AI 20.3 Final Report

Support to the IADC Space Debris Mitigation Guidelines

Foreword

This document provides the readers of IADC Space Debris Mitigation Guidelines (first edition dated 15
October 2002

)

with the purpose, feasibility, practices, and tailoring guide for each recommendation

addressed in the Guidelines. Much of this information was based on various documents, research
papers, and opinions that were introduced by IADC member agencies.

The next table depicts the category of typical debris, their causes, and recommendations from IADC.
Several national and international organisations of the space-faring nations have established Space
Debris Mitigation Standards or Handbooks to promote efforts to deal with space debris issues. The
contents of these Standards and Handbooks may be slightly different from one another, but their
fundamental principles are the same as the IADC Guidelines: (1) preventing on-orbit break-ups, (2)
removing spacecraft and orbital stages that have reached the end of their mission operations from the
densely populated orbital regimes, and (3) limiting the objects released during normal operations.

Categories

Causes

Debris Sources

operational debris

(fasteners, covers, wires)

objects released for experiments

(needles, balls, etc.)

tethers designed to be cut after experiments

objects released
by design

others

(released before retrieval)

fragments caused by ageing

(flakes of paints and blankets derived from

degradation)

tether systems cut by debris

objects released before retrieval to ensure safety

liquids with high density (leaked from the nuclear power system, etc.)

Mission-
related
objects

(Parts
Released
During
Mission
Operation)

unintentionally
released objects

particles ejected from solid motors

destruction for scientific or military experiments

(including self-

destruction, intentional collision, etc.)

destruction prior to re-entry in order to minimise ground casualty

intentional
destruction

destruction to ensure security of on-board devices and contained data

explosion caused by failure during mission operation

accidental
break-ups

explosion caused by command destruct systems, residual propellants,
batteries, etc., after mission termination

fragments caused by collision with catalogued objects

On-orbit
break-ups

on-orbit
collisions

fragments caused by collision with un-catalogued objects

Mission-terminated space
systems

systems left in near-GEO, GTO, LEO, and HEO

Table 2.1-1

Recommendations / Requirements S pecified in Major Debris Mitigation Standards

Mitigation Measures

IADC Guidelines

NASA

JAXA (NASDA)

CNES

US standard

FKA

EDMS

Operational Debris

Addressed in 5.1

Addressed

Required

Addressed

Required

Tether

Addressed in 5.1

Addressed

Required

Mission
Related
Objects

Others

Addressed in 5.1

Required

Intentional Break Up

Addressed in 5.2.3

Limited

Required

Residual Propellants

Addressed in 5.2.1

Required

Batteries

Addressed in 5.2.1

Required

Destruct Devices

Addressed in 5.2.1

Required

Pressure Vessels

Addressed in 5.2.1

Required

Probability<10

-4

Passivation within 1
year.

Required

Probability<10

-4

Passivation within 1
year.

-orbit Break

ups

Wheels

Addressed in 5.2.1

During operations;
Probability <10

-4

Depletion at end of
mission life.

Generally required
during mission
operation and after
the end of mission.

Collision Avoidance

Addressed in 5.4

(Addressed in another
document for human
flight)

Required for launch,
and Recommended
for S/C

Required
If necessary

Required for
manned mission

Required
If necessary

Collision with Large
Debris

Addressed in 5.4

Assess collision
Probability

Avoidance is
Recommended

Recommended
(Assess probability.)

Recommended
(Design & Operation)

Assess collision
Probability

Assess collision risk

Break

-up

aused by

ollision

Collision with Small
Debris

Addressed in 5.4

Assess collision
Probability

Protection is
Recommended

Recommended
(Protection, etc.)

Assess collision
Probability

Reorbit of GEO S/C

235 km
+ (1,000.Cr..A/m)

300 km
+ (1,000..A/m)

235 km
+ (1,000.Cr..A/m)

Above 36,100 km
(300km+GEO)

235 km
+ (1,000.Cr..A/m)

235 km
+ (1,000.Cr.A/m)

GEO

Lower limit of GEO
Inclination

-200 km
-15 < Inc. < 15 deg.

-500 km

-H: reorbit distance
-15 < Inc. < 15 deg.

-500 km

-H: reorbit distance
-15 < Inc. < 15 deg.

Removal

Addressed in 5.4

Within 25 years

Within 25 years, if
possible

Feasibility of disposal:
P> 0.99 Within 25 years

Within 25 years

Reduction of orbit
lifetime

Feasibility of disposal:
P> 0.99 Within 25 years

Graveyard Orbit

Not addressed
( > 2000 km)

2,500 - 19,900 km
20,500- 35,288 km

Not addressed
( > 2000 km)

2000 - (GEO-H)
H= reorbit distance

2,000 - 19,700 km

20,700 - 35,300 km

2000 - (GEO-H)
H= reorbit distance

Orbital Retrieval

Addressed in 5.3.2

Addressed

Recommended

Addressed

From

LEO

MEO

Ground Risk

Addressed in 5.3.2

Addressed

Required (Ec<10

-4

)

Required

Addressed (Ec<10

-4

)

To be assessed

(Ec<10

-4

) (TBC)

Lifetime Reduction

Addressed in 5.3.2

Addressed

Recommended

Required

Addressed

Required

Deorbit from GTO

Addressed in 5.3.1

Addressed

Recommended

Required

Addressed

Required

Mission

erminated

Systems

R/B

L/O time selection

Periodical M onitor

Addressed in 5.2.2

Required

Failure

Removal from Orbit

Addressed in 5.2.2

Addressed

Required

Addressed

Remarks for symbols a : Semi-major axis, Cr :Solar Pressure Coefficient, A/m : Area-to-mass ratio, Ec: Casualty expectation

NASA:

NASA Safety Standard 1740.14

[2]

JAXA (NASDA):

NASDA-STD-18A

[3]

CNES:

MPM-50-00-12

[4]

FKA

: Mitigation of Space Debris Population

[5]

US Standard

: Orbital Debris Mitigation Standard Practices

[6]

EDMS:

European Space Debris Safety and Mitigation Standard (draft: June 2000)

[7]

3 Terms and definitions

The following terms and definitions are added for the convenience of the readers of this document.
They should not necessarily be considered to apply more generally.

