ISO 27852:2011

Space systems — Estimation of orbit lifetime

ISO 27852:2011 describes a process for the estimation of orbit lifetime for satellites, launch vehicles, upper stages and associated debris in LEO-crossing orbits. ISO 27852:2011 also clarifies the modelling approaches and resources for solar and geomagnetic activity modelling; the resources for atmosphere model selection; and an approach for satellite ballistic coefficient estimation.

Systèmes spatiaux — Estimation de la durée de vie en orbite

General Information

• Status: Withdrawn
• Publication Date: 12-Jul-2011
• Withdrawal Date: 12-Jul-2011
• ICS: 49.140 - Space systems and operations
• Technical Committee: ISO/TC 20/SC 14 - Space systems and operations
• Drafting Committee: ISO/TC 20/SC 14/WG 3 - Operations and support systems
• Current Stage: 9599 - Withdrawal of International Standard
• Completion Date: 04-Jul-2016

Relations

• Revised: ISO 27852:2016 - Space systems — Estimation of orbit lifetime
• Effective Date: 05-Nov-2015

Standards Content (Sample)

INTERNATIONAL ISO
STANDARD 27852
First edition
2011-07-15

Space systems — Estimation of orbit
lifetime
Systèmes spatiaux — Estimation de la durée de vie en orbite




Reference number
ISO 27852:2011(E)
©
ISO 2011

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ISO 27852:2011(E)


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ISO 27852:2011(E)
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INTERNATIONAL ISO
STANDARD 27852
First edition
2011-07-15

Space systems — Estimation of orbit
lifetime
Systèmes spatiaux — Estimation de la durée de vie en orbite




Reference number
ISO 27852:2011(E)
©
ISO 2011

---------------------- Page: 1 ----------------------
ISO 27852:2011(E)

COPYRIGHT PROTECTED DOCUMENT


©  ISO 2011
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Published in Switzerland

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ISO 27852:2011(E)
Contents Page
Foreword .iv
Introduction.v
1 Scope.1
2 Terms, definitions, symbols and abbreviated terms .1
2.1 Terms and definitions .1
2.2 Symbols.4
2.3 Abbreviated terms .4
3 Orbit lifetime estimation .4
3.1 General requirements .4
3.2 Definition of orbit lifetime estimation process.5
4 Orbit lifetime estimation methods and applicability.5
4.1 General .5
4.2 Method 1 — High-precision numerical integration .6
4.3 Method 2 — Rapid semi-analytical orbit propagation .6
4.4 Methods 3 — Numerical table look-up, analysis and fit equation evaluations.7
4.5 Orbit lifetime sensitivity to sun-synchronous and high-eccentricity orbits .7
5 Atmospheric density modelling.7
5.1 General .7
5.2 Atmospheric drag models .7
5.3 Long-duration solar flux and geomagnetic indices prediction .9
5.4 Atmospheric density implications of thermospheric global cooling.14
6 Estimating ballistic coefficient.14
6.1 General .14
6.2 Estimating drag coefficient .15
6.3 Estimating cross-sectional area with tumbling and stabilization modes.15
6.4 Estimating mass .16
Annex A (informative) Space population distribution.17
Annex B (informative) 25-Year lifetime predictions using random draw approach.20
Annex C (informative) Solar radiation pressure and 3rd-body perturbations .25
Bibliography.27

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ISO 27852:2011(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.
International Standards are drafted in accordance with the rules given in the ISO/IEC Directives, Part 2.
The main task of technical committees is to prepare International Standards. Draft International Standards
adopted by the technical committees are circulated to the member bodies for voting. Publication as an
International Standard requires approval by at least 75 % of the member bodies casting a vote.
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.
ISO 27852 was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles, Subcommittee
SC 14, Space systems and operations.

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ISO 27852:2011(E)
Introduction
A spacecraft is exposed to the risk of collision with orbital debris and operational satellites throughout its
launch, early orbit and mission phases. This risk is especially high during passage through or operations
within the LEO region.
To address these concerns, the Inter-Agency Space Debris Coordination Committee (IADC) recommended to
[2]
the United Nations (section 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.” and “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.”
The Scientific and Technical Subcommittee (STSC) of the United Nations Committee on the Peaceful Uses of
Outer Space (UNCOPUOS), acknowledging the benefits of the IADC guidelines, established the Working
[3]
Group on Space Debris to develop a set of recommended guidelines based on the technical content and the
basic definitions of the IADC space debris mitigation guidelines, taking into consideration the United Nations
treaties and principles on outer space. Consistent with the IADC recommendations (listed above), STSC
Guideline 6 states that space mission planners, designers, manufacturers and operators should “Limit the
long-term presence of spacecraft and launch vehicle orbital stages in the low-Earth orbit (LEO) region after
the end of their mission.” STSC guidelines also state, “For more in-depth descriptions and recommendations
pertaining to space debris mitigation measures, Member States and international organizations can refer to
the latest version of the IADC space debris mitigation guidelines and other supporting documents, which can
be found on the IADC website (www.iadc-online.org).”
The purpose of this International Standard is to provide a common, consensus approach to determining orbit
lifetime, one that is sufficiently precise and easily implemented for the purpose of demonstrating compliance
with IADC guidelines. This International Standard offers standardized guidance and analysis methods to
estimate orbital lifetime for all LEO-crossing orbit classes.
[1]
This International Standard is a supporting document to ISO 24113 and the GEO and LEO disposal
standards that are derived from ISO 24113.

