OTB Spacecraft ODAR

0087-EX-CN-2018 Text Documents

General Atomics

2018-02-01ELS_204219

ORBITAL DEBRIS ASSESSMENT REPORT
(ODAR) FOR THE ORBITAL TEST BED (OTB)
SATELLITE

This document supports the General Atomics Electromagnetic
Systems (GA-EMS) Orbital Test Bed Satellite FCC license and has been
prepared in accordance with:
NASA Technical Standard NASA-STD-8719.14A (inc. change 1, 25-5-
2012) Process for Limiting Orbital Debris
DAS Software version used in analysis V 2.1.1

Date: December 2017

Document No.: OTB-DOC-000638-05

General Atomics Electromagnetic Systems
345 Inverness Drive South, Suite 100
Englewood, CO 80112

Tel: 303.790.0653 | Fax: 303.792.2386

DISTRIBUTION STATEMENT: Confidential and proprietary information of GA-EMS. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

Table of Contents
1 Applicable Documents .............................................................................................. 1
2 Acronyms and Abbreviations ....................................................................................... 1
3 PREAMBLE ............................................................................................................. 2
4 ODAR SELF ASSESSMENT ............................................................................................ 3
5 PROGRAM MANAGEMENT AND MISSION OVERVIEW ............................................................. 4
6 SPACECRAFT DESCRIPTION ......................................................................................... 6
7 ASSESSMENT OF SPACECRAFT DEBRIS RELEASED DURING NORMAL OPERATIONS ........................ 10
8 ASSESSMENT OF SPACECRAFT INTENTIONAL BREAKUPS AND POTENTIAL FOR EXPLOSIONS ........... 11
9 ASSESSMENT OF SPACECRAFT POTENTIAL FOR ON-ORBIT COLLISIONS .................................... 14
10 ASSESSMENT OF SPACECRAFT POSTMISSION DISPOSAL PLANS AND PROCEDURES ....................... 16
11 ASSESSMENT OF SPACECRAFT REENTRY HAZARDS ............................................................ 22
12 ASSESSMENT OF SPACECRAFT HAZARDOUS MATERIALS ...................................................... 26

List of Figures
Figure 6-1: OTB Block Diagram.......................................................................................... 6
Figure 6-2: OTB Physical Configuration (units in mm) .............................................................. 7
Figure 6-3: Outline of OTB Satellite Configuration (solar panels removed) .................................... 8
Figure 9-1: OTB DAS output for Collisions with Large Objects .................................................. 14
Figure 10-1: Stowed Drag Augmentation Device ................................................................... 16
Figure 10-2: Stowed Drag Augmentation Device (detailed view) ............................................... 17
Figure 10-3: Drag Augmentation Device Activation Diagram .................................................... 17
Figure 10-4: Deployed State of Drag Augmentation Device ...................................................... 18
Figure 10-5: Drag Augmentation Redundancy Diagram ........................................................... 19
Figure 10-6: Orbital Apogee/Perigee Altitude ..................................................................... 20
Figure 11-1: CAD Representation of the Satellite (with callouts of objects used in DAS simulation) .... 23
Figure 11-2: Cutaway View of OTB (with callouts of objects used in DAS simulation) ...................... 24
Figure 11-3: DAS Output for Requirement 4.7-1 ................................................................... 25

List of Tables
Table 9-1: Critical Units/Surfaces .................................................................................... 15
Table 10-1: Drag Augmentation Device Technology TRL ......................................................... 18
Table 11-1: Modelled Components of the Spacecraft ............................................................. 22
Table 12-1: Hazardous materials data ............................................................................... 26

ii
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

3 PREAMBLE
The Orbital Debris Assessment Report (ODAR) has been prepared by GA-EMS for the Orbital Test Bed
(OTB) satellite. It considers in turn each of the requirements for limiting the generation of orbital
debris, as specified in the NASA Technical Standard. As such, this document has been prepared in
accordance with reference [1] using the following software tool:

 DAS software version 2.1.1 [2].