3.1 Space Debris

Space debris are all man made objects including fragments and elements thereof, in Earth orbit or
re-entering the atmosphere, that are non functional.

Reference:

The term of space debris was defined in more detail as below in

Technical Report on Space

Debris

, 1999, by UN/COPUOS/STSC.

[8]

“

Space debris are all man-made objects, including their fragments and parts, whether their owners

can be identified or not, in Earth orbit or re-entering the dense layers of the atmosphere that are
non-functional with no reasonable expectation of their being able to assume or resume their
intended functions or any other functions for which they are or can be authorized”.

Detail:

As explained in Table 1.1-1, fluids can also constitute a type of debris, particularly if their densities
are high, such as NaK leaked from nuclear power systems.

3.2 Space Systems

Spacecraft and orbital stages are defined as space systems within this document.

3.2.1 Spacecraft



an orbiting object designed to perform a specific function or mission (e.g.

communications, navigation or Earth observation). A spacecraft that can no longer fulfil its
intended mission is considered non-functional. (Spacecraft in reserve or standby modes
awaiting possible reactivation are considered functional.)

3.2.2 Launch vehicle

– any vehicle constructed for ascent to outer space, and for placing one or

more objects in outer space, and any sub-orbital rocket.

3.2.3 Launch vehicle orbital stages



any stage of a launch vehicle left in Earth orbit.

3.3 Orbits and Protected Regions

3.3.1 Equatorial radius of the Earth -

the equatorial radius of the Earth is taken as 6,378 km

and this radius is used as the reference for the Earth’s surface from which the orbit regions are
defined.

3.3.2 Protected regions



any activity that takes place in outer space should be performed while

recognising the unique nature of the following regions, A and B, of outer space (see Figure 1),
to ensure their future safe and sustainable use. These regions should be protected regions
with regard to the generation of space debris.

(1)

Region A,

Low Earth Orbit

(or LEO) Region – spherical region that extends from the Earth’s

surface up to an altitude (Z) of 2,000 km

Note: The orbital region used for manned flights, of special concern due to risks of in-orbit
casualties, is included in the Region A, Low Earth Orbit Region.

(2) Region B, the

Geosynchronous Region

- a segment of the spherical shell defined by the

following:

lower altitude = geostationary altitude minus 200 km

upper altitude = geostationary altitude plus

200

-15 degrees

≤

latitude

≤

+15 degrees

geostationary altitude (Z

GEO

) = 35,786 km (the altitude of the geostationary Earth orbit)

GEO

- 200km

Z = 2000km (LEO)

GEO

+ 200km

Equator

Earth

Region A

Region B

Z = Z

GEO

Figure 1 - Protected regions

3.3.3 Geostationary Earth Orbit (GEO)



Earth orbit having zero inclination and zero

eccentricity, whose orbital period is equal to the Earth's sidereal period. The altitude of this
unique circular orbit is close to 35,786 km.

3.3.4 Geostationary Transfer Orbit (GTO)



an Earth orbit which is or can be used to transfer

space systems from lower orbits to the geosynchronous region. Such orbits typically have
perigees within LEO region and apogees near or above GEO.

3.4 Mitigation Measures

and Related Terms

3.4.1 Passivation



the elimination of all stored energy on a space system to reduce the chance

of break-up. Typical passivation measures include venting or burning excess propellant,
discharging batteries and relieving pressure vessels.

3.4.2 De-orbit



intentional changing of orbit for re-entry of a space system into the Earth’s

atmosphere to eliminate the hazard it poses to other space systems, by applying a retarding
force, usually via a propulsion system.

3.4.3 Re-orbit



intentional changing of a space system’s orbit

3.4.4 Break-up



any event that generates fragments, which are released into Earth orbit. This

includes:

(1)

An explosion caused by the chemical or thermal energy from propellants, pyrotechnics and
so on

(2)

A rupture caused by an increase in internal pressure

(3)

A break-up caused by energy from collision with other objects

However, the following events are excluded from this definition:

•

A break-up during the re-entry phase caused by aerodynamic forces

•

The generation of fragments, such as paint flakes, resulting from the ageing and
degradation of a space system.

3.5 Operational Phases

3.5.1 Launch phase -

begins when the launch vehicle is no longer in physical contact with

equipment and ground installations that made its preparation and ignition possible (or when the
launch vehicle is dropped from the carrier-aircraft, if any), and continues up to the end of the
mission assigned to the launch vehicle.

3.5.2 Mission phase -

the phase where the space system fulfils its mission. Begins at the end of

the launch phase and ends at the beginning of the disposal phase.

3.5.3

Disposal phase

- begins at the end of the mission phase for a space system and ends

when the space system has performed the actions to reduce the hazards it poses to other
space systems.

4 General Guidance

During an organisation’s planning for and operation of a space system it should take systematic
actions to reduce adverse effects on the orbital environment by introducing space debris mitigation
measures into the space system's lifecycle, from the mission requirement analysis and definition
phases.

In order to manage the implementation of space debris mitigation measures, it is recommended
that a feasible Space Debris Mitigation Plan be established and documented for each program and
project. The Mitigation Plan should include the following items:

A management plan addressing space debris mitigation activities

A plan for the assessment and mitigation of risks related to space debris, including applicable

standards

The measures minimising the hazard related to malfunctions that have a potential for

generating space debris

A plan for disposal of the space system at end of mission

Justification of choice and selection when several possibilities exist

Compliance matrix addressing the recommendations of these Guidelines.

Purpose:

Space debris mitigation measures should be taken into consideration from the very early phases
of project planning. Also, adequate decision-making is expected in each of the planning, design,
operation, and disposal phases. Section 4 recommends that space debris mitigation activities be
included in phased planning, and that organisational and systematic actions be taken according
to the authorised plan.

Practices (Phased planning):

(1) System concept, mission planning, launch configuration, operation planning, and disposal

procedures should be developed with consideration for their effects on the orbital
environment. It may be noted that major mitigation procedures should be fixed in the very
early phases (mission definition and conceptual design phases), and the issues relevant to
debris generation should be identified in the preliminary design review and be solved by a
detailed design review. (NASA Safety Standard and JAXA standard formally define two
design reviews, PDR (Preliminary Design Review) and CDR (Critical Design Review), to
assess the mitigation actions.) A Mitigation Plan should be developed to control these
activities.

