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Satellite Missions Catalogue


Jun 14, 2012


Quick facts


Mission typeNon-EO

ST8 (Space Technology 8)

ST8 is a NASA technology demonstration mission within NMP (New Millennium Program) managed by NASA/JPL. In January 2005, NASA selected four teams to develop a suite of advanced technologies, these are: 1) 2) 3) 4)

1) NGU (Next Generation Ultraflex Solar Array System). NGU is being developed by ATK Space Systems (formerly AEC-Able Engineering, Inc., Goleta, CA).

2) SAILMAST developed by AEC-Able Engineering, Inc. The SAILMAST is an ultra-light graphite mast (40 m) intended for potential solar sail propulsion as well as other applications.

3) MLHP (Miniature Loop Heat Pipe Small Spacecraft Thermal Management System). MLHP is being developed by NASA/GSFC. The MLHP can transport large heat loads over long distances with small temperature differences and without external pumping powers to provide precise temperature control and reduce the need for supplemental heaters.

4) DM (Dependable Multiprocessor), developed by Honeywell International, Inc., Clearwater, FLA. The DM will provide high rate onboard processing for science data and autonomous control functions using a COTS (Commercial-Off-The-Shelf) electronics package. Note: DM was formerly known as EAFTC (Environmentally Adaptive Fault Tolerant Computing System).


The ST8 spacecraft is based on a hybrid of the MicroStar (electronics) and LeoStar (structure) platform of OSC (Orbital Sciences Corporation Inc.), Dulles, VA, as prime contractor. The spacecraft structure consists of an aluminum honeycomb bus, a vertical payload module mounted on top, and 4 deployable solar array panels attached to the hexagonal bus. The bus structure contains all the spacecraft bus components, including electronics, battery, sensors, and actuators.. The solar arrays generate an average power of 380 W (550 W at a 32º sun angle). The mission design life is 7 months. The spacecraft has an estimated launch mass of 246 kg (there is no propulsion system available for orbital adjustments). 5) 6) 7)

The spacecraft is 3-axis stabilized (momentum biased). It will be maintained in a nadir-pointing attitude throughout the mission, with the SAILMAST pointing at zenith and the Ultraflex 175 on the anti-velocity side of the spacecraft. A single Y-axis reaction wheel provides a momentum bias and pitch control. Magnetic torquers are used for roll/yaw control and to dump reaction wheel momentum. The pointing requirements are very loose: 5º in control and knowledge and 2º knowledge of angle from Ultraflex to sun vector. Neither gyro nor star tracker is required, so the attitude sensors are limited to coarse and fine sun sensors, Earth sensors, and a magnetometer.

The command and data handling architecture is comprised of a set of three distributed processors; these are the flight computer, the battery control regulator, and the attitude and control electronics. A mission interface unit will provide RS-422 data interfaces to each experiment and will provide data storage. An onboard GPS receiver is used for spacecraft position and time knowledge.

Figure 1: The ST8 spacecraft and its experiment complement (image credit: NASA, OSC)
Figure 1: The ST8 spacecraft and its experiment complement (image credit: NASA, OSC)


- Hexagonal bus & payload module with aluminum honeycomb panels
- Deployable, fixed orientation solar arrays


- Four deployable solar arrays

- Triple-junction GaAs solar cells, average power of 380 W
- Battery: 21 Ah capacity

ADCS (Attitude Determination and Control Subsystem)

- Sensors: Fine sun sensor, Earth sensors, coarse sun sensors, magnetometer
- Actuators: Reaction wheel, magnetic torquers

CDHU (Command & Data Handling Unit)

- Distributed architecture
- GPS receiver for time and position knowledge
- 1 gigabyte (GB) memory

RF communications

- S-band: 32 kbit/s uplink; 2 Mbit/s CCSDS downlink
- Dual omni-directional antennas for near spherical coverage

Payload accommodation

- 28V power supply
- High-speed data interface
- Ancillary data

Spacecraft launch mass

< 250 kg

Nominal mission duration

7 months (1 month checkout period, 6 months of experiments)

Table 1: Overview of some spacecraft parameters

Orbit: Sun-synchronous elliptical orbit of 300 km x 1300 km, inclination = 98.5º, local equatorial crossing at 6:00 hours, the orbital period is 101 minutes.

