Orbital Debris Assessment Report and Mitigation Plan

1228-EX-ST-2019 Text Documents

Spire Global, Inc.

2019-07-03ELS_233315

2017

Spire Global, Inc.
Orbital Debris Assessment Report

LEMUR-2 Satellites
Phase IC (with Hosted Payloads) and
Phase II

Revision History

Revision Description of Revisions Release Date

1 Initial Release 10/13/2017

An orbital debris risk assessment of LEMUR-2 Phase IC (with hosted
payload) and Phase II satellites

Section 1: Program Management and Mission Overview

Program / Project Jenny Barna
Manager

Mission The purpose of the LEMUR-2 satellite constellation is to provide high-revisit global maritime and
Description aircraft domain monitoring data, weather data, and hosted payload services.

This orbital debris risk assessment report (“ODAR”) covers any LEMUR-2 Phase IC satellites with a
hosted payload and Phase II satellites proposed to be launched by Spire Global, Inc. (“Spire”).

Foreign None
Government
Involvement

Project Milestones LEMUR-2 satellites are usually launched in small deployments depending on available capacity,
quality of orbit, service and constellation replenishment needs, and risk profiles of the launch
Proposed vehicle and campaign.
Launch Date
Given the potential long lead time for the instant application and state of the low-Earth orbit
Proposed launch market for secondary payloads, Spire is filing this application early and is not capable of
Launch Vehicles providing launch parameters for the Phase II satellites at this time. However, it notes that these
satellites (similar to the Phase I, IB, and IC satellites) will only deploy at orbital altitudes from 385
Proposed
to 650 km and inclinations ranging from equatorial to polar sun-synchronous (98 degrees).
Launch Sites
This analysis considers the range of representative orbits and includes a debris assessment of the
Launch Vehicle
worst-case altitude and lifetime in order to provide the most conservative results. Spire is also
Operator
seeking authority to deploy from and above the International Space Station (“ISS”), so that orbit is
also considered.

Mission Duration The planned operational lifetime of each LEMUR-2 satellite is 2 years following deployment from
the launch vehicle.

Selection of Orbit Orbits are selected based on availability of launches, an established range of acceptable
deployment altitudes (385 km – 650 km), and inclinations (equatorial to polar sun-synchronous
(98 degrees)) that support the operational purpose of the constellation.

Potential Physical The LEMUR-2 Phase IC satellites with a hosted payload and Phase II satellites do not have any
Interference with propulsion systems to actively maintain orbital altitude. Therefore, their orbit will naturally decay
Other Orbiting following deployment from either the launch vehicle or the ISS.
Objects
As detailed in Section 5, the probability of physical interference between the LEMUR-2 satellites
and other space objects complies with Requirement 4.5 of NASA-STD-8719.14A.

Phase IC (with Spire will add a third solar “drag” panel on any LEMUR-2 Phase IC satellite bus with a hosted
hosted payload) payload and Phase II satellite bus, increasing the amount of drag on its satellite and shortening
and Phase II the orbital lifetimes by between 0.50 and 0.75 years (dependent on solar cycle changes) from

1

Satellite Bus launch at its highest orbit of 650 km. See infra § 6.
Configuration
Notes The LEMUR-2 Phase IC satellite and Phase II satellite will have a nominal launch mass
configuration of 4.5kg; however, the mass capacity may be up to 6kg maximum, which
accommodates potential other Spire or hosted payload(s). Surface area and spacecraft
specifications are otherwise identical. Both nominal and maximum cases are included in this
ODAR for collision risk and lifetime analyses. See infra §§ 5-6.