(2) The disposal phase should be clearly considered in mission planning.

Practices (Mitigation Plan):

Space debris mitigation issues should be identified and dealt with in a project like several other
issues, such as safety. It is therefore suggested to include this issue in the scope of the Product
Assurance manager of the project. It should not be necessary to issue a large amount of
documentation:

•

a plan for control of debris mitigation during the development of the project and the
operations, including the disposal and passivation of the spacecraft, and

•

the necessary technical documentation answering to this plan and identifying the measures
to limit debris generation and the assessment of the satellite disposal and its operational
aspects.

Of course, the compliance status of any recommendations is addressed via standard practices.

Moreover, it would be desirable for the Mitigation Plan to include the following elements:

(1)

Concept:

The Mitigation Plan could include management organisation, major event,

schedule, potential for generating debris, assessment plan, related documents, and the
results of design tailoring.

(2)

Organisation

: Each agency (and its contractors) may assign a group or individual bearing

responsibilities to study, plan, implement, and review space debris mitigation  activities.
The assigned group or individual should be provided with enough authority and resources
required to accomplish and fulfil this duty and  should report the progress status to the
project manager. Usually, such a role would be assigned to a Safety & Mission Assurance
department.

(3)

Management

: Major events and schedule, potential debris sources, disposal plan,

assessment plan, and related documents would be helpful.

(4)

Mitigation measures:

The technical basis for mitigation measures corresponding to each

debris source and disposal plan should be described.

(5)

Compliance matrix:

In each design phase, compliance among system requirements,

design, manufacturing, and the operation plan should be reviewed and recorded in a
compliance matrix. If some requirements or recommendations are tailored, the facts
should be recorded as specified immediately below.

Tailoring guide:

The recommendations in this document can be tailored before being applied. The results of
tailoring, however, should be agreed upon among the departments responsible for each project
and should be submitted and reviewed by the responsible review committee.

The facts of tailoring and the basis for such should be recorded in the Space Debris Mitigation
Plan. Typical examples for tailoring are as follows:

(1)

for space systems already in progress in their development phase to some extent, only
practically feasible recommendations would be applied, and

(2)

comprehensive studies for various conditions including economical, technical, and other
situations concerned with debris mitigation measures would identify the practically feasible
range of recommendations to be applied.

5 Mitigation Measures

5.1 Limit Debris Released during Normal Operations

In all operational orbit regimes, space systems should be designed not to release debris during
normal operations. Where this is not feasible any release of debris should be minimised in
number, area and orbital lifetime.

Any program, project or experiment that will release objects in orbit should not be planned unless
an adequate assessment can verify that the effect on the orbital environment, and the hazard to
other operating space systems, is acceptably low in the long-term.

The potential hazard of tethered systems should be analysed by considering both an intact and
severed system.

Purpose:

Approximately 12% of the  current catalogued objects are debris released during normal
operations.    The release of fasteners, yo-yo weights, nozzle covers, lens caps,  and  multiple
payload mechanisms should be kept to a minimum.

Feasibility:

It is relatively easy, both technically and economically, to take mitigation measures against these
objects. Many agencies have already reported to be taking such action.

Satellite manufacturers usually avoid intentional debris generation, since this debris might
remain very close to the satellite and become a danger to the satellite itself (blocking
mechanisms, obstructing the field of view, etc.). It is therefore a sound requirement for
spacecraft manufacturers to preclude intentional debris generation during normal operations.

Practices:

The number of objects released during nominal operations to become orbital debris should be

minimised by design. The following are examples of these objects.

(1) Launch vehicle connectors and fasteners: separation bolts, clamp bands, etc.

(2) Fairings: fairings and adapters for launching multiple payloads, etc.

(3) Covers: nozzle closures, etc.

(4) Others: yo-yo weights and lines, etc.

Apogee motor cases or engines should not separate or be left in an orbit passing through the
protected regions. If this is not possible, they should be left so as not to interfere with the
protected regions, and they should be passivated.

Note: Solid Rocket Motors release solid particles during and after burning. The precise nature of
the amount and distribution of the ejecta are unclear, and the improvement of solid propellants
and motor insulation to minimise the number of released objects is currently difficult.

Practice (tethers):

Tethers several thousand meters in length and a few millimetres in diameter have a large
probability to be severed by small debris. New multi-strand tether designs can reduce the risk of

severing. At the end of missions, it is recommended that tethers be retracted to reduce the
probability of collision with space systems.

Tailoring guide:

(1) Fairings:

support structural elements left in orbit during a multiple payload mission may be

released, if there are no feasible alternative measures at present. Fortunately, when
released at low altitude their orbital lifetimes can be relatively short if their area-to-mass
ratios are high.

(2) Orbital lifetime:

released objects whose orbital lifetime is short (less than 25 years, for

example) could be assessed as allowable.

(3) Mission requirements to release objects:

missions that require releasing objects should

be submitted to the review board of the agencies to assess their necessity and their effects
on orbital environment.

(4) Paint flakes and other objects released by degradation:

by exposure to the space

environment (ultraviolet radiation, atomic oxygen, thermal cycling, and micro-particle
impacts)

paint, surface materials, and possibly deployment devices can deteriorate and

generate fragments. However, it is impossible to present standards or recommendations with
regard to how many years materials should withstand the space environment.

(5) Tethers:

tethers can exacerbate the debris environment, but can also be used to reduce

orbital lifetime. In the planning of tether systems, these advantages and disadvantages
should be assessed.

5.2 Minimise the Potential for On-Orbit Break-ups

On-orbit break-ups caused by the following factors should be prevented using the measures
described in 5.2.1

−

5.2.3:

(1)

The potential for break-ups during mission should be minimised

(2)

All space systems should be designed and operated so as to prevent accidental explosions
and ruptures at end-of- mission

(3)

Intentional destructions, which will generate long-lived orbital debris, should not be planned or
conducted.

Purpose

The most common source of space debris is on-orbit break-ups of space systems. Approximately
43% of catalogued objects, and 85% of all space debris larger than 5 cm in diameter stem from on-
orbit break-ups. This section recommends the minimum effort to prevent their generation.

