Minimize ALL-STAR

ALL-STAR (Agile Low-cost Laboratory for Space Technology Acceleration and Research)

Students from the Colorado Space Grant Consortium (COSGC) have teamed with Lockheed Martin to develop a miniature satellite (3U CubeSat), known as ALL-STAR (Agile Low-cost Laboratory for Space Technology Acceleration and Research). The ALL-STAR program, designed to inspire and develop America’s future technological workforce, will provide students hands-on-experience in applying science, technology, engineering and mathematics skills to building operational space systems. The ALL-STAR concept is aimed at creating a small satellite bus with increased performance parameters.

The design of the ALL-STAR nanosatellite bus began in January 2010, the completion of the development phase is scheduled to occur in November 2011 and delivery of the system in the spring 2012.

Lockheed Martin is supporting the program and company engineers from Sunnyvale, CA, Palo Alto, CA, Newtown, PA, Albuquerque, N.M., and Denver, CO, are supplying their system engineering, program management and systems integration expertise to mentor the COSGC students as they design, develop, manufacture and deliver the CubeSat. 1)

COSGC, headquartered at the University of Colorado at Boulder, is a state-wide program that provides Colorado students access to space through innovative courses and real-world hands-on satellite programs. COSGC involves 16 Colorado colleges, universities and institutions and is funded by NASA as part of National Space Grant Program.


Figure 1: Illustration of the deployed ALL-STAR nanosatellite (image credit: COSGC) 2)

The ALL-STAR program is designed to accelerate the rate at which small research and technology based payloads are able to be launched by providing the satellite bus. Customers can purchase the ALL-STAR bus which includes the satellite structure, power, communications, attitude determination and control, command and data handling, and position knowledge.


The ALL-STAR nanosatellite is a low-cost modular 3U CubeSat form factor bus capable of supporting the 1 year on-orbit operation of a variety of spaceborne research payloads that can be configured and ready for flight in 6 months through a simplified payload hardware and software interface.

The nanosatellite consists of fully functional subsystems operating off the same amount of power as to that of a night light, designed to support one year of on-orbit operations; the bus has a size of 10 cm x 10 cm x 34 cm with a mass of ~ 4 kg. 3)

System overview: The subsystem interface has been designed so that each subsystem has an electrically identical interface to the Command and Data Handling and Flight Software systems, collectively known as CDH. The interface has three layers: the physical layer, the data link layer, and the network layer. The physical layer specification contains all requirements and details on the electrical aspect of the communication between subsystems and CDH. It will fully define the hardware interfaces need by each subsystem and CDH to complete a link. The data link layer specification details the low-level protocol needed to abstract the higher levels away from specific hardware limitations within the link. Finally, the network layer defines the high-level protocol used to transfer information and commands from one subsystem to another. This protocol forms the backbone of communication within the satellite.4)


Figure 2: The ALL-STAR nanosatellite in launch configuration (with and without solar cells), image credit: COSGC

Spacecraft structure

- Standard 3U CubeSat
- Payload volume: ~ 1U; payload allotted mass: 2 kg
- Interface structure provided
- Physical size: 10 cm x 10 cm x 34 cm (launch configuration)
- Total launch mass: ≤ 4 kg

EPS (Electrical Power Subsystem)

- ≥ 10 W nominal draw (orbit average)
- 30 W for 15 min peak support period
- Payload: 4.5 W nominal draw

Attitude knowledge and pointing

- Knowledge accuracy: ≤ ±0.2º (1σ)
- Pointing accuracy: ≤ ±1º (1σ)

Position knowledge

- Onboard GPS receiver, knowledge: ≤ ±100 m

Command and data handling

- Handles all subsystem interactions
- Memory storage for payload
- Reads digital and analog sensors from payload


Modular, easily-changeable design

RF communications

- 250 kbit/s (min) information rate; goal of 1 Mbit/s information rate

Optional propulsion module

- ΔV of ≥ 10 m/s
- Part of the payload section

Mission lifetime

≥ 1 year

Spacecraft flight configuration

- Deploys solar panels and external structure
- PEZ (Payload Extension Zone) split 50/50 between bus and payload

Spacecraft flight orientation

Attitude primarily determined by the payload requirements

Table 1: Overview of ALL-SAT performance requirements 5)


Figure 3: Designation of spacecraft elements (image credit: COSGC)

STR (Structures):

The nanosatellite is comprised of three separate integrated structures (Figure 4): 6)

