 18.6 MPa

Operating Voltage Range.......20 to 42 vdc

Weight…………………………………0.136 kg

Response Time…………..……….< 5 ms

Thermal Capability……………..−34"°"C to 60 °C Cylinder DOT Specification: 3AA-2015


Weight………………….10.7 kg

Service Pressure………13.89 MPa Computer Systems Power Management
Thermal Control
Attitude Control
Navigation Solutions
Command and Data Handling Data Architecture Centralized Architecture
Reliable Data-Flow Analysis Picture Data Mission Flow Chart Project Management Min: -23 C Max: 73 C Min:-11 C
Max:103 C Ultima Finium Inc. Architecture Looking Ahead Autonomous Flight Manager Current SolidWorks Model Current Solidworks Model Current SolidWorks Model Mass Properties Total Mass – 60.04 kg
Envelope – 76.2cm x 60.3cm x 60.2cm
Center of Mass (from geometric center)
x = 1.79 mm y = -3.55 mm z = 0.88 mm
Moments of Inertia (kg*m^2)
Ixx = 2.64 Iyy = 3.24 Izz = 3.70
Ixy = 0.14 Ixz = 0.015 Iyz = -0.075
First Mode Natural Frequency – 44.3 Hz Max Stress: 61 Mpa SolidWorks Testing – Stress Max Deflection: 0.96 mm SolidWorks Testing – Deflection Adding a 50% margin to each value, worst case temperature range is approximately 54.2 °C to -12.3 °C Conclusions:
Polished Aluminum for Earth Side
Solar Cell for Sun Side Preliminary Thermal Calculations Revisited - Full-scale
- Bus constructed from scratch using aluminum and fastened with screws and brackets
- Components modeled with foam
- Aided in visualization of interior and allowed for a better estimate of piping and wiring Satellite Mock-Up Sinclair Interplanetary – RW-1.0-28 Typical Performance Specifications

Nominal Momentum…………….. 1.0 Nm-sec

Nominal Torque…………………… 100 mNm Reaction Wheel Sinclair Interplanetary – TQ-15 Magnetic Torque Rod Typical Performance Specifications

Nominal Dipole………………….. 15 Am

Maximum Dipole….…………….. 19 Am CSS Sun Sensor Sun Sensor - Redundant attitude estimation

- Accuracy of 1°

- Optimal power SSBV GPS Receiver GPS/INS NV-GI100 2 GPS Units
4 GPS Receivers
- 4 reference points
- Increase position accuracy
- High fidelity attitude determination
- Built-in IMU for redundancy GPS ASC Inc. - TE-2809-PC TigerEye 3D Flash LIDAR

For ProxOps within 20 meters of Target

Sends pulses of light and calculates the time it takes for the light to reflect

Determines position & orientation of Target LIDAR Cold Gas Thruster Typical Performance Specifications

Nominal Thrust…………………….…....... 3.6 N (0.80 lbf)

Minimum Impulse………………..……..... 4.43 mNs

Maximum Operating Pressure………... 7.6 MPa Solenoid Actuated Thruster MOOG Model 58-118 MOOG Model 58-118 Nozzle Prototype Cold Gas Thruster Setup Plume Battery Energy Simulations Downlink Data-Flow Analysis Communications Summary Gain : 6 dBi
Opening Angle : 85°
Input Power : Up to 10 W
Frequency: 2.2 – 2.3 GHz Patch Antenna Gain : -2.5 dB
Input Power : Up to 10 W
Frequency : 120 – 900 MHz Bullet Antenna Link Budget NanoCam C1U
- 3MP sensor
- 10 bit color
- Internal data processing
- JPEG compression Optical Camera Picture Data Maximum pass time: 500 seconds/day
Useful pass time: About half Downlink Budget Main Uses
- Power Management
- Thermal Control
- Pressure Control
- Attitude Control
- Navigation Solutions
- Command and Data Handling Specifications
- 2 GB of Flash
- 64 MB of SDRAM
- Supports GPS daughter card
- Parallel camera interfaces
- Operating System : LINUX (C) Andrew Model 160 High Performance Flight Computer Computer System Color Code:
C&DH Cold Gas System *1 picture is approximately 0.3 MB Components: $1,066,000
Manufacturing: $700,000
Testing: $2,000,000 Leftover: $1,234,000 Leftover Margin: 4.96 kg -

- 2

2 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 Professor David Kanipe Typical Thermal Requirements for Spacecraft Components Worst Case Scenario HOMER Open loop vs Closed loop
Full transcript