According to the NASA database provided by NASA/JSC (Orbital Debris Program Office), at the
end of February 2004 at least 174 orbital fragmentations (excluding aerodynamic break-ups) have
occurred. Among them, 79 events (45%) had propulsion-related causes, 54 events (31%) were
due to intentional destruction, and 8 events (5%) were due to the rupture of battery cases. Only
one case was due to an accidental on-orbit collision. The causes of the remaining 18% are
unknown.

Figure 5.2-1 shows the cause of break-ups, and Fig.5.2-2 shows differences of the causes
depending on spacecraft and rocket bodies. Intentional destruction is the major cause for
spacecraft break-up, while the propulsion system is the major responsible for rocket-body break-
up. No spacecraft has yet been observed to have broken up as a result of liquid propulsion failure,
and no rocket body as a result of battery failure.

5.2.1 Minimise the potential for post mission break-ups resulting from stored energy

In order to limit the risk to other space systems from accidental break-ups after the completion of
mission operations, all on-board sources of stored energy of a space system, such as residual
propellants, batteries, high-pressure vessels, self-destructive devices, flywheels and momentum
wheels, should be depleted or safed when they are no longer required for mission operations or
post-mission disposal. Depletion should occur as soon as this operation does not pose an
unacceptable risk to the payload. Mitigation measures should be carefully designed not to create
other risks.

Purpose

The most important and effective measure is the prevention of break-ups. Expenditure of
residual propellants and high-pressure fluids and the switching-off of battery charging lines are
typical measures. More detailed recommendations are addressed below.

(1) Residual propellants and other fluids, such as pressurants, should be depleted as thoroughly

as possible, either by depletion burns or venting, to prevent accidental break-ups by over-
pressurisation or chemical reaction.

Purpose

Residual propellant is the most common cause of on-orbit break-ups. Many accidental break-ups
have been caused by orbital stages possessing hypergolic propulsion systems with common
bulkhead tanks. But even cryogenic propulsion systems have apparently ruptured as a result of
propellant evaporation and resulting over-pressurisation.

The above recommendation can prevent such propulsion-related break-ups. However, it is
sometimes difficult to know the exact amount of remaining propellant, since sensors can give
incorrect information, for example at the end of life of a satellite.

Practices in design and operation of LV

Accidental mixing of hypergolic propellants should be prevented by design. For example, the
common bulkhead or the lines having a path between the oxidiser and fuel feeding systems, that
would increase the risk of mixing of oxidiser and fuel, should be properly

designed and used

. In

cases where a common bulkhead tank system is designed, the pressure of the inner tank should
be kept higher than the outer tank in order to prevent a rupture of the common bulkhead. This
effort to keep differential pressure should be also applied during the final venting or burning
operation to prevent bulkhead breakage.

Even in the case of a monopropellant or cryogenic propellant system or a separated tank
system, residual propellant should be vented or burned at the end of mission. Venting lines
should be designed to prevent blockage from freezing propellants.

Consequently, an adequate  sequence of  valve operation, sufficient electric power to sustain
vent-valve  operation, and a monitoring system to sense complete depletion are recommended.
The sequence of events should be officially planned and reviewed.

Impact on the operating spacecraft

Depletion burns and venting may generate impulses that will disturb the attitude of spacecraft or
rocket body. Especially in the case of venting propellants, a specific design (torque-free venting
system) or operation may be required to cancel the impulse.

Tailoring guide:

Some propellant may be allowed to become trapped in lines as long as the amount is insufficient
to cause a break-up by ignition or pressure increase.

(2) Batteries should be adequately designed and manufactured, both structurally and electrically,

to prevent break-ups. Pressure increase in battery cells and assemblies could be prevented by
mechanical measures unless these measures cause an excessive reduction of mission
assurance. At the end of operations battery charging lines should be de-activated

Purpose

Historically, eight accidental  satellite  break-ups  have been caused by  battery ruptures. The
above guideline  recommends  considering  measures during  design,  manufacturing, and
operation to prevent such malfunctions.

Practices:

The main causes of  battery  break-ups are inadequate  design and manufacturing in both
structural and electrical aspects, as well as operational errors. Usually, battery cases have
enough strength to withstand the increase of inner pressure under normal conditions and will not
cause  a satellite  break-up. However,  system qualification for long periods can be difficult. In
addition, there is a break-up risk in case of hypervelocity impact.  Shutting-off charging lines and
completely discharging the battery will substantially reduce the break-up risk.

Relays (and the command line) to shut off the charging lines and heaters or other high power
loads to discharge batteries are recommended.

In any case, there are electrical and chemical events able to generate gas inside the cells, and
then to make the pressure increase without limitation. The space debris limitation should rely on
electrical protection, rather than on battery mechanical re-enforcement, for example,

in case of a potential leakage current, the implementation of a resistor between battery
and structure might be recommended or

a high depth of discharging (DOD) may lead to a cell being inversely polarised if the cell
is not homogeneous.

It is therefore recommended to perform a specific power subsystem study aimed at defining an
adequate architecture that would be able to cope with end-of-life electrical passivation needs for
the various families of satellites.

Tailoring guide

For the passivation itself, some documents recommend the implementation of a relay or relays
for disconnection from the charging lines and the associated command line. From French
experience, such a disconnection capability was implemented on the SPOT platform (and this
command was used for the disposal of SPOT 1), but it is usually not implemented on
Telecommunication satellites or many small satellites. An erroneous command to a system
employing a solitary relay could be a single point of failure, which is often considered
unacceptable in satellite design. An alternative would be to install independent relays in parallel.

Pressure relief valves for battery cells might reduce reliability. Such measures have been taken
for the battery cells and assemblies on some launch vehicles (e.g., Ag-Zn batteries), but less
often for spacecraft.

(3) High-pressure vessels should be vented to a level guaranteeing that no break-ups can occur.

Leak-before-burst designs are beneficial but are not sufficient to meet all passivation
recommendations of propulsion and pressurisation systems. Heat pipes may be left
pressurised if the probability of rupture can be demonstrated to be very low.

Purpose

This recommendation is mainly applied to regulated systems that consist of an upstream high-
pressure vessel and a downstream, regulated-pressure vessel.

Practices

(1) blow down system

: The upstream pressurant should be vented at least to less than the mean

operational pressure of the downstream vessel.

(2) tanks with a bladder:

Tanks in which fuel and pressurant are separated by a bladder should

contain a mechanism for totally venting gases. In cases where such a mechanism is not
implemented, enough safety margin to prevent break-up under expected solar heating should
be adopted.