This document supports GA-EMS’ application for an experimental license [3]. Further information on
OTB’s RF system is provided in reference [4].

2
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

4 ODAR SELF ASSESSMENT
Orbital Debris Assessment Report Evaluation: Orbital Test Bed (OTB) Mission.

This provides a summary overview of the compliance against requirements. The details for each
requirement are contained within the body of this report. Note: GA-EMS is not responsible for the
launch vehicle ODAR.

Reqm’t
Launch Vehicle Spacecraft Comments
#

Complaint Not Incomplete Standard Compliant Not Incomplete Note: GA-EMS is not responsible for
Compliant Non- or N/A Compliant
Compliant the launch vehicle ODAR

4.3-1.a x Compliant – no debris released

4.3-1b x Compliant – no debris released

4.3-2 x Compliant – no debris released

4.4-1 x Compliant – no credible risk of
explosion

4.4-2 x Compliant – end of mission
passivation

4.4-3 x N/A – no planned explosions or
intentional collisions

4.4-4 x N/A – no planned explosions or
intentional collisions

4.5-1 x Compliant – probability of collision
0.00031

4.5-2 x Compliant – probability of collision
0.000506

4.6-1(a) x Compliant – deorbit 23.76 years from
EOM or 25.76 years from launch

4.6-1(b) x N/A – atmospheric re-entry disposal

4.6-1(c) x N/A – atmospheric re-entry disposal

4.6-2 x N/A – OTB will be in LEO

4.6-3 x N/A – OTB will be in LEO

4.6-4 x Compliant – significant spacecraft
heritage and EOM HDRM heritage

4.7-1 x Compliant – no objects expected to
survive re-entry, 1:100000000

4.8-1 x N/A – no tether on OTB

3
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

saturation of the reaction wheels. In addition, sun sensors, gyros and magnetometers also form part of
the Attitude & Orbit Control System as shown in the block diagram, Figure 6-1.

6.7 Range Safety or other Pyrotechnic Devices
There are no pyrotechnic devices within the OTB satellite.

6.8 Electrical Generation and Storage System
The OTB bus has four deployable solar panels and four body-mounted panels, of which two are hosted
payloads, see Figure 6-2. The power from the solar panels is used to charge a flight proven ABSL 8s10p
Lithium-Ion battery manufactured by ABSL (http://www.enersys.com/) which provides an unregulated
power supply to downstream subsystems. The system and battery have been designed in order to
complete the mission even if there is a string failure on the battery or the solar arrays.

Power is supplied from the battery to the Power Distribution Module (PDM) units. The PDM controls the
switching of the 28V unregulated battery supply to spacecraft subsystems.

6.9 Other Sources of Stored Energy
The deployable solar panels are deployed using the SSTL hold down and release system (HDRS, see
section 7.1). No debris is produced during solar panel deployment.

6.10 Radioactive Materials
There are no radioactive materials onboard the OTB satellite.

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All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

7 ASSESSMENT OF SPACECRAFT DEBRIS RELEASED DURING NORMAL
OPERATIONS
7.1 Objects Expected to be released from the Spacecraft any time after Launch
There are no intentional releases from OTB; all deployment systems are self-contained entities
designed not to release debris into the environment.

The deployment systems are:

1) SSTL Hold Down and Release System

The Hold Down and Release System (HDRS) for the deployable solar panels utilizes captive devices on
both sides of the interface (retention of bolt and springs) and as such does not generate debris. This
system has been used successfully on previous SSTL spacecraft and as such is TRL9 Flight Proven.

2) Planetary Systems Corporation Separation System (‘Lightband’)

Planetary Systems Corporation states within their 2000785 Rev C User’s Manual for Mark II Lightband,
2nd April 2013 (section 3, point 5): Non-pyrotechnic. The Lightband generates no debris on or after
separation’ and is a TRL 9 system. http://www.planetarysystemscorp.com/#!__products/mark-ii-
motorized-lightband.