2

ODAR Section 2: Spacecraft Description
Physical Description:

Property Value

Total Mass at Launch 4.5 kg nominal; 6 kg maximum

Dry Mass at Launch 4.5 kg nominal; 6 kg maximum (no propellant/propulsion system)

Form Factor 3U cubesat

COG <3 cm radius from geometric center

Envelope (stowed) 100 mm x 100 mm x 340.5 mm (excluding dynamic envelope)

Envelope (deployed) 1 m x 1 m x 300 mm

Propulsion Systems None

Fluid Systems None

AOCS Stabilization/pointing with 3x orthogonal reaction wheels, desaturation + coarse
pointing with magnetorquers, and Global Positioning System (“GPS”) navigation

Range Safety / None
Pyrotechnic Devices

Electrical Generation Triple-junction GaAs solar panels

Electrical Storage Rechargeable Lithium-Ion battery pack

Radioactive Materials None

3

ODAR Section 3: Assessment of Debris Released During
Normal Operations

Spire’s LEMUR-2 Phase IC satellites with a hosted payload and Phase II satellites do not release objects during
deployment or operation. Therefore, Requirements 4.3-1 and 4.3-2 of NASA-STD-8719.14A are not applicable.

4

ODAR Section 4: Assessment of Spacecraft Intentional
Breakups and Potential for Explosions
Potential causes for spacecraft breakup:

LEMUR-2 Phase IC satellites with a hosted payload and Phase II satellites have no propulsion and accordingly do
not carry highly volatile propellant. The only energy sources (kinetic, chemical, or otherwise) onboard the
spacecraft are a Lithium-Ion battery system and reaction wheels. Thus, the only two plausible causes for breakup
of these LEMUR-2 satellites are the following:
1. energy released from onboard batteries and
2. mechanical failure of the reaction wheels.

Summary of failure modes and effects analysis of all credible failure modes, which may lead to an
accidental explosion:

The battery aboard these LEMUR-2 satellites is an 80Wh Lithium-Ion battery pack, which represents the only
credible failure mode during which stored energy is released. The main failure modes associated with Lithium Ion
batteries result from overcharging, over-discharging, internal shorts, and external shorts.

The only failure mode of the reaction wheel assemblies that could lead to creation of debris would be breakup of
the wheels themselves due to mechanical failure while operating at a high angular rate.

Risk mitigation plan:

The battery pack onboard these LEMUR-2 satellites has been designed and qualified to comply with controls /
process requirements identified in NASA Report JSC-20793 ‘Crewed Space Vehicle Battery Safety Requirements’ to
mitigate the chance of any accidental venting / explosion caused by the above failure modes.

The reaction wheels on board these LEMUR-2 satellites are limited with respect to maximum rotational speed of
the wheels and are contained within a sealed compartment, thus mitigating any risk of breakup of the wheels
themselves into debris.

Detailed plan for any designed spacecraft breakup, including explosions and intentional collisions:

There is no planned breakup of the satellites on-orbit.

Rationale for all items required to be passivated that cannot be due to design:

N/A

Assessment of spacecraft compliance with Requirements 4.4-1 through 4.4-4:

4.4-1, Limiting the risk to other space systems from accidental explosions during COMPLIANT
deployment and mission operations while in orbit about Earth or the Moon:

4.4-2, Design for passivation after completion of mission operations while in orbit N/A
about Earth or the Moon:

5

4.4-3, Limiting the long-term risk to other space systems from planned breakups: N/A

There are no planned breakups of any of the satellites.

4.4-4, Limiting the short-term risk to other space systems from planned breakups: N/A

There are no planned breakups of any of the satellites.

6

ODAR Section 5: Assessment of Spacecraft Potential for On-
Orbit Collisions
Probability for collision with objects larger than 10 cm:

The probability of a collision of any of any LEMUR-2 Phase IC satellites with a hosted payload or Phase II satellites
with an orbiting object larger than 10 cm in diameter was calculated using the National Aeronautics and Space
Administration’s (“NASA’s”) Debris Assessment Software (“DAS”) 2.0.2 software. Table 1 below shows the risk for
all orbits into which LEMUR-2 satellites may be deployed in each of five different area/mass ratio scenarios,
including a worst-case scenario. The table shows the risk both at the expected nominal orbital dwell time and at
the worst-case dead-on-arrival orbital dwell time. ISS deployments are from in front and below the ISS, typically
in a range of 385 km to 400 km, at the time of deployment as directed by the ISS Program. Table 2 below shows a
worst-case analysis of 400 km. Certain deployments have similar inclinations but slightly different altitudes.
Where the altitude is slightly different, Spire groups the launches together under the worst-case (highest) altitude.

