Space systems -- Disposal of satellites operating at geosynchronous altitude

This document specifies requirements for the following: — planning for disposal of a spacecraft operating at geosynchronous altitude to ensure that final disposal is sufficiently characterized and that adequate propellant will be reserved for the manoeuvre; — selecting final disposal orbits where the spacecraft will not re-enter the operational region within the next 100 years; — executing the disposal manoeuvre successfully; — depleting all energy sources on board the vehicle before the end of its life to minimize the possibility of an event that can produce debris. This document provides techniques for planning and executing the disposal of space hardware that reflect current internationally accepted guidelines and consider current operational procedures and best practices.

Systèmes spatiaux -- Élimination des satellites opérant à une altitude géostionnaire

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Publication Date
02-Jul-2019
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9092 - International Standard to be revised
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INTERNATIONAL ISO
STANDARD 26872
Second edition
2019-07
Space systems — Disposal of satellites
operating at geosynchronous altitude
Systèmes spatiaux — Élimination des satellites opérant à une altitude
géostionnaire
Reference number
ISO 26872:2019(E)
ISO 2019
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ISO 26872:2019(E)
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© ISO 2019

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Published in Switzerland
ii © ISO 2019 – All rights reserved
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ISO 26872:2019(E)
Contents Page

Foreword ........................................................................................................................................................................................................................................iv

Introduction ..................................................................................................................................................................................................................................v

1 Scope ................................................................................................................................................................................................................................. 1

2 Normative references ...................................................................................................................................................................................... 1

3 Terms and definitions ..................................................................................................................................................................................... 1

4 Symbols and abbreviated terms ........................................................................................................................................................... 2

4.1 Symbols ......................................................................................................................................................................................................... 2

4.2 Abbreviated terms ............................................................................................................................................................................... 3

5 Geosynchronous region ................................................................................................................................................................................ 3

6 Protected region ................................................................................................................................................................................................... 3

7 Primary requirements.................................................................................................................................................................................... 5

7.1 Disposal manoeuvre planning ................................................................................................................................................... 5

7.2 Probability of successful disposal........................................................................................................................................... 6

7.3 Criteria for executing disposal action .................................................................................................................................. 6

7.4 Contingency planning ....................................................................................................................................................................... 6

8 Disposal planning requirements ......................................................................................................................................................... 6

8.1 General ........................................................................................................................................................................................................... 6

8.2 Estimating propellant reserves ................................................................................................................................................ 6

8.3 Computing the initial perigee increase .............................................................................................................................. 6

8.4 Developing basic manoeuvre requirements for a stable disposal orbit ................................................ 7

8.5 Developing long-term (100-year) disposal orbit characteristics ................................................................ 7

8.6 Determining the manoeuvre sequence .............................................................................................................................. 7

8.7 Developing a vehicle securing plan ....................................................................................................................................... 8

8.8 Developing a contingency plan ................................................................................................................................................. 8

Annex A (informative) Tabulated values of the optimal eccentricity vector...............................................................9

Annex B (informative) Optimal manoeuvre sequences .................................................................................................................27

Annex C (informative) Example calculations ............................................................................................................................................33

Annex D (informative) Disposal strategy and analysis for sample GEO spacecraft ..........................................39

Bibliography .............................................................................................................................................................................................................................46

© ISO 2019 – All rights reserved iii
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ISO 26872:2019(E)
Foreword

ISO (the International Organization for Standardization) is a worldwide federation of national standards

bodies (ISO member bodies). The work of preparing International Standards is normally carried out

through ISO technical committees. Each member body interested in a subject for which a technical

committee has been established has the right to be represented on that committee. International

organizations, governmental and non-governmental, in liaison with ISO, also take part in the work.

ISO collaborates closely with the International Electrotechnical Commission (IEC) on all matters of

electrotechnical standardization.

The procedures used to develop this document and those intended for its further maintenance are

described in the ISO/IEC Directives, Part 1. In particular, the different approval criteria needed for the

different types of ISO documents should be noted. This document was drafted in accordance with the

editorial rules of the ISO/IEC Directives, Part 2 (see www .iso .org/directives).

Attention is drawn to the possibility that some of the elements of this document may be the subject of

patent rights. ISO shall not be held responsible for identifying any or all such patent rights. Details of

any patent rights identified during the development of the document will be in the Introduction and/or

on the ISO list of patent declarations received (see www .iso .org/patents).

Any trade name used in this document is information given for the convenience of users and does not

constitute an endorsement.