Launch: A launch of the ST8 spacecraft is planned for late 2009 on a Pegasus-XL vehicle (air launch from an L-1011 carrier aircraft) of OSC.

RF communications: Uplink (32 kbit/s) and downlink (2 Mbit/s) communication is provided by an S-band transceiver and dual omni antennas.

Mission operations are conducted out of Orbital's MOC (Mission Operations Center), Dulles, VA, using the MAESTRO (Mission Adaptive Environment for Spacecraft Test and Real-time Operations) package of OSC. During normal operations, data will be sent from the MOC of Orbital to the experimenters' facilities via Internet.

The current operations scenario calls for performing the Ultraflex 175 deployment and experiment operations as the first experiment activity following a one month spacecraft checkout. This activity will only take a few days, at most. The Thermal Loop and DM experiments will both be performed over a longer term, perhaps for much of the 6 month experiment operations phase, in order to collect data during several of the experiments' modes. The final experiment activity will be the SAILMAST deployment and subsequent structural tests. This experiment is performed last to reduce the risk of a failure impacting the spacecraft and other experiments.

Figure 2: Alternate view of the ST8 spacecraft (image credit: NASA, OSC)
Figure 2: Alternate view of the ST8 spacecraft (image credit: NASA, OSC)

Experiment complement: (NGU, SAILMAST, MLPH, DM)

NGU (Next Generation Ultraflex Solar Array System), or Ultraflex 175 experiment:

NGU is designed and built by ATK Space Systems, in collaboration with the NASA/GRC (Glenn Research Center and Emcore Photovoltaics (EPV). The NGU is an ultra-lightweight flexible-blanket (accordion fanfold) solar array that deploys to provide a significant advancement in performance over existing state-of-the-art for high power arrays. The Ultraflex 175 array is comprised of ten primary interconnected isoscelestriangular shaped ultra-lightweight substrates (Figure 3). When Ultraflex 175 is delpoying in a rotational "fan" fashion, each interconnected triangular shaped substrate (also known as a gore) unfolds; upon full deployment the structure becomes tensioned into a rigid shallow umbrella-shaped structure. Radial spar elements attached to each substrate elastically deflect to predetermined positions when completely deployed to maintain the deployed structure in a preload and high-stiffness state.

The triangular gores are the building blocks of the Ultraflex 175 system, consisting of an open mesh Vectran substrate to which the cell circuits are bonded using a patented ultralightweight.

process. The use of the lightweight open mesh substrate allows the Ultraflex 175 to have a very low non-power producing mass per unit area, and allows the cells to radiate directly from their (partially open) backsides. When stowed, the Ultraflex 175 array gore substrates are folded in a flat-pack accordion manner and sandwiched between two rigid panels (static and pivot panels) to produce a compact launch volume.

The Ultraflex 175 experiment (3.2 m diameter) is an evolved version of ATK's 1st generation Ultraflex design developed for the MSP (Mars-01 Lander Program). The proposed Ultraflex 175 baseline system offers exceptional scaled-up performance exceeding the NMP ST8 requirements, and provides many additional features desired by NASA. In addition, many breakthrough technologies are proposed to be developed/implemented within the Ultraflex 175 system to provide significant performance growth capability (see Table 2). The proposed Ultraflex 175 technology enhancement and resulting performance growth will demonstrate the potential and flexibility of the Ultraflex 175 to be the standard high-performance solar array platform for future NASA, commercial and military space missions. 8) 9)

Figure 3: Illustration of major Ultraflex 175 subsystems and assemblies (image credit: NASA, ATK)
Figure 3: Illustration of major Ultraflex 175 subsystems and assemblies (image credit: NASA, ATK)