• Payload section

• Bus section

• External solar panel shell with deployable PEZ (Payload Extension Zone)

The PEZ system allows for the CubeSat to deploy its external solar array structure, exposing the internal structure and the science payload section, while still conforming to the 3U CubeSat standard during launch. In order to accomplish this, the ALL-STAR team integrated simple mechanical elements that make up the three phases of the deployment system. These phases are restraint & activation, mechanical deployment, and locking mechanisms. 7) 8)

Restraint & activation: The deployable solar panels are constrained during launch by three mechanisms. The first of these mechanisms is the simplest in design consisting of simply a fixed clip that is attached to the internal structure that hooks into a bracket on the back of the outermost solar panel. The middle of the solar panels are restrained through the use of a modified hinge that restrains the outer most panel from rotating away from other panel until both panels have rotated through a minimum angle. The final restraint is a rotating claw mechanism that hooks into the solar panels in the same way as the fixed clips but in distinctly different. The PEZ deployment is restrained during launch through the use of a mini-Frangibolt actuator.


Figure 4: Exploded view of the ALL-STAR nanosatellite in flight configuration (image credit: COSGC)

The main deployment of the PEZ utilizes two constant force springs, shown in Figure 5, that are held in tension in the launch configuration. When the Frangibolt fractures, these constant force springs will begin to retract. The two structures interface through smooth sliding surfaces that allow the springs to extend the structures to the final configuration.


Figure 5: Constant force spring implementation (image credit: COSGC)

Once the ALL-STAR structure has fully deployed, it is important that the satellite remains as ridged as possible and that the mechanisms for the deployment don’t collapse. To prevent such events, the mechanical team has also incorporated several locking mechanisms. The first mechanism is an integrated spring plunger and hard stop on the external structure. The hard stop insures that the constant force springs can’t over deploy the structure to a point in which the structures would separate. Also the integrated spring plunger falls into a locking hole that insures that the structures don’t re-collapse.


Figure 6: Functional block diagram of the nanosatellite structure (image credit: COSGC)


Figure 7: Bus electronics layout (image credit: COSGC)

ADCS (Attitude Determination and Control Subsystem):

ADCS, also referred to as ACS, provides 3-axis spacecraft stabilization. Attitude sensing is provided by a 3-axis magnetometer, a star tracker, and gyroscopes; actuation is provided by reaction wheels (3 wheels mounted orthogonally) and magnetorquer wheels (used for momentum dumping). ADCS is capable is support of the following maneuvers: 9)

• Ground target tracking: short duration, quickly changing the target vector (> 60º in < 10 minutes)

• Sun spin:

- Perform spin with minimal losses (zenith within 10º of sun vector)

- For orbital altitude of 350 km: 5 min for 180º

• Change of target vectors: (used for payload pointing and RF communications)

- 2 minutes (~2.5% of orbit)

- 75º half angular distance across Earth

• Correct external torques: Max aerodynamic torque of ~6 x 10-5 Nm @ 350 km altitude.


Figure 8: Components of the ADCS (image credit: COSGC)

ADCS microcontroller: AVR32UC3 (commonality with CDH)

• Tracks attitude knowledge and commands

• Implements control law and controls actuators

• SPI (Serial Peripheral Interface)

- Receive inertial reference from star camera

- Read gyroscope ADC

- Command motor and magnetorquer controllers

• Two wire interface(I2C): Read magnetometer.

Star tracker:

• Provides inertial attitude reference to attitude determination system

• Built in-house due to cost constraints

• Image processing and star identification performed using a dedicated FPGA (Field Programmable Gate Array).

GPS receiver:

• Selection of the COTS GPS receiver, the Novatel OEMV-1.

• GPS antenna: VTGPSIA-3 antenna with built in LNA

• The receiver provides both position and time solutions to the onboard subsystems.

CDH (Command and Data Handling) subsystem:

The CDH subsystem provides wire-free integration with subsystems and payload.

• Provides 32-bit microcontroller to FSW (Flight Software). Also utilizes 8-bit microcontroller for interface support to power switching components and GPS

• MRAM (Magnetoresistive Random-Access Memory) provides flexible SRAM (Static Random Access Memory ) extension / configuration storage space

• Micro SD (Secure Digital) cards allow for expandable storage in small form factor.