For an explanation of the voluntary nature of standards, the meaning of ISO specific terms and

expressions related to conformity assessment, as well as information about ISO's adherence to the

World Trade Organization (WTO) principles in the Technical Barriers to Trade (TBT) see www .iso

.org/iso/foreword .html.

This document was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles,

Subcommittee SC 14, Space systems and operations.

This second edition cancels and replaces the first edition (ISO 26872:2010), which has been technically

revised. The main changes compared to the previous edition are as follows:

— to be consistent with ISO 24113, the word “satellite” has been replaced by “spacecraft”;

— ISO 24113 has been incorporated by reference, such that its normative content serves as requirements

in this document as well;

— to be consistent with ISO 24113, Post-Mission Disposal is no longer defined as a conditional

probability.

Any feedback or questions on this document should be directed to the user’s national standards body. A

complete listing of these bodies can be found at www .iso .org/members .html.
iv © ISO 2019 – All rights reserved
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ISO 26872:2019(E)
Introduction

This document prescribes requirements for planning and executing manoeuvres and operations to

remove an operating spacecraft from the geosynchronous orbit at the end of its mission and place

it in an orbit for final disposal where it will not pose a future hazard to spacecraft operating in the

geosynchronous ring.
This document includes requirements related to the following:
— when the disposal action needs to be initiated,
— selecting the final disposal orbit,
— executing the disposal action successfully, and
— depleting all energy sources to prevent explosions after disposal.

End-of-mission disposal of an Earth-orbiting spacecraft broadly means the following:

a) removing the spacecraft from the region of space where other spacecrafts are operating, so as not

to interfere or collide with these other users of space in the future, and

b) ensuring that the disposed object is left in an inert state and is incapable of generating an explosive

event that could release debris which might threaten the operating spacecraft, see ISO 16127.

For a spacecraft operating in the geosynchronous belt, the most effective means of disposal is first to

re-orbit the spacecraft to a super-synchronous orbit above the region of the operating spacecraft and

the manoeuvre corridor used for relocating the operating spacecraft to new longitudinal slots, and then

to discharge batteries and vent propellants and take other actions to preclude a debris-producing event.

© ISO 2019 – All rights reserved v
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INTERNATIONAL STANDARD ISO 26872:2019(E)
Space systems — Disposal of satellites operating at
geosynchronous altitude

IMPORTANT — The electronic file of this document contains colours which are considered to be

useful for the correct understanding of the document. Users should therefore consider printing

this document using a colour printer.
1 Scope
This document specifies requirements for the following:

— planning for disposal of a spacecraft operating at geosynchronous altitude to ensure that final

disposal is sufficiently characterized and that adequate propellant will be reserved for the

manoeuvre;

— selecting final disposal orbits where the spacecraft will not re-enter the operational region within

the next 100 years;
— executing the disposal manoeuvre successfully;

— depleting all energy sources on board the vehicle before the end of its life to minimize the possibility

of an event that can produce debris.

This document provides techniques for planning and executing the disposal of space hardware that

reflect current internationally accepted guidelines and consider current operational procedures and

best practices.
2 Normative references

The following documents are referred to in the text in such a way that some or all of their content

constitutes requirements of this document. For dated references, only the edition cited applies. For

undated references, the latest edition of the referenced document (including any amendments) applies.

ISO 24113:2019, Space systems — Space debris mitigation requirements
3 Terms and definitions

For the purposes of this document, the terms and definitions given in ISO 24113 and the following apply.

ISO and IEC maintain terminological databases for use in standardization at the following addresses:

— ISO Online browsing platform: available at https: //www .iso .org/obp
— IEC Electropedia: available at http: //www .electropedia .org/
3.1
inclination excursion region

region in space occupied either by a non-operational geostationary spacecraft (3.4) or by an operational

geosynchronous spacecraft without inclination station-keeping
3.2
re-orbit manoeuvre
action of moving a spacecraft (3.4) to a new orbit
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ISO 26872:2019(E)
3.3
satellite

manufactured object or vehicle intended to orbit the Earth, the moon or another celestial body

3.4
spacecraft

system designed to perform a set of tasks or functions in outer space, excluding launch vehicle

[SOURCE: ISO 24113:2019, 3.25]
4 Symbols and abbreviated terms
4.1 Symbols
a semi-major axis
C solar radiation pressure coefficient of the spacecraft (0 < C < 2)
R R
NOTE In some references, C is defined as the index of surface reflection.
e eccentricity
h perigee altitude
i inclination
I specific impulse
L solar longitude
M mean anomaly
p semilatus rectum or semi-parameter [p = a(1 − e )]
r radius of orbit
v true anomaly
μ Earth gravitational constant