(3) LBB design

: Leak-before-burst (LBB) designs are beneficial but not sufficient in preventing

potential break-up scenarios. They are normally effective when the rise in pressure is gradual.
On the other hand, the cause of the significant 1996 Pegasus HAPS break-up has been
assessed to be the rapid over-pressurisation and failure of the main propellant tank (which had
a leak-before-burst design) when a regulator between the propellant tank and pressurant tank
failed.

Tailoring guide:

(1) Although helium bottles of launch vehicles sometimes do not have vent mechanisms, a bleed

valve of the pressure regulator will gradually decrease the inner pressure to avoid unsafe
levels.

(2) In some propellant tanks with a bladder and no vent valve, the pressurising gas might be

trapped in the tanks and cannot be vented. Usually the pressure will decrease during normal
operations to safe levels (less than one tenth of initial pressure), but enough margin should be
taken for the case that some failure would keep the initial pressure, e.g., main engine failure.

(3) Heat pipes are highly pressurised and, therefore, a source of stored energy. However, in the

usual design process, they have enough structural integrity to prevent such accidents. NASA
has recognised that no particular action is usually needed to de-pressurise heat pipes and
intends to revise NSS 1740.14 for such cases.

(4) Self-destruct systems should be designed not to cause unintentional destruction due to

inadvertent commands, thermal heating, or radio frequency interference.

Purpose

Unintentional triggering of self-destruct systems can produce break-ups.

Practices

Unintentional activation of self-destruct systems is a complex topic and many sources may
trigger it, for example static electricity discharge, impact, …

Destruction command receivers should be turned off as soon as they are no longer needed.

Thermal insulation should protect the explosive charge to keep its temperature less than its
cook-off temperature.

(5) Power to flywheels and momentum wheels should be terminated during the disposal phase.

Tailoring Guide

Usually no action will be required if the batteries have been discharged. Flywheels and
momentum wheels will usually stop shortly after cutting off the power supply due to friction.

(6) Other forms of stored energy should be assessed and adequate mitigation measures should be
applied.

Purpose

“Other forms” covers all other possible sources of break-ups that have not be mentioned above.
Such forms might be design-dependent and should be assessed; adequate mitigation measures
should then be applied.

Practices

A list of all elements with stored energy (mechanical, thermal, chemical, etc.) should be
established and subject to assessment on each project. Examples are as follows:
1)  chemical experimental devices,
2)  mechanical devices that might retain a large amount of stress or kinetic energy,
3)  thermal devices, and

pyrotechnic devices.

5.2.2 Minimise the potential for break-ups during operational phases

During the design of a space system, each program or project should demonstrate, using failure
mode and effects analyses or an equivalent analysis, that there is no probable failure mode leading
to accidental break-ups. If such failures cannot be excluded, the design or operational procedures
should minimise the probability of their occurrence.

During the operational phases, a space system should be periodically monitored to detect
malfunctions that could lead to a break-up or loss of control function. In the case that a
malfunction is detected, adequate recovery measures should be planned and conducted;
otherwise disposal and passivation measures for the system should be planned and conducted.

Purpose:

Mission assurance is not explicitly a space debris issue. However, considering the effect of on-
orbit break-ups, an intentional decrease in reliability that is induced by cost reductions, lack of
technology, or time-saving should be avoided for the sake of other operating space systems and
the orbital environment.

Practice:

It is standard practice on satellites, even on the cheapest ones, to identify potential failure modes
and their effects and to monitor on-board or on-ground (depending on the needed reaction delay)
the technological parameters indicating that

(1) a failure has occurred and is likely to propagate to other functions of the vehicle or

(2) a failure is likely to occur (indicated by parameter drift).

Monitoring then allows the ground or the on-board satellite management to take all necessary
passivation measures, in order to eliminate the risk of failure propagation.

A primary recommendation would then be to make sure that all necessary measurement points
are implemented on-board to monitor the physical characteristics (pressure, temperature,
current, etc.) and their drift, in order to detect failures with the potential to lead to debris
generation.

Concerning propulsion, depending on the selected architecture, these actions may consist of
closing or opening some valves to isolate the critical section.

5.2.3 Avoidance of intentional destruction and other harmful activities

Intentional destruction of a space system, (self-destruction, intentional collision, etc.), and other
harmful activities that may significantly increase collision risks to other systems should be avoided.
For instance, intentional break-ups should be conducted at sufficiently low altitudes so that orbital
fragments are short lived.

Purpose:

Intentional destructions have been conducted for the purpose of engineering tests, experiments,
or security assurance (data and technology security) for on-board information. Such activities
should be avoided whenever possible.

In the past, deliberate activities detrimental to the space environment have taken place. Large
numbers of needles were scattered in-orbit for a communications experiment in the 1960’s.

When conducted, intentional destruction or potentially harmful activities should be assessed for
possible damage to other spacecraft.

Tailoring Guide

In rare cases, destruction may be planned to reduce the risk to people on Earth from re-entering
debris objects,  but  this should be conducted at low altitude,  e.g.,  lower than 90 km. However,
keeping the destruct devices in-orbit during mission operation could increase the risk of an  on-
orbit explosion, even if  the  mission duration  is short. Also, to control the destruction in low
altitude may not be easy because of difficulty in  attitude control, protection from aero-heating,
and the maintenance of command lines.

5.3 Post Mission Disposal

5.3.1 Geosynchronous Region

Spacecraft that have terminated their mission should be manoeuvred far enough away from GEO
so as not to cause interference with space systems still in geostationary orbit. The recommended
minimum increase in perigee altitude at the end of re-orbiting, which takes into account all orbital
perturbations, is:

235 km + (1000

x A/m)

where C

: Solar radiation pressure coefficient (typical values are between 1 & 2),

A/m: Aspect area to dry mass ratio [m

/kg]

235 km: Sum of the upper altitude of the GEO protected region (200 km) and the

maximum descent of a re-orbited space system due to luni-solar and
geopotential perturbations (35 km).

The propulsion system for a GEO spacecraft should be designed not to be separated from the
spacecraft. In the case that there are unavoidable reasons that require separation, the propulsion
system should be designed to be left in an orbit that is, and will remain, outside of the protected
geosynchronous region. Regardless of whether it is separated or not, a propulsion system should
be designed for passivation.