ST8 requirement and/or NASA need

Ultraflex 175 predicted performance for a 7 kW sized wing system


Specific power at BOL

> 175 W/kg for a 7 kW wing system

> 175 W/kg for NGU-S > 220 W/Kg for NGU-LW

ST8 requirement met with standard MJ PV. Much higher specific power achievable with lightweight MJ PV. The Ultraflex 175 platform provides "road map" growth beyond the ST8 requirement

Deployed first mode frequency

> 0.1 Hz

> 0.3 Hz

Extremely lightweight tensioned structural platform with sufficient depth provides deployed stiffness significantly higher than classical systems

Stowed specific volume

> 31.8 kW/m3

> 33 kW/m3

Ultraflex 175 occupies an extremely compact launch volume and footprint compared to classical systems

High voltage capability

> 100 VDC operation

> 100 VDC operation

Ultraflex 175 is inherently suited to high voltage operation because of its serpentine circuit configuration/layout and lack of conductive substrate. Conventional high voltage design solutions can be applied to Ultraflex 175 at lower mass than classical systems

Multi-AU operation

to 5 AU (Astronomical Unit)

to 5 AU

Ultraflex 175 employs use of EPV ATJ PV. These devices have undergone preliminary LILT testing and shown their suitability for multi-AU missions. Ultraflex 175 high-temp capability materials also allow for < 1.0 operation.


7 kW wing size

to 15 kW wing size and beyond

The Ultraflex 175 is scalable to wing sizes beyond 7 kW with the simple optimization of spar structural elements (to optimize bending and torsional stiffness). Recent studies for advanced JPL applications have sized Ultraflex 175 wings as high as 12 kW.

Stowed transfer orbit power


Provides stowed transfer orbit power

Ultraflex 175 provides the ability to incorporate stowed transfer orbit power (need for most GEO comsat missions), enhancing commercialization (addressed in study phase)

Radiation hardness

Operation in high radiation environment

Operates in high radiation environments

Ultraflex 175 employs proven radiation hard MJ PV technologies. Rear side and front side shielding can be applied to increase radiation hardness.


High reliability

High reliability

Ultraflex 175 is a high-reliability system because of its simple and redundant mechanisms and features, and its 1st generation heritage from the Ultraflex Mars 01-Lander system


Low (reasonable)

Low (reasonable)

Ultraflex 175 has the ability to be cost-competitive with classical systems once completely developed and its significant mass benefits are included in overall mission costs. The simplicity of the mechanical design and low cost substrates should allow for Ultraflex 175 recurring costs to be no greater than classical systems.

Table 2: Ultraflex 175 performance versus ST8 requirements and NASA needs

SAILMAST experiment:

An experiment under development for the ST8 flight program is referred to as SAILMAST (Scalable Architecture for the Investigation of the Load Managing Attributes of a Slender Truss). The SAILMAST experiment plan provides a thorough investigation of the fundamental attributes of an ultra-lightweight (slender) coilable truss, allowing extrapolation to generalized gossamer truss structures that may be different in geometry, loading, or design (Figure ). Obviously, this technology is particularly suitable for solar sailing applications.

The SAILMAST experiment will provide validation of the most fundamental building block of gossamer space structure technology and, in particular, the essential element needed for near term solar sailcraft to support key NASA roadmap missions. The ST8 mission will use a SAILMAST of 40 m in deployed length intended as a future solar sail mast built by ATK Space Systems. The mast diameter is 24 cm. In its stowed phase, the SAILMAST is < 1% of its deployed length. The linear mass density of this structure is 34 g/m. 10) 11)

Figure 4: Illustration of the SAILMAST concept (image credit: NASA)
Figure 4: Illustration of the SAILMAST concept (image credit: NASA)

MLHP experiment:

MLHP (Miniature Loop Heat Pipe Small Spacecraft Thermal Management System) developed by NASA/GSFC. MLHP is a miniature loop heat pipe consisting of two evaporators, two condensers, a body mounted radiator and a deployable radiator. Other elements include deployable radiators, TECs (Thermoelectric Coolers) on the LHP compensation chambers (CCs), a capillary flow regulator, and an aluminum coupling block between the vapor line and liquid line. 12)

The two most important features of the Thermal Loop layout are the integration of multiple evaporators into a single LHP (Loop Heat Pipe), and the use of miniature evaporators. The Thermal Loop combines the functions of variable conductance heat pipes (VCHPs), thermal switches, thermal diodes, and state-of-the-art LHPs into a single integrated thermal system. Major technology advances of the system design are:

1) Miniaturization of the evaporator, i.e. reducing the evaporator diameter from 25 mm to 13 mm

2) Multiple evaporators and multiple condensers in a single LHP

3) TECs for temperature control and start-up success

4) A transient LHP model and scaling rules.

The objectives of the MLHP flight experiment and ground modeling are to demonstrate the following items:

- The operation of MLHP with two evaporators and two condensers

- Reliable and repeatable MLHP start-up and shut down

- MLHP operation when the two evaporators receive various heat loads (even and uneven heat loads)

- Heat load sharing between the two evaporators

- MLHP operation when two radiators are exposed to different thermal environments

- The ability of the TECs to control the Compensation Chamber (CC) temperature

- Analytical models that predict MLHP thermal performance in one-G and zero-G environments

- Scaling rules that establish applicability and scalability of the MLHP technology for future missions.

Figure 5: Schematic of the thermal loop system (image credit: NASA)
Figure 5: Schematic of the thermal loop system (image credit: NASA)

Principle of LHP operation: An LHP utilizes boiling and condensation of the working fluid to transfer heat, and surface tension forces developed by the evaporator wick to circulate the fluid. It can transport large heat loads over long distances with small temperature differences. This process is passive and self-regulating in that the evaporator will draw as much liquid as necessary to be completely converted to vapor according to the applied heat load.

When multiple evaporators are placed in parallel in a single loop, each evaporator will still work passively. No control valves are needed to distribute the fluid flows. All evaporators will yield the same vapor temperature as liquid vaporizes inside individual evaporators regardless of their heat loads. The loop provides a single interface temperature for all instruments. Furthermore, when an evaporator is exposed to a heat sink, such as when the attached instrument is turned off, the evaporator will receive heat from other evaporators servicing the operating instruments.

This will eliminate the need for supplemental electrical heaters while maintaining all instruments close to the saturation temperature. The evaporators can automatically switch between evaporating and condensing modes based on the surrounding thermal conditions. Therefore, each instrument can operate independently without affecting other instruments.

The fluid flow distribution among multiple, parallel condensers is also passive and self regulating. Each condenser will receive an appropriate mass flow rate so that the conservation laws of mass, momentum and energy are satisfied in the condenser section.

Technology item


Thermal Loop technology advances

Integral Thermal Subsystem - Thermal Loop with TECs on CCs

Louvers, Heat Pipes, LHPs, Heaters, Thermostats

Flexible locations of heat dissipating components, heat load sharing, TEC for temperature control and start-up enhancement

LHP configuration

Single evaporators

Multiple evaporators

LHP evaporator outer diameter

25 mm

13 mm

Analytical modeling of LHPs

- Top-level transient models for single evaporator LHPs
- No scaling rules

-Detailed transient models for multi-evaporator LHPs
- Scaling rules established

LHP start-up method

Starter heaters on evaporator (20 W to 40 W)

TEC on CC (< 5 W)

LHP temperature control

- Control heater on CC; cold biased,
- Heating only, no cooling
- Heater power: 5 W to 10 W

- TEC on CC plus coupling blocks on transport lines
- Both heating and cooling
- Heater Power: 0.5 W to 2 W

Table 3: Summary of technology advances of MLHP

DM (Dependable Multiprocessor) experiment:

The DM project was formerly known as EAFTC (Environmentally Adaptive Fault Tolerant Computing System). The goal of the Dependable Multiprocessor experiment, developed at Honeywell, is to provide a spacecraft/payload processing capability 10x - 100x what is available today, enabling heretofore unrealizable levels of science and autonomy. The objective is to combine high-performance, fault tolerant, COTS-based cluster processing and fault tolerant middleware in an architecture and software framework capable of supporting a wide variety of next-generation mission applications. The DM project builds on earlier projects at JPL, Honeywell and Raytheon, which were sponsored by NASA, DARPA, and USAF. 13) 14)

A cluster computer comprises a set of single board computers, interconnected by a high speed switched network, running a file-oriented multi-threading operating system and a "middleware" which controls and coordinates parallel processing applications. DM employs a Linux operating system (Monte Vista version) common to all processors in the cluster (Figure 8).

Dependable Multiprocessor technology comprises four key elements:

• An architecture and methodology which enables the use of COTS-based, high-performance, scalable, multi-computer systems in a space environment, incorporating reconfigurable co-processors, and supporting parallel/distributed processing for science codes, and accommodating future COTS parts/standards through upgrades

• An application software development and runtime environment that is familiar to science application developers, and facilitates porting of applications from the laboratory to the spacecraft payload data processor

• An autonomous and adaptive controller for fault tolerance configuration, responsive to environment, application criticality, and system mode, that maintains required dependability and availability while optimizing resource utilization and system efficiency

• A methodology and tools which allow the prediction of the system's behavior in the space environment, including: predictions of availability, dependability, fault rates/types, and system level performance.

The basic architecture (Figure 6) consists of a redundant radiation-hardened system controller which acts as the controller for a parallel processing cluster of COTS-based, high-performance, data processing nodes, and a redundant network interconnect. The implementation employs a radiation-hardened system controller, a Honeywell 603e Power PC (PPC) based SBC (Single-Board Computer), the high-performance data processing nodes are PPC750FX compute nodes with FPGA (Field Programmable Gate Array) co-processor accelerators. The interconnection of the parallel processing cluster is via Gigabit Ethernet. The system can be augmented with mission-specific elements, including mass storage, custom interfaces, and radiation sensors, as required.

Figure 6: Illustration of the DM hardware architecture (image credit: Honeywell)
Figure 6: Illustration of the DM hardware architecture (image credit: Honeywell)
Figure 7: Hardware architecture of a data processor (image credit: Honeywell)
Figure 7: Hardware architecture of a data processor (image credit: Honeywell)

The specific mission objectives of the DM experiment validation are four-fold:

1) To expose a COTS-based, high-performance processing cluster to the real space radiation environment

2) To characterize the radiation environment

3) To correlate the radiation performance of the COTS components with the environment

4) To assess the radiation performance of the COTS components and the DM system response in order to validate the predictive reliability, availability, and performance models for the DM flight experiment and for future NASA missions.

The DM experiment is considered to be a free running experiment, collecting radiation environment characterization and radiation event (SEU) data, correlating the environment and detected events with S/C ephemeris, and monitoring and reporting Dependable Multiprocessor response. The DM is expected to be run continuously for four of the six month ST8 mission to maximize the amount of data collected.

The DM experiment encompasses the measurement of component and system parameters that can only be validated in a real space environment. Primarily, these are the component fault/error rates dues to radiation, and the accuracy of the predictive fault/error model. The primary data categories to be taken during the flight experiment are:

• Spacecraft ephemeris

• Fault rates, types, and error detection and recovery coverages

• SEU (Single Event Upset) alarm measurements

• Environment diagnostic sensor measurements

The spacecraft ephemeris will be used to correlate the radiation performance of the COTS components with the orbit location.