7

Table 1 –LEMUR-2 Phase IC (with hosted payload) or Phase II Satellites
Collision Risk with Objects Larger Than 10 cm (Run at Worst-Case Orbit of 650 km, 98 deg)

Nominal Mass Configuration (4.5 kg) Maximum Mass Configuration (6 kg)

650 km, 98 degrees (Worst-Case Orbit) 650 km, 98 degrees (Worst-Case Orbit)

Collision Risk
Effective Area/Mass Orbital Dwell Time Effective Area/Mass Orbital Dwell Time Collision Risk
Satellite Operational State per NASA DAS
(m2/kg) (years): (m2/kg) (years): per NASA DAS Analysis
Analysis

Satellite Nonfunctional 0.0082 19.7 3 x 10-6 0.0061 21.8 1 x 10-5

ADCS Nonfunctional,
0.0140 18.01 4 x 10-6 0.0105 21.8 1 x 10-5
Partial Deploy

ADCS Nonfunctional, Fully
0.0201 14.7 4 x 10-6 0.0151 16.8 1 x 10-5
Deployed

Operational, 14.9 4 x 10-6 0.0146
0.0194 17.2 1 x 10-5
Partial Deploy

Operational, 7.5 4 x 10-6 0.0217 14.2 1 x 10-5
0.0290
Nominal

8

Table 2 –LEMUR-2 Phase IC (with hosted payload) or Phase II Satellites
Collision Risk with Objects Larger Than 10 cm (Run at ISS Orbit of 400 km, 51.6 deg)

Nominal Mass Configuration (4.5 kg) Maximum Mass Configuration (6 kg)

400 km, 51.6 degrees (ISS Orbit) 400 km, 51.6 degrees (ISS Orbit)

Collision Risk
Effective Area/Mass Orbital Dwell Time Effective Area/Mass Orbital Dwell Time Collision Risk
Satellite Operational State per NASA DAS
(m2/kg) (years): (m2/kg) (years): per NASA DAS Analysis
Analysis

Satellite Nonfunctional 0.0082 2.5 0 0.0061 2.7 0

ADCS Nonfunctional,
0.0140 2.0 0 0.0105 2.3 0
Partial Deploy

ADCS Nonfunctional, Fully
0.0201 1.6 0 0.0151 1.97 0
Deployed

Operational, 1.6 0 0.0146
0.0194 2.0 0
Partial Deploy

Operational, 1.0 0 0.0217 1.4 0
0.0290
Nominal

9

Probability for collision with objects 10 cm or less:

NASA’s DAS returned a response of Compliant with Requirement 4.5-2 of NASA-STD-8719.14A in a number of
potential orbits and configurations, including a worst-case scenario of 650 km, 98 degrees.

Assessment of spacecraft compliance with Requirement 4.5-1 and 4.5-2:

4.5-1, Probability of collision with large objects: COMPLIANT

4.5-2, Probability of damage from small objects: COMPLIANT

10

ODAR Section 6: Assessment of Spacecraft Post-Mission
Disposal Plans and Procedures
Description of disposal option selected:

Following its deployment, a LEMUR-2 Phase IC satellite with a hosted payload and Phase II satellite will naturally
decay until it reenters the atmosphere. Table 3 describes the mission scenarios for which lifetime analysis of these
LEMUR-2 satellites was considered and the effective area-to-mass ratio of the satellite in each scenario. The ratio
was calculated using the external dimensions of the LEMUR-2 satellite and deployed arrays. Note that Spire will
add a third solar “drag” panel on any Phase IC satellite bus with a hosted payload and Phase II satellite bus,
increasing the amount of drag on its satellite and shortening the orbital lifetimes (compared to its Phase I
satellites).1

For purposes of Section 6, drag area from deployed antennas was omitted; as such, the effective area-to-mass
calculated below is a conservative case.