Operators should avoid the long term presence of launch vehicle orbital stages in the
geosynchronous region.

Purpose:

To preserve the GEO environment, where the removal of objects by natural forces normally will
require extremely long periods, objects should be moved to a higher region when no longer
useful.

Definitions:

A/m: Aspect area (in m

) over dry mass (in kg) :

A is the effective cross-sectional area of the spacecraft (m

) in the condition when the space

system is sent to super-GEO, usually with solar arrays and antennas in their deployed positions.
The NASA Safety Standard

[2]

on orbital debris provides a simple calculational method for

determining the approximate cross-sectional area for a tumbling vehicle, i.e., “the average cross-
sectional area is 1/4 the surface area. For a simple convex spacecraft body with solar panel
wings, calculate the average cross-sectional area from the surface areas of the spacecraft body,

body

, and the solar panels,

, as

) / 4

body

.”

Mass (kg) is the actual mass at the time that the space system is sent to super-GEO. Usually
this can be considered equal to the dry mass, if all fluids have been burned or released.

(Solar Pressure Radiation Coefficient):

The actual value of C

depends on the surface characteristics (insulators, solar arrays, radiators,

antennas, etc.), their areas, and the vehicle attitude with respect to the sun. There will be some
difference between the case of the golden colour of aluminised Kapton and the black Kapton, but
the total value of C

will not vary significantly because of the large area of the solar paddles and

other components. So C

may be in the range of about 1.2 to 1.5. In addition, the value is

typically expected to decrease with ageing, but usually the value at the beginning of life will be
used as a conservative measure.

Practice (Sending to super-GEO):

To prevent on-orbit collisions in the GEO region, spent space systems (including separated
apogee propulsion systems, that should not ordinarily be separated, and orbital stages of launch
vehicles that conduct direct injection of spacecraft into GEO) might be sent to an orbit higher
than GEO.

The minimum re-orbit distance was determined as the sum of the upper side of the bandwidth of
protected region and the expected orbital change of the re-boosted object due to natural
perturbations.

The following figure shows the GEO protected space region (Fig. 5.3.1-1).

Figure 5.3.1-1 Protected GEO Orbital Region and Reorbit Distance from GEO

Practice (control of eccentricity):

There is no mention of the eccentricity of final orbit, but the eccentricity might be minimized,
since (1) it will help to attain highest perigee altitude if there is much uncertainty in the estimated
quantity of residual propellant, (2) it will minimise the deviation between the apogee and perigee
altitudes which consequently permits a higher relative perigee altitude, and (3) it will increase the
stability of the orbit from luni-solar perturbation.

Penalty

The penalty of this re-orbit operation may be equivalent to 3 months of operation for a typical
N/S and E/W station-keeping GEO satellite.

The following are typical values for the required velocity increase and mass fraction for re-orbit
manoeuvres.

- 200 km

+ 200 km

Protected Region

Effect of luni

-solar and geopotential perturbations = 35 km

- 15degree

Earth

GEO

Effect of Solar radiation pressure = 1000 x C

x A/m

Reorbit distance 235 km + (1000 x C

x A/m)

Super-
GEO

Apogee propulsion system:

In the past, some types of spacecraft have separated their apogee propulsion systems to obtain
better characteristics and efficiencies from the aspect of attitude control, thermal control, and
field of view. Liquid engines are more hazardous than solid motor cases particularly in the event
that they separate while containing residual propellants as sources of break-up energy (for
instance NASDA’s ETS-VI). Such residual propellants should be vented before separation.
Otherwise specific devices (to control venting and to provide energy to open the valves), should
be required to vent immediately after separation.

If unavoidable reasons arise which require separation, the propulsion system should be
designed to be left in a higher orbit as recommended for mission-terminated GEO spacecraft.

Note that current satellite designs or future ones do not usually include separable propulsion
stages. This may be the case for interplanetary missions, which are not in the scope of this report.

Direct Injection into GEO

In the case of direct injection of payloads into orbits near GEO (e.g. US Centaur upper stage)
the best solution might be to insert the upper stage and payload directly into a recommended
disposal orbit above GEO and to have the payload then perform a minor maneuver to place
itself into GEO.

Practice (GTO objects):

To avoid the long-term presence of launch vehicle orbital stages in the geosynchronous region,
the NASA Safety Standard 1704.14

[2]

and NASDA-STD-18A

[3]

recommend that apogee should

decrease to 500 km lower than GEO within 25 years. In JAXA, as a practical procedure, with
considering the perturbation effect on GTO, it is requested that the apogee should descend 550
km lower than GEO.

é h

s p ü

2 0 0

Figure 5.3.1-2 Velocity Increase (

∆

V) and Propellant Mass Fraction with

Parent S/C Dry Mass (

∆

M/M) as Function of Reorbit Distance

5.3.2 Objects Passing Through the LEO Region

Whenever possible space systems that are terminating their operational phases in orbits that pass
through the LEO region, or have the potential to interfere with the LEO region, should be de-
orbited (direct re-entry is preferred) or where appropriate manoeuvred into an orbit with a reduced
lifetime. Retrieval is also a disposal option.

A space system  should be left in an orbit in which, using an accepted nominal projection for solar
activity, atmospheric drag will limit the orbital lifetime after completion of operations. A study on the
effect of post-mission orbital lifetime limitation on collision rate and debris population growth has
been performed by the IADC. This IADC and some other studies and a number of existing national
guidelines have found 25 years to be a reasonable and appropriate lifetime limit. If a space system
is to be disposed of by re-entry into the atmosphere, debris that survives to reach the surface of
the Earth should not pose an undue risk to people or property.    This may be accomplished by
limiting the amount of surviving debris or confining the debris to uninhabited regions, such as broad
ocean areas.  Also,  ground environmental pollution, caused by  radioactive substances, toxic
substances or any other environmental pollutants resulting from  on-board articles, should be
prevented or minimised in order to be accepted as permissible.

In the case of a controlled re-entry of a space system, the operator of the system should inform the
relevant air traffic and maritime traffic authorities of the re-entry time and trajectory and the
associated ground area.