Figure 8: DM system software architecture (image credit: Honeywell)
Figure 8: DM system software architecture (image credit: Honeywell)
Figure 9: Configuration of the DM experiment (image credit: NASA, Honeywell)
Figure 9: Configuration of the DM experiment (image credit: NASA, Honeywell)

Todays COTS computer generation has become more resistant to the debilitating effects of radiation than previous generations. Many commercial parts can withstand many 10 s of krad of TID (Total Ionizing Dose) and are immune to catastrophic Single Event Latchup (SEL). The primary issue preventing the deployment of a COTS-based spaceborne cluster computer is their continued SEU (Single Event Upset) susceptibility, which cause only soft, transient errors, not permanent hardware failures. Further, the latest generation of computer electronics, using SOI-CMOS (Silicon on Insulator-Complementary Metal-Oxide Semiconductor) technology, has proven to be approximately an order of magnitude less susceptible to SEU than previous bulk CMOS.

If the DM technology allows a system to withstand a few errors per day per processor, without unduly impacting system dependability, it will be possible to fly, essentially commercial, cluster computers. Not only would this provide mission enabling performance and performance density levels, but it would significantly lower the cost of development, as standard laboratory science codes could be easily ported to these systems without the expensive and error prone process normally associated with moving complex codes from the lab to a new flight platform.

Figure 10: Artist's rendition of the ST8 spacecraft in orbit (image credit: OSC)
Figure 10: Artist's rendition of the ST8 spacecraft in orbit (image credit: OSC)


2) P. R. Turner, L. M. Herrell, "A Pragmatic Access to Space Approach for the ST8 Mission," 2005 IEEE Big Sky Conference, IEEEAC paper #1428, 17 December 2004

3) P. R. Turner, L. M. Herrell, "Formulation Refinement and Access to Space for the ST8 Mission," Proceedings of the 2006 IEEE/AIAA Aerospace Conference, Big Sky, MT, USA, March 4-11, 2006

4) S. Franklin, J. Ku, B. Spence, M. McEachen, S. White, J. Samson, R. Some, J. Zsoldos, "The Space Technology 8 Mission," Proceedings of the 2006 IEEE/AIAA Aerospace Conference, Big Sky, MT, USA, March 4-11, 2006

5) H. Abakians, A. Chmielewski, S. Franklin, "NASA's New Millennium Program, The Space Technology 8 (ST8) Mission," Proceedings of the Sixth Annual NASA Earth Science Technology Conference (ESTC 2006), College Park, MD, USA, June 27-29, 2006, URL:





10) B. Spence, S. White, N. Wilder, T. Gregory, M. Douglas, R. Takeda, N. Mardesich, T. Peterson, B. Hillard, P. Sharps, N. Fatemi, "Next Generation UltraFlex Solar Array for NASA's New Millennium Program Space Technology 8," Proceedings of IEEE Aerospace Conference 2005, Big Sky, MT, USA, March 5-12, 2005 URL:

11) M. E. McEachen, T. A. Trautt, D. M. Murphy, "The ST8 SAILMAST Validation Experiment," Proceedings of the 46th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics & Materials Conference, April 18-21, 2005, Austin, Texas, USA, URL:

12) J. Ku, L. Ottenstein, D. Douglas, M. Pauken, G. Birur, "Miniature Loop Heat Pipe with Multiple Evaporators for Thermal Control of Small Spacecraft," Government Microcircuit Applications and Critical Technology Conference, April 4-7, 2005, Las Vegas, NV, URL:

13) J. Ramos, J. Samson, D. Lupia, V. Aggarwal, M. Patel, I. Troxel, R. Subramaniyan, A. Jacobs, J. Greco, G. Cieslewski, J. Curreri, M. Fischer, E. Grobelny, A. George, R. Some, "High-Performance, Dependable Multiprocessor," Proceedings of the 2006 IEEE/AIAA Aerospace Conference, Big Sky, MT, USA, March 4-11, 2006, URL:

14) J. R. Samson Jr., J. Ramos, M. Patel, A. D. George, R. Some, "Technology Validation: NMP ST8 Dependable Multiprocessor Project," Proceedings of the 2006 IEEE/AIAA Aerospace Conference, Big Sky, MT, USA, March 4-11, 2006, URL:

This description was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" - comments and corrections to this article are welcomed by the author.

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