Table 3 - Area-to-Mass Ratio of LEMUR-2 Phase IC (with hosted payload) and Phase II Satellites
in Various Mission Scenarios

Effective Area/Mass Ratio
Scenario Description (m2/kg)

Nominal Mass 4.5 kg Maximum Mass 6 kg2

Operational, Nominal ▪ Spacecraft pointing, position is
nominal, operational 0.0290 0.0217
▪ Solar arrays deployed

Operational, ▪ Spacecraft pointing, position is
Partial Deploy Failure nominal, operational 0.0194 0.0146
▪ 1 of 2 solar arrays deployed

ADCS Nonfunctional ▪ Spacecraft tumbling randomly
0.0201 0.0151
Fully Deployed ▪ Both solar panels deployed

ADCS Nonfunctional ▪ Spacecraft tumbling randomly
0.0140 0.0105
Partially Deployed ▪ 1 of 2 solar panel deployed

1
See Application of Spire Global, Inc., File No. SAT-AMD-20161114-00107, Orbital Debris Assessment Report: 100 LEMUR-2
Phase IB and IC Satellites, Exhibit C (filed Nov. 14, 2016).
2
As mentioned, LEMUR 2 Phase IC satellites and Phase II satellites will have capacity to add up to 1.5kg of total mass in the
accommodation of new Spire or hosted payload(s). This orbital debris assessment evaluates lifetime and collision risk with
both the nominal and maximum possible mass configurations. Surface area and spacecraft specifications are otherwise
identical.

11

Satellite ▪ Spacecraft tumbling randomly
Nonfunctional ▪ No solar panels deployed 0.0082 for 5 years 0.0061 for 5 years
0.0201 thereafter3 0.0151 thereafter3

Table 4 below shows the simulated orbital dwell time for a LEMUR-2 Phase IC satellite with a hosted payload and
Phase II satellite in a number of potential orbits and configurations, including a worst-case scenario of 650 km, 98
degrees.

3
This calculation conservatively assumes that the solar panels do not deploy in the first 5 years and that deployment only
occurs after nylon burn wire degrades in natural sunlight (i.e., double-fault situation). See infra note *.

12

Table 4 – Orbit Dwell Time for Phase IC (with hosted payload) and LEMUR-2 Phase II Satellites in Representative Low-Earth Orbits4

Nominal Mass (4.5 kg) Maximum Mass (6 kg)

Effective Effective
Spacecraft 400 km, 450 km, 500 km, 600 km, 650 km, 400 km, 450 km, 500 km, 600 km, 650 km,
Area/Mass Area/Mass
Operational State 51.6 deg 98 deg 98 deg 98 deg 98 deg 51.6 deg 98 deg 98 deg 98 deg 98 deg
(m2/kg) (m2/kg)

Satellite
0.0082 2.5 3.3 4.4 10.7* 19.7* 0.0061 2.7 3.6 5 12.2* 21.8*
Nonfunctional

ADCS Nonfunctional,
Partial Deploy 0.0140 2 2.8 3.6 8.14* 18.01* 0.0105 2.3 3 4 12.2* 21.8*

ADCS Nonfunctional,
0.0201 1.6 2.6 3.2 5.7 14.7 0.0151 1.97 2.8 3.5 7.2 16.8
Fully Deployed

Operational, 1.6 2.6 3.2 5.8 14.9 0.0146 2 2.8 3.5 7.5 17.2
0.0194
Partial Deploy