σ standard deviation or the positive root of the variance, which measures the dispersion of

the data
Ω right ascension of ascending node (RAAN)
ω argument of perigee

A/m effective area-to-mass ratio: projected area of the spacecraft perpendicular to the sun's ray

divided by the mass of the spacecraft
ΔH change in altitude
ΔV delta velocity or total velocity change
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ISO 26872:2019(E)
4.2 Abbreviated terms
EGM Earth gravitational model
EOMDP end-of-mission disposal plan
GEO geosynchronous (geostationary) Earth orbit
RAAN right ascension of ascending node
5 Geosynchronous region

The geosynchronous region is a circular ring around the Earth in the equatorial plane. Within this region,

an object in space moves along the ring at a mean angular rate that is equal or very close to the Earth's

rotation, meaning that the spacecraft appears to be positioned over a fixed location on the ground.

Without so-called north-south station-keeping, the inclination of a GEO spacecraft will gradually cycle

between 0° (equatorial orbit) and a maximum of approximately 14,6° and back again. In addition to

maintaining the accuracy of its inclination, a GEO spacecraft must execute station-keeping manoeuvres

to maintain longitudinal accuracy, so as to prevent a naturally occurring drift to the east or to the west

caused by asymmetries in the Earth’s gravitational field, unless the spacecraft is located at one of the

two “gravity wells” on the geostationary arc.

Figure 1 shows a three-dimensional view of the geosynchronous ring with a cross-section defining the

approximate size of the ring. Figure 2 gives the dimensions of three regions of the cross-section. The

cross-section is defined by two axes: the latitudinal axis and radial axis. This plane of the cross-section

is perpendicular to the Earth's equatorial plane.

The three concentric boxes shown in Figure 2 give the approximate boundaries for three types of

orbits. The smallest box represents the region where a geostationary spacecraft will be confined under

station-keeping, and the next larger box approximates the region where a geosynchronous spacecraft

may be located when its inclination is not controlled but it remains under a mission-specified value. For

example, the upper value for some specific geosynchronous missions may range from 3° to 5° depending

on the ground user's antenna design. The largest box represents the inclination excursion region

for a non-operational GEO spacecraft and the ±200 km protected region. For most communication

spacecrafts, the longitude station-keeping limit is ±0,1°.
6 Protected region

The GEO protected region, defined by ISO 24113 and indicated by 3 in Figure 1, includes the rectangular

toroid centred on geostationary altitude, with an extent 200 km above and below this altitude and

with inclination limits of +15° to −15°. While operations are usually conducted within about 75 km

of geostationary altitude, the GEO protected region is extended in altitude to create a manoeuvre

corridor for relocating the spacecraft. Passivation of the disposed spacecraft is necessary to ensure

that accidental explosions from on-board energy sources do not create debris that could re-enter the

protected region.
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ISO 26872:2019(E)
Key
1 Earth
2 equator
3 GEO region
4 LEO (low Earth orbit) region
Z altitude measured with respect to a spherical Earth whose radius is 6 378 km

Z altitude of the geostationary orbit with respect to a spherical Earth whose radius is 6 378 km

GEO
NOTE The dimensions in the figure are not to scale.
Figure 1 — View in the equatorial plane of Earth and the protected regions
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ISO 26872:2019(E)
Dimensions in kilometres
Key
x radial (away from Earth)
y latitude (north)
1 protected region
2 geostationary control box (±37,5 km × ±37,5 km)
3 geosynchronous control box (±40 km radial; ±3° to ±5° in inclination)
NOTE The dimensions in the figure are not to scale.
Figure 2 — Cross-section of the geosynchronous ring
7 Primary requirements
7.1 Disposal manoeuvre planning

An EOMDP shall be developed, maintained and updated in all phases of mission and spacecraft design

and operation. The EOMDP shall be an integral part of the space debris mitigation plan specified by