3) Drag Augmentation Device

The drag augmentation device is based on flight proven technologies and components. This system is
further explained in section 10.

7.2 Spacecraft Compliance with Requirements 4.3-1 and 4.3-2
Requirement 4.3-1: Debris passing through LEO – released debris with diameters of 1mm or larger:

a) Requirement 4.3-1a: All debris released during the deployment, operation, and disposal phases
shall be limited to a maximum orbital lifetime of 25 years from date of release (Requirement
56398).

Compliant ― no debris released during normal operations.

b) Requirement 4.3-1b: The total object-time product shall be no larger than 100 object-years per
mission (Requirement 56399). The object-time product is the sum of all debris of the total time
spent below 2,000 km altitude during the orbital lifetime of each object. (See section 4.3.4.2
for methods to calculate the object-time product.)

Compliant ― no debris released during normal operations.

Requirement 4.3-2: Debris passing near GEO: For missions leaving debris in orbits with the potential of
traversing GEO (GEO altitude +/- 200 km and +/- 15 degrees latitude), released debris with diameters
of 5 cm or greater shall be left in orbits which will ensure that within 25 years after release the apogee
will no longer exceed GEO - 200 km (Requirement 56400). DAS will be used to assess compliance
against these requirements.

Compliant ― no debris released during normal operations.

10
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

8 ASSESSMENT OF SPACECRAFT INTENTIONAL BREAKUPS AND
POTENTIAL FOR EXPLOSIONS
8.1 Potential Causes of Spacecraft Breakup during Deployment and Mission
Operations
The spacecraft is a heritage design and has undergone rigorous structural acceptance testing during
qualification. There are no pressurized systems on board, and no propulsion system.

OTB will fly the flight proven ABSL 8s10p Lithium-ion battery that employs the commercial SONY
18650HC cell. The contents of this cell are held at 1 bar with all cells having a pressure release device
that provides a safe rupture to vent ratio in accordance with recommended NASA safety standards.

No credible potential causes of spacecraft breakup during deployment or mission operations are
foreseen.

8.2 Summary of Failure Modes and Effects Analyses of all Credible Failure Modes
which may lead to an Accidental Explosion
The only device that stores a sufficient amount of energy to cause an accidental explosion is the flight
proven, heritage ABSL 8s10p battery. This unit has flown on over 50 spacecraft to date and the SONY
18650HC cell that it employs has accrued more than 50,000 cell years in orbit without any safety
events in space. Before launch, the battery is charged and then disconnected from the spacecraft. It is
only connected to the spacecraft power system upon separation from the launch vehicle so it presents
no credible risk during the launch phase of the mission.

Upon separation and during operations, the battery will be cycled at a benign depth of discharge of
less than 15% with an end of charge voltage of 32.8V. The power system is a heritage design and has
battery charge and discharge safety functionality that has been proven on numerous other SSTL
spacecraft including CFESat, RapidEye, UK-DMC-2, Deimos and exactView-1.

There are three possible failure modes for the battery:

1) Overcharge of Lithium-ion batteries above the recommended maximum end of charge voltage
of 33.6V could result in thermal runaway. To prevent this, the battery voltage is constantly
monitored by the power subsystem and charging is halted if this voltage is exceeded. This
functionality has been tested and verified using the OTB power subsystem. As a second level of
protection, in the event of the failure of this safety circuitry, the ABSL 8s10p cells contain
internal protection devices that result in an open circuit shutting down the entire battery. This
functionality has been demonstrated via a number of ground tests.

2) Short circuit of Lithium-ion batteries can result in very high current draw that leads to heating
and could invoke thermal runaway. To prevent this, the spacecraft has fused or switched power
lines to all components so that high current draw effects are halted before the battery can be
affected. As an additional level of protection, in the event of fuse failure, the ABSL 8s10p cells
carry short circuit protection devices that detect a rise in temperature and increase the
internal resistance reducing current flow mitigating a rise in temperature.