Operational, Nominal 0.0290 1 2.4 2.9 4.8 7.5 0.0217 1.4 2.5 3.1 5.5 14.2

* To ensure Spire exceeds the NASA standard in all scenarios, Spire has included a double fault-tolerant solar panel deployment mechanism, which will provide sufficient surface area and drag to comply with the NASA
standard even if the LEMUR-2 Phase IC satellites with a hosted payload and Phase II satellites are dead on arrival. These LEMUR-2 satellite’s solar panels are part of a built-in, post-deployment sequence programmed
into onboard software prior to launch, which requires no direction from the ground. If for some reason the onboard sequence fails, solar array deployment can be commanded from the ground. If a LEMUR-2 satellite is
non-communicative, an entirely passive, redundant fail-safe is included on all LEMUR-2 satellites in the form of a burn wire. The tensile strength of the burn wire has been tested and verified to degrade to a breaking
point after 3600 hours or 150 days of UV radiation exposure. Spire’s worst-case scenario for dwell time above conservatively models 5 years of non-deployed solar panels and no loss of altitude during those 5 years,
followed by the dwell times for an Attitude Determination and Control System (“ADCS”) nonfunctional satellite, even though a non-deployed solar panel LEMUR-2 would still have some surface area that would cause
some loss of altitude during that period. As such, the scenario is a conservative worst-case one.

13

Identification of systems required for post-mission disposal: None

Plan for spacecraft maneuvers required for post-mission disposal: N/A

Calculation of final area-to-mass Ratio if atmospheric reentry not selected: N/A

Assessment of Spacecraft Compliance with Requirements 4.6-1 through 4.6-4:

4.6-1, Disposal for space structures passing through low-Earth orbit (“LEO”): COMPLIANT

All satellites will reenter the atmosphere within 25 years of launch

4.6-2, Disposal for space structures passing through geostationary orbit (“GEO”): N/A

4.6-3, Disposal for space structures between LEO and GEO: N/A

4.6-4, Reliability of post-mission disposal operations: N/A

14

ODAR Section 7: Assessment of Spacecraft Reentry Hazards
NASA DAS was used to test the major spacecraft components for re-entry hazards. The major components
tested included the following.

• Solar panels and cells
• GPS antennas
• PCB circuit boards
• Primary structure
• Reaction wheel assembly

Summary of objects expected to survive an uncontrolled reentry (using DAS 2.0.2 software): None

Calculation of probability of human casualty for expected reentry year and inclination: 0%

Assessment of spacecraft compliance with Requirement 4.7-1:

4.7-1, Casualty risk from reentry debris: COMPLIANT

ODAR Section 7A: Assessment of Spacecraft Hazardous
Materials
Summary of hazardous materials contained on spacecraft: None

15

ODAR Section 8: Assessment for Tether Missions
Type of tether: N/A

Description of tether system: N/A

Determination of minimum size of object that will cause the tether to be severed: N/A

Tether mission plan, including duration and post-mission disposal: N/A

Probability of tether colliding with large space objects: N/A

Probability of tether being severed during mission or after post-mission disposal: N/A

Maximum orbital lifetime of a severed tether fragment: N/A

Assessment of compliance with Requirement 4.8-1:

4.8-1, Collision hazards of space tethers: N/A

16

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

Spire Global, Inc. (“Spire”) Orbital Debris Risk Mitigation Plan

Spire believes that (i) LEMUR-2 Phase IC satellites with a potential hosted payload and

Phase II satellites create relatively little additional orbital debris risks compared to existing

systems approved by the Federal Communications Commission (“Commission”) and (ii) these

satellites meet applicable orbital debris requirements as listed in Section 25.114(d)(14) of the

Commission’s rules.1

Any Phase IC satellite bus with a hosted payload and Phase II satellite bus may have

slightly different surface area and mass values from those values associated with the Phase I

satellites bus.

• Spire will add a third solar “drag” panel on all of its Phase IC satellites with a

hosted payload and Phase II satellites, increasing the amount of drag on the

satellites and shortening the orbital lifetimes by between 0.50 and 0.75 years

(dependent on solar cycle changes) from launch at their highest orbit of 650 km.