• Interfaces:

- Provide each subsystem an identical electrical interface

- Payload supplied data bus, digital and analog sensors/indicators

- System watchdog helps mitigate SEUs (Single Event Upsets) and other lockup conditions

- 1-Wire temperature bus allows temperature sensing with satellite unpowered.




Temperature range




-40ºC to +85ºC

GSE Interface



-40ºC to +85ºC

Data memory



-25ºC to +85ºC

Configuration memory



-40ºC to +85ºC

Payload ADC

Texas Instruments


-40ºC to +85ºC

GPS interface microcontroller



-40ºC to +85ºC

Subsystem connector


LS2-125-01-S-D –RA1


Backplane connector




Table 2: CDH components list


Figure 9: Schematic view of CDH configuration (image credit: COSGC)


Figure 10: Functional block diagram of CDH (image credit: COSGC) 10)


Figure 11: Overview of CDH interfaces (image credit: COSGC)

EPS (Electrical Power Subsystem):

EPS design features: 11)

• 10 W nominal power (5 W to the bus, 5 W to the payload)

• 30 W peak power (in COM passes, etc.) for 15 minutes / 2 orbit

• Total of 918 solar cells: (cell efficiency of 27%)

• 4 Li-Ion battery cells, 2 sets in parallel; cell capacity of 2.6 Ah at 3.65 V each

• The power bus is divided into 3.3 V and 12 V unregulated battery lines

• Solar interface: MPPT (Maximum Power Point Tracking) from solar cells to batteries.

• The EPS is monitored and controlled by the master processor.


Figure 12: Illustration of the EPS (image credit: COSGC)

RF communications:

A full duplex communications subsystem is implemented, featuring SDR (Software Defined Radio) architecture, S-band (2.4 GHz) downlink with a data rate of 250 kbit/s with BPSK (Binary Phase Shift Keying) modulation and convolutional coding (Reed Solomon). The uplink is in UHF (435 MHz) with a data rate of 9.6 kbit/s. In addition, a CW beacon is implemented. 12)


Figure 13: Schematic view of the RF communications subsystem (image credit: COSGC)

PROP (Propulsion Subsystem):

PROP is an optional independent module that can be attached to the payload section of the ALL-STAR bus. The PROP design features and capabilities are: 13)

• A cold gas propulsion system (butane propellant with a mass of 72 g)

• 1 valve/nozzle provides 0.05 N of thrust

• Achieve a minimum ΔV of 10 m/s

• Approximate time for ΔV of 10 m/s is 190 minutes.

The total mass budget of PROP is 485 gram.


Figure 14: Functional block diagram of PROP (image credit: COSGC)


Figure 15: Illustration of the propulsion subsystem (image credit: COSGC)

Launch: ALL-STAR / THEIA was selected to participate in NASA’s ELaNa-5 (Educational Launch of Nanosatellites) initiative, and has been manifested on the CRS-03 (Commercial Resupply Services) mission to the ISS on a Falcon 9/Dragon vehicle of SpaceX, expected to launch in the fall of 2012. 14)

Orbit: Circular orbit, altitude of ~ 325 km, inclination = 51.6º, period: ~ 91 minutes.

Mission title

University / Institution

CubeSat size


MEC (Medgar Evers College) and CUNY (City University of New York)


KYSat–2 (Re-flight)

Kentucky Space


Hermes–2 (Re-flight)

University of Colorado, Boulder





TechCube-1 (Technology Demonstration CubeSat-1)




COSGC (Colorado Space Grant Consortium)





Ho‘oponopono Radar Calibration CubeSat

University of Hawaii


Table 3: Summary of secondary payloads manifested on ELaNa-5 (Ref. 14)

Sensor complement: (THEIA)

THEIA (Telescopic High-definition Earth Imaging Apparatus):

THEIA, named for the titan goddess of sight in Greek mythology, is an optical remote sensing payload, compatible with the COSGC ALL-STAR bus, that will take images of Earth's surface in full color in order to verify all capabilities of the ALL-STAR bus on orbit, providing a marketable proof of concept. 15) 16) 17) 18) 19) 20) 21)

The THEIA instrument is being designed and developed by the THEIA team at the University of Colorado (UC) and at COSGC (Colorado Space Grant Consortium).

THEIA must meet the mass and volume constraints, as stated in the ALL-STAR ICD (Interface Control Document). The payload constraints of the ALL-STAR bus structure are:

• Volume (max): 1215 cm3

• Length (max): 16.5 cm (payload+ PROP)

• Mass (max): 2 kg (including the payload structure).


