ISO 26872:2019

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Second edition
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Space systems —
Disposal of satellites
operating at
geosynchronous
altitude
Systèmes spatiaux —
Élimination des satellites
opérant à une altitude
géostionnaire

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Contents Page
Foreword .iii
Introduction . iv
1  Scope . 1
2  Normative references . 1
3  Terms and definitions . 1
4  Symbols and abbreviated terms . 2
5  Geosynchronous region . 3
6  Protected region . 3
7  Primary requirements . 4
8  Disposal planning requirements . 5
Annex A (informative) Tabulated values of the optimal eccentricity vector . 8
Annex B (informative) Optimal manoeuvre sequences . 28
B.1  General . 28
B.2  Type A manoeuvre sequence . 28
B.3  Type B manoeuvre sequence . 29
B.4  Sample manoeuvre sequence . 30
B.5  Manoeuvre sequence for low-thrust propulsion . 32
Annex C (informative) Example calculations . 34
C.1  Closed-form formulae for computing off-perigee burn locations . 34
C.2  Sample 100-year histories of sun-pointing disposal orbits . 35
Annex D (informative) Disposal strategy and analysis for sample GEO spacecraft . 40
Bibliography . 47

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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).
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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).
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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.
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d.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.
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ml
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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.
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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
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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
Deleted: ISO 24113:—
1
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debris mitigation
3 Terms and definitions
requirements¶
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
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p
— IEC Electropedia: available at http://www.electropedia.org/
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a.org/
3.1
inclination excursion region
region in space occupied either by a non‐operational geostationary spacecraft (3.4) or by an operational
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geosynchronous spacecraft without inclination station‐keeping
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3.2
re-orbit manoeuvre
action of moving a spacecraft (3.4) to a new orbit
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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
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[SOURCE: ISO 24113:2019, 3.25]
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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.
R
e eccentricity
h perigee altitude
p
i inclination
I specific impulse
sp
L solar longitude
S
M mean anomaly
p 2
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
4.2 Abbreviated terms
EGM Earth gravitational model
EOMDP end‐of‐mission disposal plan
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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
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ground.
Without so‐called north‐south station‐keeping, the inclination of a GEO spacecraft will gradually cycle
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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
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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
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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‐
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keeping, and the next larger box approximates the region where a geosynchronous spacecraft may be
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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
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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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Key
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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
ZGEO altitude of the geostationary orbit with respect to a spherical Earth whose radius is 6 378 km
NOTE The dimensions in the figure are not to scale.
Figure 1 — View in the equatorial plane of Earth and the protected regions
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;
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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;
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
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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.
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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
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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):
(1)
HC2351000 A/m
R
The minimum value of C for computing the initial perigee increase shall be no less than 1,5 (a
R
conservative estimate for C , so as to adequately predict the solar radiation pressure effect).
R
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.
R
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
R
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
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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
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disposal orbit, have the optimal near‐sun‐pointing perigee and exhaust all the propellant on board. The
disposal 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
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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;
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c) securing the spacecraft after the move;
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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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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
R
square metres per kilogram, that will yield the highest perigee over the next 100 years. The optimal
−5
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
S
0,01 × C × A/m. These charts can be interpolated to find the optimal vector for any particular
R
spacecraft at a given time. However, the following should be noted when using these tables.
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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 value was found. Different RAANs were then
checked and it was found that the relevant angular parameter was the argument of perigee plus RAAN;
if this value is held constant, then the results will again be consistent with 1 km to 2 km, irrespective of
the particular RAAN.
In addition, care should be taken if interpolating the values. In searching for optimal values in the
angular argument, it was found that, at times, there was not one pure maximum, but multiple local
maximums. As either the time or C × A/m advanced, the true maximum jumped from one peak to
R
2
another. For example, consider the 2008‐05‐01 disposal. A C × A/m of 0,015 m /kg has an optimal
R
2
eccentricity of 0,000 04 and a longitude of periapsis and 262,3°, whereas the C × A/m of 0,03 m /kg
R
has optimal values of 0,000 09 and 37,3°. Linearly interpolating would imply optimal values of
2
0,000 057 and 307,3° for a C × A/m of 0,02 m /kg. Instead, the C × A/m = 0,02 optimal value was
R R
actually 0,000 015 and 47,3°. In this case, the optimal point switched from one maximum to another,
and therefore the intermediate maximum would actually be close to one point or the other.
Therefore, when confronted with angular changes greater than 90°, it is recommended that
interpolation not be performed. Instead, the closer value should be used either directly or as a starting
point for a more refined search.
A few final comments on the general behaviour of the system are warranted. When the A/m was small
(C × A/m < 0,01), the optimal angle was pointed at the lunar apogee; when the A/m was large
R
(C × A/m > 0,03), the solar radiation pressure force became dominant and the optimal angle was
R
directed toward the sun.
2 2
Table A.1 — Optimal eccentricity vector for C × A/m = 0,00 m /kg, C × A/m = 0,005 m /kg and
R R
2
C × A/m = 0,01 m /kg
R
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2
C × A
...