• The Phase IC satellite bus and Phase II satellite bus will have a nominal launch

mass configuration of 4.5 kg; however, the mass capacity may be up to 6 kg

maximum, which accommodates potential other Spire or hosted payload(s).

Surface area and spacecraft specifications are otherwise identical. Both nominal

and maximum cases are included in the ODAR for collision risk and lifetime

analyses.2

1
See 47 C.F.R § 25.114(d)(14); see also Orbital Debris Assessment Report: LEMUR-2 Satellites Phase IC (with
Hosted Payloads) and Phase II, Exhibit C (“Exhibit C”).
2
See Exhibit C.

1

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

Each section below addresses specific measures taken by Spire, as required under Section

25.114(d)(14), to limit the possibility that its space station operations will generate orbital debris.

Like the Phase I satellites that preceded them, Phase IC satellites with a potential hosted

payload and Phase II satellites continue to be one of the lowest risk satellite busses ever

approved by the Commission.

I. Orbital Dwell and Post-Mission Disposal

The Commission’s rules call for indication of the anticipated evolution over time of the

satellites’ orbits.3 Specifically, using the National Aeronautics and Space Administration

(“NASA”) Debris Assessment Software (“DAS”), Spire has calculated the dwell times of the

Phase IC satellites with a hosted payload and Phase II satellites.

At the highest orbit sought of 650 km, total orbital lifetime would not exceed 21.8 years

from launch in a conservative worst-case scenario. This calculation is based on a conservative

worst-case scenario of a dead-on-arrival LEMUR-2 satellite, launched in its maximum mass

configuration,4 and is still well within the standard of twenty-five years of mission completion

and thirty years of launch set forth in Requirement 4.6.1 of NASA-STD-8719.14A

(“Requirement 4.6.1”).5 This analysis is more conservative than the analysis conducted by most

other operators, who do not calculate orbital dwell time and do not limit themselves to an orbit

based on a worst-case, dead-on-arrival basis.6 The actual expected lifetime is seven to fourteen

years at this worst-case altitude, depending on the initial mass of the satellite.

3
See 47 C.F.R. § 25.114(d)(14)(iii).
4
As mentioned, LEMUR-2 Phase IC and Phase II satellites have a nominal launch configuration of 4.5 kg; however,
the mass capacity may be up to 6 kg maximum to accommodate any hosted payload scenarios. Both nominal and
maximum cases are included in the ODAR for collision risk and lifetime analyses. See Exhibit C.
5
See Process for Limiting Orbital Debris, NASA-STD-8719.14A § 4.6.1 (Dec. 2011).
6
As of today, no LEMUR-2 satellite has been dead on arrival.

2

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

Spire has run an analysis measuring dwell times at inclinations from equatorial to sun

synchronous to ensure that changes in inclination do not cause Spire to violate Requirement 4.6.1

at a maximum deployment apogee of 650 km. The results indicate that changes in inclination do

not meaningfully affect orbital dwell times and that at any inclination at 650 km the 25-year

requirement is met. 7

Orbital Lifetime (years) vs. Altitude & Inclination (km)
SSO
30
Equatorial
25
10 deg

Max Altitude
20
20 deg
15 30 deg
10 40 deg

5 50deg
60 deg
0
375 400 425 450 475 500 525 550 575 600 625 650 675 700 70 deg
Altitude (km) 80 deg
90 deg

Spire has also run an analysis measuring dwell times across the entire solar cycle to

ensure that changes in launch schedule do not cause Spire to violate Requirement 4.6.1 at a

maximum deployment apogee of 650 km.

7
See NASA-STD-8719.14A § 4.6.1; see also Mitigation of Orbital Debris, Second Report and Order, 19 FCC Rcd
11567 ¶¶ 61, 83 (2004).