ISO 24113. The EOMDP shall include the following:
a) details of the nominal mission orbit;
b) details of the targeted disposal orbit;
c) estimates of the propellant required for the disposal action;

d) identity of systems and capabilities required for successful completion of the disposal action;

e) criteria that, when met, shall dictate initiation of the disposal action;
f) identities of energy sources required to be depleted before end of life;
g) timeline for initiating and executing the disposal action;
h) timeline for depleting the remaining energy sources;
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ISO 26872:2019(E)

i) those individuals or entities, or individuals and entities to be notified of the end of mission and

disposal and a timeline for notification.
7.2 Probability of successful disposal

In accordance with the requirements of ISO 24113:2019, 6.3.1, a spacecraft shall be designed such that

the joint probability of having sufficient energy (propellant) remaining to achieve the final disposal

orbit and successfully executing commands to deplete energy sources equals or exceeds 0,9 at the time

disposal is executed. Details of the design that provide the basis for the probability estimate shall be

included in the EOMDP.
7.3 Criteria for executing disposal action

Specific criteria for initiating the disposal action shall be developed, included in the EOMDP and

monitored throughout the mission life.

EXAMPLE Propellant amount remaining; redundancy remaining; status of electrical power; status of

systems critical to a successful disposal action; time required to execute disposal action.

Projections of mission life based on these criteria shall be made as a regular part of mission status

reviews.
7.4 Contingency planning

Independent of the success or failure of other aspects of a disposal action, a contingency plan shall be

developed to deplete all energy sources and secure the vehicle before the final demise of the spacecraft.

The objective shall be to ensure that actions necessary to secure the vehicle are taken before end of life.

The contingency plan shall include criteria that define when the securing actions are to be taken, the

rationale for each criterion, and a schedule for securing actions. The contingency plan shall be included

in the EOMDP.
8 Disposal planning requirements
8.1 General

Planning activities for end-of-mission disposal shall start in the mission design phase. Planning for the

actual disposal action should begin at least six months before the date of re-orbit manoeuvres. The steps

described in 8.2 to 8.8 shall be followed in all mission phases and shall be documented in the EOMDP.

8.2 Estimating propellant reserves

The amount of fuel necessary to perform spacecraft disposal shall be estimated from the design phase,

in accordance with the needed accuracy level, and reserved for the disposal phase. The minimum ΔV

capability (3 − σ) to reach the targeted disposal orbit shall be determined and specified in the EOMDP.

The fuel required to provide this ΔV shall be maintained for end-of-life disposal, see ISO 23339.

8.3 Computing the initial perigee increase

In accordance with the requirements of ISO 24113:2019, 6.3.2, a spacecraft operating within the GEO

protected region shall, after completion of its GEO disposal manoeuvres, have an orbital state that

satisfies at least one of the two conditions outlined below.

a) The orbit has an initial eccentricity of less than 0,003, and a minimum perigee altitude, ΔH,

expressed in kilometres, above the geostationary altitude (35 786 km) calculated according to

Formula (1):
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ISO 26872:2019(E)
Δ HC=+235 1 000××Am/ (1)

The minimum value of C for computing the initial perigee increase shall be no less than 1,5 (a

conservative estimate for C , so as to adequately predict the solar radiation pressure effect).

Justification shall be provided for using a value less than 1,5. Formula (1) was derived to ensure

that the long-term perturbations will not cause the GEO debris to re-enter a protected zone of GEO

plus 200 km.

b) The orbit has a perigee altitude sufficiently above the geostationary altitude that the spacecraft will

not enter the GEO protected region within 100 years, irrespective of long-term perturbation forces.

8.4 Developing basic manoeuvre requirements for a stable disposal orbit

A stable disposal orbit shall be established by one of the two options described below.

a) Use Formula (1) and the eccentricity constraint to determine initial disposal orbit conditions.

b) Perform long-term (100-year) numerical integrations of the selected disposal orbit. The predicted

minimum perigee altitude shall be greater than the 200 km protected region (see 8.5). It is

recommended that the optimal eccentricity vector be determined from Tables A.1 to A.3, as a

function of the date of orbital insertion and the value of C × A/m.

The altitude stability will be improved for either method if the following apply:

— the initial disposal perigee points toward the sun (perigee is sun-pointing);

— the disposal manoeuvres are performed in the most favourable season of the year, such that the

same amount of perigee altitude increase will give the largest clearance over 100 years.

NOTE 1 The true optimal direction will differ slightly from the actual sun-pointing direction as a result of

lunar perturbations.

NOTE 2 See Annex A for the optimal eccentricity and argument of perigee as a function of time for various

values of C × A/m. Disposal orbits defined in accordance with Formula (1) are stable if the final eccentricity is

less than 0,003. Tables A.1 to A.3 can be used to select the initial guess if option b) is used to determine the initial

orbit parameters.