Purpose:

The  LEO region is a collection of  useful orbits that many  countries use for Earth observation,
micro-gravity experiments, communications, space scientific observation and experiments, and
so on. It also includes manned missions conducted during the past 30 years up to an altitude of
600 km. Preserving the orbital environment of this region is very important both for the use of this
region and also for passing through this region to GEO and beyond.  Consequently, the removal
of objects from  LEO as soon as possible after the end of  a  mission is beneficial. Fortunately,
natural forces, especially drag, work to clean debris from this region, although  this is

effective

primarily for satellites below 700 km. I

t is recommended that orbital lifetime be reduced to less than

25 years at the end of mission (approximately 750 km circular orbit for A/m = 0.05 m

/kg, and

approximately 600 km circular orbit for A/m=0.005 m

/kg, depending on solar activity to be more

exact). For a given amount of propellant, lowering perigee only will minimise the remaining
orbital lifetime, compared with lowering both apogee and perigee to a new, lower circular orbit.

This guideline is appropriate for all space systems, regardless of size: satellites without
propulsion systems should not be launched to the orbits within the LEO protected region if their
post-mission lifetime is greater than 25 years.

Practice (Reduction of orbital lifetime):

Computations related to orbital lifetime as a function of initial orbit, air drag and area-to-mass
ratios may be found in many documents. Similarly, the fuel required for decreasing a low orbit
perigee down to a given value is easy to compute. The IADC recommendation is to ensure that
the lifetime after disposal will not exceed 25 years.

IADC Working Group 2 studied the effect of de-orbiting and the result is shown below. [Ref. End-
of-life Disposal of Space Systems in the Low Earth Orbit Region, IADC/WG2]

[11]

General

A combination of mission-related object elimination, passivation and post-mission de-
orbiting to a limited lifetime orbit was found to be successful at mitigating the future LEO

debris environment in the long-term (assuming that launch traffic does not increase
significantly above that seen in recent years).

Post-mission De-orbiting to a Limited Lifetime Orbit

All post-mission lifetimes considered in the study (0, 25 and 50 years) were able to stabilise
overall LEO debris population levels over the next 100 years, and were therefore deemed to
be beneficial.

Longer post-mission lifetimes generally led to higher stabilised population levels (at altitudes
<800 km) and therefore it is desirable to shorten post-mission lifetime as far as possible in
order to reduce population levels and collision risks in the long-term. However, shorter post-
mission lifetimes are more costly for space systems to achieve using on-board propulsion
systems.

Only a modest near-linear increase in de-orbit manoeuvre propellant consumption would be
needed to reduce post-mission lifetime over much of the range considered in this study.
However, it has been found that decreasing post-mission lifetime to very short times would
involve a substantial exponential growth in the de-orbit propellant requirement.

Hence, based on the analysed post-mission lifetimes, a 25-year post-mission lifetime was
found to be the shortest possible without significant and disproportionate increases in de-
orbit propellant consumption.

Therefore, a 25-year post-mission lifetime appears to be a good compromise between an
immediate (or very short lifetime) de-orbit policy which is very effective but much more
expensive to implement, and a 50 or 100 year lifetime de-orbit policy which is less costly to
implement but can lead to higher collision risks in the long-term.

- Any concern for low-altitude manned mission safety in connection with post-mission de-

orbiting is not warranted. Though the population of >10 cm objects will slightly increase in this
region mainly due to perigee lowering, these large disposed objects can be, and are, tracked
and avoided. The benefit to low-LEO altitudes attained by post-mission de-orbiting is a low
and stabilised overall LEO collision rate. This directly prevents significant growth in the
untrackable (but hazardous) centimetre-sized object population at all LEO altitudes, including
low-LEO altitudes where manned missions are operating.

>1cm Population Evolution

EVOLVE Model Projections

200000

400000

600000

800000

1000000

1200000

1400000

1600000

1800000

2000000

100

Projection Year

No. of objects >1cm in LEO

Business As Usual (BAU)

MRO/Explosion Prevention

MRO/Expl Prev, 0-yr De-orbit

MRO/Expl Prev, 25-yr De-orbit

MRO/Expl Prev, 50-yr De-orbit

Figure 5.3.2-1 Debris ( > 1cm) Average Population Evolution from Evolve

MRO: Mission Related Objects are refrained to be released,

Expl Prev: Explosions and other break-up events are prevented,

N-yr De-orbit: S/C & Rocket Bodies are removed within N years from orbit.

[Ref. End-of-life Disposal of Space Systems in the Low Earth Orbit Region, IADC/WG2]

[11]

Case-1: Business as Usual (No mitigation measures are

taken.)

Case-2: Mission Related Objects are refrained to be

released, and explosions are prevented.

Case-3: In addition to case-2, systems are removed

within 50 years after mission termination.

Case-4: In addition to case-2, systems are removed

within 25 years after mission termination.

Case-5: In addition to case-2, systems are removed

immediately after mission termination.

Estimation of Penalty

The propellant requirement to achieve a specified orbital lifetime will be higher if the operating
orbit is high. For example, if orbital lifetime is limited to 25 years after mission completion,  an
amount of propellant equal to  4% of the mass of the vehicle will be required for the disposal
operations from an altitude of 1000 km.

Table 5.3.2-1 Required propellant for lifetime reduction within 25 years

(Isp = 200 sec, A/m = 0.05 m

/kg)

Initial Altitude

Descending

altitude

Final perigee

Altitude

Delta Velocity

Mass Fraction

(Propellant / Dry Mass)

800 km

70 km

730 km

18 m/s

0.8%

1,000 km

370 km

630 km

88 m/s

4.3%

1,500 km

965 km

535 km

236 m/s

11%

2,000 km

1505 km

495 km

349 m/s

17%

[Ref: Space Debris Handbook NASDA -CRT-98006, 1998]

[10]

The IADC WG2 report [End-of-life Disposal of Space Systems in the Low Earth Orbit Region,
IADC/WG2]

[11]

also shows propellant mass for re-orbit as shown in Fig. 5.3.2-2 (in the case of

Isp=260 sec).