3

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

onboard sequence fails, solar array deployment can be commanded from the ground. If a

satellite is non-communicative, an entirely passive and redundant fail-safe is included on all

satellites in the form of a burn wire. The tensile strength of the burn wire has been tested and

verified to degrade to a breaking point after 3600 hours or 150 days of UV radiation exposure. 9

Spire’s worst-case scenario for dwell time conservatively models five years of non-deployed

solar panels and no loss of altitude during those five years even though a dead-on-arrival satellite

still has surface area that would cause at least some altitude loss.

II. Re-entry Hazards

Spire’s post-mission disposal plan is to allow its satellites to passively re-enter the

atmosphere and completely burn up upon re-entry.10 Spire has used NASA DAS to review the

survivability of major components upon re-entry and found that no objects are expected to

survive re-entry, putting the risk to human life (both on the ground and in aircraft) at 0. This

calculation is orders of magnitude lower than legacy satellite busses.

III. Planned Release of Debris

Spire’s Phase IC satellites with a hosted payload and Phase II satellites will not undergo

any planned release of debris. Spire also conducts extensive acceptance level environmental

testing of all of its satellites to provide further confidence in the structural integrity of the

satellite in launch and space environments. In fact, because Spire launches with every major

launch rocket that takes secondary payloads, including Falcon 9; Antares/Cygnus; Atlas-5;

PSLV; Soyuz; and H-II, the satellite will be subjected to a battery of different testing standards,

9
See Application of Spire Global, Inc., File No. SAT-LOA-20151123-00078, Test Summary: Tensile Properties
Test with Accelerated UV Aging A Demonstration of NOAA DeOrbit Guideline Compliance in an ‘Edge Case’
Scenario, Exhibit E (filed Nov. 23, 2015).
10
See 47 C.F.R. § 25.114(d)(14)(iv).

5

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

including those required by NASA for International Space Station (“ISS”) deployments.

IV. Limiting the Probability of Accidental Explosions

Phase IC satellites with a hosted payload and Phase II satellites have no propulsion and

accordingly do not carry highly volatile rocket propellant. The only energy sources (kinetic,

chemical, or otherwise) onboard the spacecraft is a Lithium-Ion battery system and reaction

wheels.

The battery pack on board the Phase IC satellites with a hosted payload and Phase II

satellites complies with all controls/process requirements identified in NASA Report JSC-20793

Section 5.4.3 to mitigate the chance of any accidental venting/explosion. 11 A battery cell

protection circuit manages the charging cycle, performs battery balancing, and protects against

over and undercharge conditions. The batteries will not be passivated at End-of-Mission due to

the low risk and low impact of explosive rupturing. The maximum total chemical energy stored

in the battery pack is ~144kJ (~288kJ total).

The only failure mode of the reaction wheel assemblies that could lead to creation of

debris would be breakup of the wheels themselves due to mechanical failure while operating at a

high angular rate. Risk mitigation strategies for breakups due to the reaction wheels include

limiting the maximum rotational speed of the wheels and containing them within a sealed

compartment.

V. Collisions with Large Debris

The collision risk posed by the Phase IC satellites with a hosted payload and Phase II

satellites continues to be among the lowest in the satellite industry due to their very small surface

11
See Crewed Space Vehicle Battery Safety Requirements, NASA Report JSC-20793 § 5.4.3 (Jan. 2014),
https://standards.nasa.gov/file/657/download?token=DUcHF-J7.

6

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

area and mass. Using NASA DAS, Spire has calculated the risk of collision for all deployments

sought in this application. The highest probability of collision occurs for the highest orbit of 650

km. Even that probability is 1 x 10-5 over its entire orbital lifetime for a fully deployed satellite

(i.e., the maximum surface area). 12 This probability is hundreds of times lower than legacy

satellite busses’ probability of collision in their worst-case orbits. Full details of the NASA DAS

analysis with respect to collision with large objects for each deployment sought in this

application are contained in the ODAR attached to this application.13

Spire participates in a sharing agreement with the Joint Space Operations Center

(“JSpOC”) to better coordinate collision avoidance measures and receive conjunction threat

reports. Spire’s satellites carry onboard Global Positioning System (“GPS”) receivers that

provide for precise orbital position determination. Spire also receives from JSpOC updated two-

line element sets, or “TLEs,” which facilitate the identification and tracking of Spire’s satellites.