3) Over discharge of the ABSL 8s10p, where the battery is taken below the minimum
recommended operating voltage of 20V, has been proven by test to degrade battery
performance but not pose a safety risk. Over-discharge tests where ABSL 8s10p cells have been
driven to negative voltage have been shown to result in a safe failure with cells effectively
becoming resistors. In addition, the power system of the satellite will switch off below 23V to
limit the discharge of the battery below this point.

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All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

The SONY 18650HC cells in the ABSL 8s10p battery have had thermal runaway testing performed, which
showed that in the event of a single cell thermal runaway with other cells in healthy conditions
propagation does not occur. As such, even thermal runaway of a single cell (which is protected against
through the mechanisms described above) has not been shown to cause an explosion.

In summary, there are no credible failure modes which would cause an accidental explosion and the
battery design has been sufficiently tested to insure that thermal runaway propagation does not occur.

8.3 Detailed Plan for any Designed Spacecraft Breakup
There are no intentional breakups ― explosions nor intentional collisions ― planned.

8.4 List of Components which are Passivated at EOM
At End of Mission (EOM), all units except the power system and the RF receivers will be turned off. The
battery will continue to charge and discharge at a very low depth of discharge to the nominal end of
charge level set via telecommand. The design of the heritage power system is such that the battery
cannot be disconnected from the power system at EOM.

Reaction wheels are stopped in a controlled manner as part of EOM operations and switched off.

An end of charge voltage of 31V will be selected (the minimum possible) that will result in the battery
being at a state of charge < 50% and minimize the energy available for any failure event. Ground and
in-orbit life test data indicate an expected graceful degradation of the battery with effective battery
capacity reducing over time, leading to a benign state through the absence of useful capacity. With the
end of charge voltage set to 31 V any additional charging capacity will be dissipated from the radiator
surface.

8.5 Rationale for Items which cannot be Passivated due to their Design
Please see response to section 8.4 on the battery. The battery cannot be completely passivated
(discharged to 0% state of charge and disconnected from the power system) due to heritage system
design and power system which have successfully flown on numerous SSTL missions.

8.6 Spacecraft Compliance with Requirements 4.4-1 through 4.4-4
Requirement 4.4-1: Limiting the risk to other space systems from accidental explosions during
deployment and mission operations while in orbit about Earth or the Moon: For each spacecraft and
launch vehicle orbital stage employed for a mission, the program or project shall demonstrate, via
failure mode and effects analyses or equivalent analyses, that the integrated probability of explosion
for all credible failure modes of each spacecraft and launch vehicle is less than 0.001 (excluding small
particle impacts) (Requirement 56449).

Compliant ― GA-EMS see no credible risk of spacecraft explosion (caused by the spacecraft itself
without external impact) during launch, nominal operations and the time between EOM and Earth re-
entry (Refer to section 8.1).

Requirement 4.4-2: Design for passivation after completion of mission operations while in orbit about
Earth or the Moon: Design of all spacecraft and launch vehicle orbital stages shall include the ability
and a plan to deplete all on-board sources of stored energy and disconnect all energy generation
sources when they are no longer required for mission operations or post mission disposal or control to a
level which cannot cause an explosion or deflagration large enough to release orbital debris or break up
the spacecraft (Requirement 56450)

Compliant.

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purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

Requirement 4.4-3: Limiting the long-term risk to other space systems from planned breakups: Planned
explosions or intentional collisions shall:

a) Be conducted at an altitude such that for orbital debris fragments larger than 10 cm the
object-time product does not exceed 100 object-years (Requirement 56453). For example, if
the debris fragments greater than 10cm decay in the maximum allowed 1 year, a maximum of
100 such fragments can be generated by the breakup.

b) Not generate debris larger than 1 mm that remains in Earth orbit longer than one year
(Requirement 56454).

Not applicable ― OTB will not undertake any planned explosions or intentional collisions.

Requirement 4.4-4: Limiting the short-term risk to other space systems from planned breakups:
Immediately before a planned explosion or intentional collision, the probability of debris, orbital or
ballistic, larger than 1 mm colliding with any operating spacecraft within 24 hours of the breakup shall
be verified to not exceed 10-6 (Requirement 56455).