Practice (on-orbit retrieval):

With current technology, this option is not feasible for most spacecraft owner/operators.  The
only practical measure is  that of using  the US Space Shuttle. However, NASA would not
encourage this option because even  for the US the use of the Shuttle for satellite retrieval is
normally not warranted, and the risk to the Shuttle crew may exceed the risk to people on Earth

Chemical Propulsion De-orbiting - Fuel (I

=260s)

100

Lifetime Limit (yrs)

Fuel Mass Margin (% sat. mass)

800 km alt, 0.005 m^2/kg

800 km alt, 0.05 m^2/kg

1400 km alt, 0.005 m^2/kg

1400 km alt, 0.05 m^2/kg

Re-orbit (2 burns):

1400 km to 2000 km

Figure 5.3.2-2 Cost of N-year post-mission lifetimes in terms of added fuel mass

assuming use of conventional chemical propulsion systems.

[Ref. End-of-life Disposal of Space Sys tems in the Low Earth Orbit Region, IADC/WG2]

[11]

from uncontrolled re-entry. Providing the Shuttle on a commercial  basis for retrieval is not
possible.   Furthermore, the target object  orbit must be lowered to the Shuttle orbit (600 km at
highest case), propellants and deployed objects must be  passivated, and they must  offer the
proper interfaces. So, until such time that direct retrieval is a more commonly available option
(perhaps by robotic means), this is not a practical solution.

Tailoring guide (Reduction of orbital lifetime)

One can take advantage of anticipated residual propellants set aside for other purposes, e.g.,
initial orbital injection, in determining propellant reserves for disposal manoeuvres.

Purpose (Ground Safety from Objects Surviving Re-entry ):

One effective space debris mitigation measure is the removal of mission-terminated space
objects from useful orbit regions and the disposal of them by aerodynamic heating during re-
entry, if possible. However, the ground casualties that might be caused by fragments surviving
atmospheric re-entry should be carefully considered in planning uncontrolled re-entry, particularly
for large spacecraft.

To assess the  human casualty risk of impact by  objects that survive re-entry, assessment
parameters and  their  allowable levels,   reliable analysis tools for survivability, and acceptable
analysis conditions should be developed.

Practice (Assessment of Re-entry Safety):

More than 4,300 missions to Earth orbit (more than 5,000 tons in mass) have been
accomplished since 1957. More than 50 large objects (system level objects) typically fall back
to Earth every year.

The re-entries of Cosmos 954 on Canadian territory in January 1978 and Skylab in the oceans

and on Australia in July 1979 are well-known. Large objects that have re-entered since the
1980’s are listed in the following table.

Table 5.3.2-1 Large objects re-entered after 1980

Name

Nationality Mass [kg]

Date of Decay

Mode

Salyut 6/Cosmos 1267

Russia

35,000

29-Jul-82

Controlled Re-entry

Cosmos 1443

Russia

15,000

19-Sep-83

Controlled Re-entry

Apollo 9 CSM BP-16

USA

16,700

10-Jul-85

Natural Re-entry

Apollo 8 CSM BP-26

USA

16,700

8-Jul-89

Natural Re-entry

Salyut 7/Cosmos 1686

Russia

40,000

7-Feb-91

Natural Re-entry

Compton GRO

USA

14,910

4-Jun-00

Controlled Re-entry

Mir

Russia

120,000

23-Mar -01

Controlled Re-entry

[URLhttp://www.aero.org/cords/faq3.html]

Typical parameters to assess re-entry safety are casualty area and the casualty expectation
(Ec). An allowable Ec is not currently recommended in the IADC Guidelines, while NASA Safety
Standard 1740.14

[2]

, U.S. Government Orbital Debris Mitigation Standard Practices

[6]

, and

NASDA Space Debris Mitigation Standard (NASDA-STD18A)

[3]

limit the value of Ec to less than

-4

[persons per event].

5.3.3 Other Orbits

Space systems that are terminating their operational phases in other orbital regions should be
manoeuvred to reduce their orbital lifetime,

commensurate with LEO lifetime limitations, or

relocated if they cause interference with highly utilised orbit regions.

Purpose

The Navigation Satellites orbital region, particularly the circular 12-hour-orbit, is also a useful

orbital region, which may be subject of further studies, although the number of satellites
currently residing in that orbital region is not yet large with respect to protected orbital regions.

5.4 Prevention of On-Orbit Collisions

In developing the design and mission profile of a space system, a program or project  should
estimate and limit the probability of accidental collision with known objects during the system's
orbital lifetime. If reliable  orbital data  is available, avoidance manoeuvres for spacecraft and  co-
ordination of launch windows may be considered if the  collision risk is not considered negligible.
Spacecraft design should limit the probability of collision with small debris which could cause a loss
of control, thus preventing post-mission disposal.

Purpose:

The above recommendation addresses

(1) estimation of collision probability and taking measures, if necessary, in the planning phase;

(2)

collisions with large objects during mission operations (collision avoidance); and

[This may be applied for large debris or orbiting vehicles (already tracked), and by an
operational action (authorisation for launcher lift-off, collision avoidance manoeuvre). Such
measures are already in place for some manned and unmanned spacecraft.]

(3) collision with small debris during mission operations.

[This may be applied for small or very small debris (on the order of 1mm) with additional
satellite shielding, a specific lay-out to protect the most sensitive components, or a
separation of redundant components]

Practice (avoidance of on-orbit collision):

The United States Space Surveillance Network (SSN) and the Russian Space Surveillance
System (SSS) monitor the LEO environment to warn crewed spacecraft if an object is projected
to come within a few kilometres. NASA computes for the Space Shuttle a probability of collision
with a conjuncting object and if the probability of collision is high enough (typically 1/10000), an
avoidance manoeuvre may be performed. During the period 1999-2003, the International Space
Station executed seven evasive manoeuvres using similar conjunction assessment techniques.
The Russian SSS and the Russian Space Agency performed similar collision avoidance
assessments for the Mir space station.

For GEO spacecraft, coordinated stationkeeping is beneficial. Inclination and eccentricity vector
separation strategies can be efficiently employed to maintain co-located GEO spacecraft at safe
distances. Eccentricity vector control may also be employed to reduce the risk of collision
between members of a given LEO satellite constellation.

Practice (avoidance of collision with new launch):

Collision between an ascending launch vehicle and manned systems should be avoided. In the
USA, collision avoidance analysis for new launches is conducted and safe launch windows are
established. In the event of a predicted conjunction, the launch is delayed. JAXA estimates the
collision probability with manned systems before H-IIA lift-off and confirms that they can keep a
distance of 200 km x 50 km x 50 km from any manned systems in orbit.