Figure 16: ALL-STAR bus payload compartment constraint model (image credit: COSGC)

The THEIA imaging system is comprised of a refractor and a CMOS imaging sensor. The refractor chosen for this design is an achromatic doublet lens that measures 50.8 mm in diameter and has a 150 mm focal length. An achromatic doublet lens is composed of two lenses that made of two types of glass that are cemented together with optical-grade adhesive, as shown in Figure 17. Placing two types of lenses in extremely close proximity is a common optical practice for image clarity since the second lens effectively reverses the prism effect of the first. Furthermore, the diameter and focal of the chosen refractor fits within the volume constraints, but utilizes of the majority of the allotted volume to achieve the required angular resolution and field of view.

Baseline mission:

• 50 mm aperture refractor telescope

• Full color, Earth imaging (spectral range: 0.4-1.0 µm)

• 1º FOV (Field of View)

• Required angular resolution of 0.04º

• Goal angular resolution of 0.0017º

• 5 Mpixel CMOS camera.


Figure 17: Schematic view of the refractor imaging system (image credit: COSGC, UC)

The full structural system design is shown in Figure 18 with the key components of the THEIA design labeled. The payload structure is a dual-element structure with nearly equal lengths of Invar-36 and Titanium-64. The thermal expansion and compression of this dual-element structure closely follows the change in focal length of the optics system with temperature. Furthermore, the light baffle structure can be seen between the refractor and the imaging sensor. Within the structure, four light baffles are used to eliminate off-axis light that enters THEIA from the edge of the refractor.

The lens and electronics mounting is also visible. The single optic will be mounted with Invar-36, which exhibits minimal expansion with temperature and will produce minimal rotational and translational misalignment at differing temperatures. The CMOS board is mounted at the focal point of the refractor, and the FPGA is in its own housing structure on one side of the structure. FPGA is mounted on the side of THEIA to allow for maximum focal length between the refractor and imaging sensor.


Figure 18: CAD model of the THEIA imaging system (image credit: COSGC, UC)

The system functional block diagram is displayed in Figure 19. It can be seen that the ALL-STAR bus provides 3.3 V, 12 V, and unregulated battery power line to THEIA. THEIA will then regulate and distribute the provided to the CMOS imaging sensor, the FPGA processor, and supporting electronics. When THEIA is taking an image, incoming light will be focused on the imaging sensor by the refractor and captured with the imaging sensor’s electronic shutter. The image will then be sent over I2C interface to the FPGA, which will process the image. Finally, the processed image and sensor data will be transferred to the ALL-STAR bus over an SPI interface and saved. On orbit, the ALL-STAR bus will downlink the data on the next available ground pass.


Figure 19: Functional block diagram of THEIA (image credit: COSGC, UC)

Concept of operations:

The THEIA project has three separate concepts of operations, depending on the situation at hand. The first concept of operations describes the imaging process of the THEIA system, and can be seen in Figure 20. Note that the process repeats itself every time an image is taken at the maximum rate of one image per minute.


Figure 20: THEIA concept of operations - image capture diagram (image credit: COSGC, UC)

In orbit, the ALL-STAR bus will orient in the direction of the desired object (target area) before THEIA captures an image. After THEIA captures an image, THEIA will transfer the processed image to the ALL-STAR bus, which will store the image until the next COM pass when it will downlink the processed image.


Figure 21: Illustration of THEIA's FOV (image credit: COSGC, UC)

In order to validate the pointing capabilities and pointing knowledge of the bus, the payload must be able to see the desired target even if the bus is off by its maximum error which is 1º. Figure 21 illustrates how the FOV is related to the pointing accuracy of the ALL-STAR bus and what this covers on the ground. The pointing knowledge of the ALL-STAR bus can be validated if the resolution of the telescope is around 5 times better than the bus’ knowledge. Therefore the angular resolution of the telescope must be at least 0.04º. The goal however is to be able to see much finer detail than what would be seen with 0.04º of angular resolution, so an additional stretch goal requirement was set based around the goal of seeing a football field sized object on the ground.


Figure 22: Orbit concept of operations illustrating the life cycle of the mission (image credit: COSGC, UC)

1) Launch

2) ALL-STAR start-up mode

• The ALL-STAR bus is deployed from P-POD and launch vehicle

• The ALL-STAR bus solar panels receive power and deploy. The ALL-STAR bus subsystems turn on and check health and status

• During the first high speed COM pass, the ground operator checks health and status of the ALL-STAR bus and approves it for normal operation.