Not applicable ― OTB will not undertake any planned explosions or intentional collisions.

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All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

pointing until the deorbit device is deployed. The critical surface of the units will therefore be in the
plane of the direction of motion.

During the EOM phase, the flight computer will receive the deorbit command from the ground through
the S-band patch antennas and the receiver. The onboard computer will then command the deorbit
device to deploy, which will begin to deorbit the spacecraft. For these actions to occur, the power
system must also be active.

An extra level of protection to the critical units is provided by our standard dual-redundant design
approach. Each of the units listed in Table 9-1, aside from the deorbit device, are redundant. In the
unlikely event of a unit being damaged during the operational phase, the deorbit device could still be
deployed by the redundant components. See section 10 for further discussion.

Table 9-1: Critical Units/Surfaces

Critical Surface Probability of Penetration
Drag Augmentation HDRM 0.000000
OBC 0.000067
PDM 0.000014
Battery 0.000010
RF Tray 0.000012
Antennas 0.000405
TOTAL PROBABILITY OF PENETRATION 0.000508

The total probability of post-mission disposal failure is 0.000508, which is compliant with the
requirement.

9.3 Spacecraft Compliance with Requirements 4.5-1 and 4.5-2
Requirement 4.5-1: Limiting debris generated by collisions with large objects when operating in Earth
orbit: For each spacecraft and launch vehicle orbital stage in or passing through LEO, the program or
project shall demonstrate that, during the orbital lifetime of each spacecraft and orbital stage, the
probability of accidental collision with space objects larger than 10 cm in diameter is less than 0.001
(Requirement 56506).

Compliant ― the probability of collision with large objects was calculated by DAS [2] to be 0.00031
during the orbital lifetime, which is compliant with the requirement, Figure 9-1.

Requirement 4.5-2: Limiting debris generated by collisions with small objects when operating in Earth
or lunar orbit: For each spacecraft, the program or project shall demonstrate that, during the mission
of the spacecraft, the probability of accidental collision with orbital debris and meteoroids sufficient
to prevent compliance with the applicable post mission disposal requirements is less than 0.01
(Requirement 56507).

Compliant ― the probability of accidental collision with orbital debris and meteoroids sufficient to
interfere with post-mission disposal was calculated by DAS [2] to be 0.000508, which is compliant with
the requirement, Table 9-1.

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All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

Activation redundancy is achieved through the spacecraft platform design, including the electrical
power system and switch architecture. The deployment of the drag augmentation device is controlled
by the power system through a system of series switches. This allows for Arm and Fire commanded
switches to ensure that the deployment only takes place when commanded at EOM. These Arm and Fire
switches are also configured so that each pair has a redundant circuit in the event of a failure of one of
the sets. The commanding of these switches also contains a built in redundancy in that the commands
come directly from the onboard computer (OBC), but in the event of a critical system failure with the
OBC the switches can still be commanded simply through the RF system via emergency, low-level
commanding.

The redundancy inherent in the platform system design for the subsystems required to activate the
drag augmentation device is illustrated in Figure 10-5. There are no single point failures present in the
platform elements required for activation, with at least one-layer of redundancy. This is our standard
design approach, used on numerous missions, which have demonstrated 100% mission success over the
last 15 years.

Primary Drag
Primary On-Board Primary TiNi Heater
Primary RF Receiver Augmentation
Computer Element
Device Switches
Power System –
Mission Operation Power Disribution De-Orbit Device
Center Commanding (Internally Frangibolt TiNi
Redundant)
Redundant Drag
Redundant RF Redundant On- Secondary TiNi
Augmentation
Receiver Board Computer Heater Element
Device Switches

Figure 10-5: Drag Augmentation Redundancy Diagram

Release redundancy is inherit to the TiNi Frangibolts through the use of primary and secondary heater
circuits. In the event that the primary heating element is damaged or is unable to release the
Frangibolt the secondary heating circuit can be used. This ensures that the Frangibolt successfully
deploys.