Should the intention be to operate the vehicle after placing it in a disposal orbit, the effects of such

operation on the orbit shall be estimated; and this estimate and computations verifying that the

operations will not compromise the long-term stability of the orbit (i.e. perigee shall remain above the

protected region for 100 years) shall be included in the EOMDP. In all cases, the spacecraft shall be

passivated (see 8.7) prior to end of life.
8.5 Developing long-term (100-year) disposal orbit characteristics

Long-term (100-year) orbit histories are needed only when the second option [see 8.4 b)] is chosen to

establish a stable disposal orbit. If 8.4 b) is chosen, orbit propagation results developed by a reliable

orbit propagator, either semi-analytic or numerical, shall be used to predict histories of perigee heights

above GEO for a period of 100 years after initial insertion into the disposal orbit. The orbit propagator

shall be of high precision and include as a minimum the perturbing forces of Earth's gravitational

harmonics (up to a degree/order of 6 by 6), lunisolar attractions and solar radiation pressure. The

precision of long-term propagation of the propagator shall be verified against another well-established

orbit propagator. Details on the orbit propagator used, assumptions made and analysis results shall be

included in the EOMDP.
8.6 Determining the manoeuvre sequence

The manoeuvre sequence shall be determined that will place the GEO spacecraft in the required disposal

orbit, have the optimal near-sun-pointing perigee and exhaust all the propellant on board. The disposal

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ISO 26872:2019(E)

orbit is obtained after passivation and complete tank depletion, which can have unpredictable effects

on orbital parameters and altitude. See Annex B for examples. The initial conditions of the disposal

orbit shall be determined using the steps outlined in 8.4 and 8.5.
8.7 Developing a vehicle securing plan

Depletion of propellant creates forces that can affect a vehicle's orbit. The vehicle securing plan shall

specify the following:

a) steps to deplete on-board energy sources after the spacecraft has been placed into the disposal orbit;

b) the effects the depletion action will have on the final orbit of the vehicle (the goal should be either

to increase altitude or at least to limit a possible decrease in altitude);
c) criteria for when the plan will be executed;
d) a schedule to be followed.
8.8 Developing a contingency plan

If a malfunction or other circumstance makes it necessary to proceed to the disposal phase earlier than

planned, a contingency plan shall be developed that includes provisions for the following:

a) selecting an alternative orbit that is the least likely to interfere with the protected area (see

Annex C): the contingency plan shall include criteria and techniques for selecting this orbit;

b) manoeuvring the spacecraft to the alternative orbit;
c) securing the spacecraft after the move;

d) securing the vehicle if specified criteria are met at any time in the mission.

Annex D provides an example in which the quantity of propellant remaining is uncertain.

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ISO 26872:2019(E)
Annex A
(informative)
Tabulated values of the optimal eccentricity vector

Tables A.1 to A.3 contain the optimal eccentricity vector [eccentricity and argument of perigee plus

RAAN (or longitude of periapsis)] as a function of time and a function of (C × A/m), expressed in

square metres per kilogram, that will yield the highest perigee over the next 100 years. The optimal

values were calculated in a brute-force fashion using increments of 2,3 × 10 in eccentricity and 5° in

longitude of periapsis. The benefit gained from using the optimal vector over the sun-pointing strategy

varied from 0 km to 20 km (the average was approximately 9 km). However, if the sun-pointing strategy

is chosen for the disposed vehicle, then the longitude of periapsis should be set equal to the value of

the solar longitude (depicted as L in Tables A.1 to A.3) with an eccentricity equal to 0,01 × C × A/m.

S R

These charts can be interpolated to find the optimal vector for any particular spacecraft at a given

time. However, the following should be noted when using these tables.

The initial conditions used to generate the data assumed a constant semi-major axis of 300 km above

GEO (i.e. a constant ΔV was used in the disposal), mean anomaly of 180° (i.e. the last burn occurs at

apogee, raising the perigee so that the eccentricity is equal to the tabulated value), an inclination of

7,74° (maximum at end of life if inclination drift is allowed) and an epoch of 0:00 Universal Time on

the first day of each month. Additional analysis has shown that the optimal vector depends little upon

these elements (the minimum perigee altitude may vary by approximately 2 km for each component),

but if a high level of accuracy is required for a given disposal, the interpolated values found from the

tables should be used as an initial guess so as to find the optimum for a particular disposal situation.

The exception is the RAAN: in the search process, the initial RAAN was set to 62,3° and the argument

of perigee was changed in 5° increments until the optimal valu
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