JSpOC has a direct line to Spire’s satellite operations team that is accessible twenty-four hours

per day/seven days per week to ensure that Spire can take immediate action to coordinate

collision avoidance measures. Spire’s GPS-Radio Occultation instrument has capabilities that

allow it to determine the precise location of a satellite down to two centimeters. To Spire’s

knowledge, such precise location capabilities are non-existent outside the context of very large

government satellites and do not exist for any commercial operator.14 This hyper-precise

location data will allow the satellites to have orbits projected out with extreme precision, thus

greatly lowering the number of false positive conjunction alerts and making collision avoidance

12
See Exhibit C.
13
See id.
14
For reference, Spire’s agreement with Orbcomm License Corp. specifies location accuracy of 20 meters as a
threshold. See Application of Spire Global, Inc., File No. SAT-LOA-20151123-00078, Spire Global - ORBCOMM
Agreement (filed Sept. 15, 2016). TLEs received from JSpOC have accuracy measured in kilometers.

7

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

measures far more well informed. Spire currently provides ephemeris data, available from

public sources, online.15 In the near future, Spire will begin publically sharing enhanced

ephemeris data, using the hyper-precise location data from its proprietary GPS-Radio

Occultation instrument.

Special care is also given to minimizing the potential for collision with manned

spacecraft, including the ISS. The operational altitude of the ISS is approximately 400 km.

Spire will coordinate with NASA to assure protection of the ISS on an ongoing basis. Because

Spire participates in many ISS deployments (including above station deployments expressly

approved by the ISS program on a launch-by-launch basis), ISS program management has a

detailed understanding of the Phase IC satellites with potential hosted payloads and Phase II

satellites.

Spire will work closely with its launch providers to ensure that the satellites are deployed

in such a way as to minimize the potential for in-plane collision. The risk is further mitigated

with the typical small deployments undertaken by Spire.

Further, in advance of this filing, Spire has reached out to the other low-Earth orbit

operators at or below 650 km that are identified in the Commission’s Approved Space Station

List and has informed them of Spire’s intention to coordinate to further mitigate any collision

risks.16

The Commission’s rules call upon applicants to specify the accuracy, if any, with which

the orbital parameters of their non-geostationary satellite orbit space stations will be

15
See Open TLE Service, Spire, tle.spire.com (last viewed Oct. 17, 2017). To obtain the ephemeris data for any
particular LEMUR-2 satellite, type in the LEMUR-2’s NORAD ID after “tle.spire.com/” in the URL bar.
16
See Approved Space Station List, FCC, https://www.fcc.gov/approved-space-station-list (last updated Sept. 25,
2017).

8

Exhibit B
Spire Global, Inc.
47 C.F.R. § 25.114(d)(14) Submission

maintained.17 Because the Phase IC satellites with a hosted payload and Phase II satellites will

not carry maneuvering propellant, Spire will not maintain satellite inclination angles, apogees,

perigees, and right ascension of the ascending node to any specified degrees of accuracy.

VI. Collisions with Small Debris or Meteoroids

Spire used NASA DAS to confirm that the Phase IC satellites with a hosted payload and

Phase II satellites meet the requirements of 4.5-2.18

17
See 47 C.F.R § 25.114(d)(14)(iii).
18
See Process for Limiting Orbital Debris, NASA-STD-8719.14A § 4.5-2 (Dec. 2011).

9

Document Created: 2019-07-03 06:43:06
Document Modified: 2019-07-03 06:43:06

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