The drag augmentation system has also been designed to have inherit performance redundancy by
being designed to deorbit the system well within the required timeline, as described below. This
additional margin provides further confidence for uncertainty in atmospheric models.

Figure 10-6 shows the lifetime plot for the OTB satellite after deployment of the drag augmentation
device at end of mission. Assuming a launch date of 2018, with a 2 year mission lifetime, OTB will
deorbit within 25.76 years from launch (less than the required 30 years for total on-orbit time). Note
that the calculation assumes random tumble of the OTB satellite after deployment.

19
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

OTB-DOC-000638-05 ODAR FOR THE ORBITAL TEST BED (OTB) SATELLITE December 2017

 Leave the space structure in an orbit in which natural forces will lead to atmospheric
reentry within 25 years after the completion of mission but no more than 30 years after
launch; or
 Maneuver the space structure into a controlled deorbit trajectory as soon as practical after
completion of mission.

b) Storage orbit option: Maneuver the space structure into an orbit with perigee altitude greater
than 2000 km and apogee less than GEO - 500 km.

c) Direct retrieval: Retrieve the space structure and remove it from orbit within 10 years after
completion of mission.

The OTB satellite reentry is compliant using reentry option “a”.

Requirements 4-6-1(b) and 4-6-1(c) are not applicable.

Requirement 4.6-2: Disposal for space structures near GEO: A spacecraft or orbital stage in an orbit
near GEO shall be maneuvered at EOM to a disposal orbit above GEO with a predicted minimum perigee
of GEO +200 km (35,986 km) or below GEO with an apogee of GEO – 200 km (35,586 km) for a period of
at least 100 years after disposal (Requirement 56563).

Not applicable. OTB will be in LEO.

Requirement 4.6-3: Disposal for space structures between LEO and GEO:

a. A spacecraft or orbital stage shall be left in an orbit with a perigee greater than 2000 km above the
Earth’s surface and apogee less than 500 km below GEO (Requirement 56565).

b. A spacecraft or orbital stage shall not use nearly circular disposal orbits near regions of high value
operational space structures, such as between 19,200 km and 20,700 km (Requirement 56566).

Not applicable. OTB will be in LEO.

Requirement 4.6-4: Reliability of post mission disposal operations in Earth orbit: NASA space programs
and projects shall ensure that all post mission disposal operations to meet Requirements 4.6-1, 4.6-2,
and/or 4.6-3 are designed for a probability of success as follows: (Requirement 56567)

a) Be no less than 0.90 at EOM.

b) For controlled reentry, the probability of success at the time of reentry burn must be
sufficiently high so as not to cause a violation of Requirement 4.7-1 pertaining to limiting the
risk of human casualty.

SSTL has a long heritage of successful missions, 49 spacecraft launched to date with over 500 years
collectively on orbit, in part due to the multiple points of redundancy built into every mission design
with a design life of 5 years or more. The drag augmentation device, Frangibolt based, has a minimum
reliability of 0.99999 ― this is the supplier published reliability for the TiNi Frangibolt and is the
industry standard with 20 years of space heritage. The OTB satellite is an “SSTL-150 ESPA” which is
listed in the NASA Rapid III catalogue; it has a Ps of 92%. (Noting that SSTL have not had a mission
failure on-orbit in over 15 years, and none with this class of satellite.)This gives us confidence that
OTB will be able to complete its 2 year mission and successfully deploy the drag augmentation devices
at EOM, with a Ps = 0.91999.

4.6.4b N/A ― OTB does not have a controlled re-entry.

21
All information contained in this document is proprietary and confidential to GA-EMS and shall not, in whole or in part, be reproduced or disclosed, or used for any
purpose other than for which it is provided, without GA-EMS prior written consent. All rights reserved.

Document Created: 2018-01-30 16:56:38
Document Modified: 2018-01-30 16:56:38

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