3) COM (communications) pass 2

• The ground operator turns on THEIA

• The health and status of THEIA is checked by CDH and relayed to the ground.

4) Payload operations mode

• The ALL-STAR bus identifies when the satellite approaches the desired picture location, and points THEIA to the ground

• THEIA takes multiple pictures of the ground

• THEIA processes images and sends data to the ALL-STAR bus

• The ALL-STAR bus returns to nominal Nadir attitude.

5) Typical COM pass mode

• The ALL-STAR bus sends health and status of the ALL-STAR bus and THEIA

• The ALL-STAR bus sends processed images, data, and corresponding GPS location to ground station

6) End of life

• The ALL-STAR bus attitude system orients ALL-STAR to maximize drag

• The ALL-STAR bus and THEIA are shut down.


Figure 23: Exploded view of the THEIA optical payload (image credit: University of Colorado)

Legend to Figure 23: The left side displays the optics lens mounting which attaches to the light baffle assembly. At the end of the light baffle tube is the CMOS image sensor board (green). The optics assembly is supported by an invar plate (brown, bottom of picture) and is housed in aluminum.

1) “Colorado Space Grant Consortium, Lockheed Martin To Develop CubeSat,” Lockheed Martin, Aug. 10, 2010, URL:


3) ALL-STAR documentation package, URL:

4) Jessica Brown, Riley Pack, Tyler Murphy, Ryan Kophs, “Systems and Management Backup Documentation,” ALL-STAR System Requirements, SYS.200.2, Jan. 3, 2011

5) Jessica Brown, Riley Pack, Tyler Murphy, Ryan Kophs, “Systems and Management,” Critical Design Review Slides, URL:

6) Simon Hodgson, Jon Kanaber, Matt Gosche, “Structures and Mechanisms,” Critical Design Review Slides, URL:

7) Tyler Murphy, Chris Koehler, “PEZ: Expanding CubeSat Capabilities through Innovative Mechanism Design,” 2011 COSGC Space Research Symposium, URL:

8) Tyler Murphy, Jon Kanaber, Christopher Koehler, “PEZ: Expanding CubeSat Capabilities through Innovative Mechanism Design,” Proceedings of the 25th Annual AIAA/USU Conference on Small Satellites, Logan, UT, USA, Aug. 8-11, 2011, paper: SSC11-XII-5

9) Aaron Russert, Brian Roth, Jamey Graham, Stanislav Stoytchev, “Attitude Determination and Control,” Critical Design Review Slides, URL:

10) Riley Pack, “Command and Data Handling,” Critical Design Review Slides, URL:

11) Greg Stahl, Eric Huckenpahler, “ EPS (Electrical Power Subsystem),” Critical Design Review Slides, URL:

12) Riley Pack, Logan Finch, Logan Smith, Eric Pahlke, “Communications,” Critical Design Review Slides, URL:

13) Carolyn Maurus, Jackson Beall, Jesse Ellison, “Propulsion,” Critical Design Review Slides, URL:

14) Garret Skrobot, “ELaNa - Educational Launch of Nanosatellite,” 8th Annual CubeSat Developers’ Workshop, CalPoly, San Luis Obispo, CA, USA, April 20-22, 2011, URL:

15) Christopher Koehler, Scott Palo, Herbert Kroehl, Jean Koster, Russ Mellon, Peter Bradley, Jessica Brown, Justin Clark, Nicole Doyle, Kyle Florentine, Kyle Kemble, Carolyn Maurus, Jennifer McGraw, Brian Roth, Malcolm Young, “Telescopic High-definition Earth Imaging Apparatus (THEIA),” Project Final Report (PFR), April 28th, 2010, URL:

16) “Telescopic High-definition Earth Imaging Apparatus,” URL:

17) “Telescopic High-definition Earth Imaging Apparatus (THEIA), Conceptual Design Document,” University of Colorado, Sept. 23, 2010, URL:

18) “Telescopic High-definition Earth Imaging Apparatus,” CDR (Critical Design Review), Dec. 1, 2010, URL:

19) “THEIA - Telescopic High-definition Earth Imaging Apparatus,” Project Definition Document (PDD), Sept. 9, 2010, URL:

20) “An Athermal System Design to Improve Optical Performance across a Mission Profile,” April 9, 2011, URL:

21) “Telescopic High-definition Earth Imaging Apparatus,” URL:

The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.