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INTERNATIONAL STANDARD ISO 27852:2011(E)

Space systems — Estimation of orbit lifetime
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 International Standard describes a process for the estimation of orbit lifetime for satellites, launch
vehicles, upper stages and associated debris in LEO-crossing orbits.
It also clarifies the following:
⎯ modelling approaches and resources for solar and geomagnetic activity modelling;
⎯ resources for atmosphere model selection;
⎯ approaches for satellite ballistic coefficient estimation.
2 Terms, definitions, symbols and abbreviated terms
2.1 Terms and definitions
For the purposes of this document, the following terms and definitions apply.
2.1.1
orbit lifetime
elapsed time between the orbiting satellite's initial or reference position and orbit demise/reentry
NOTE 1 An example of the orbiting satellite's reference position is the post-mission orbit.
NOTE 2 The orbit's decay is typically represented by the reduction in perigee and apogee altitudes (or radii) as shown
in Figure 1.
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ISO 27852:2011(E)
Perigee and Apogee Height vs Time
(β=860.72cm^/kg, 300 x 1250 km, i=70° Jac71 F10=75,81,Ap=5)
1 400
SPIN_Hp(km)
SPIN_Ha(km)
OPUS_Hp(km)
1 200
OPUS_Ha(km)
1 000
800
600
400
200
0
02/22/08 06/01/08 09/09/08 12/18/08 03/28/09 07/06/09 10/14/09 01/22/10 05/02/10
Time (GMT)

Figure 1 — Sample of orbit lifetime decay profile
2.1.2
disposal phase
interval during which a spacecraft or launch vehicle orbital stage completes its disposal actions
2.1.3
earth equatorial radius
equatorial radius of the Earth
NOTE The equatorial radius of the Earth is taken as 6 378,137 km and this radius is used as the reference for the
Earth's surface from which the orbit regions are defined.
2.1.4
LEO-crossing orbit
low-earth orbit, defined as an orbit with perigee altitude of 2 000 km or less
NOTE As can be seen in Figure A.1, orbits having this definition encompass the majority of the high spatial density
spike of satellites and space debris.
2.1.5
long-duration orbit lifetime prediction
orbit lifetime prediction spanning two solar cycles or more (e.g. 25-year orbit lifetime)
2.1.6
mission phase
phase where the space system fulfills its mission
NOTE Begins at the end of the launch phase and ends at the beginning of the disposal phase.
2.1.7
post-mission orbit lifetime
duration of the orbit after completion of the mission phase
NOTE The disposal phase duration is a component of post-mission duration.
2 © ISO 2011 – All rights reserved

Apsidal Heights (km)

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ISO 27852:2011(E)
2.1.8
satellite
system designed to perform specific tasks or functions in outer space
NOTE A spacecraft that can no longer fulfill its intended mission is considered as non-functional. Spacecraft in
reserve or standby modes awaiting possible reactivation are considered functional.
2.1.9
space debris
all man-made objects, including fragments and elements thereof, in Earth orbit or re-entering the atmosphere,
that are non-functional
2.1.10
space object
man-made object in outer space
2.1.11
orbit
path followed by a space object
2.1.12
solar cycle
≈11-year solar cycle based on the 13-month running mean for monthly sunspot number and is highly
correlated with the 13-month running mean for monthly solar radio flux measurements at the 10,7 cm
wavelength
NOTE 1 Historical records back to the earliest recorded data (1947) are shown in Figure 2.
NOTE 2 For reference, the current 25-year post-mission IADC orbit lifetime recommendation is overlaid onto the
historical data; it can be seen that multiple solar cycles are encapsulated by this long time duration.
Adjusted Daily Ottawa/Penticton Solar Flux (10,7 cm wavelength)
350
25-year IADC Lifetime Recommendation
300
250
200
150
100
50
1947 1957 1967 1977 1987 1997
Date

Figure 2 — Solar cycle (≈11 year duration)
© ISO 2011 – All rights reserved 3

F -cm Solar Radio Flux
10,7

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ISO 27852:2011(E)
2.2 Symbols
a orbit semi-major axis
A satellite cross-sectional area with respect to the relative wind
A earth daily geomagnetic index
p
β ballistic coefficient of satellite, equal to C × A/m

D
C satellite drag coefficient
D
C satellite reflectivity coefficient
R
e orbit eccentricity
−22 −2 −1
F solar radio flux observed daily at 2 800 MHz (10,7 cm) in solar flux units (10 W m Hz )
10,7
F solar radio flux at 2 800 MHz (10,7 cm), averaged over three solar rotations
10,7
H apogee altitude, equal to a(1 + e) − R
a e
H perigee altitude, equal to a(1 − e) − R

p e
m mass of a satellite
R equatorial radius of the earth
e
2.3 Abbreviated terms
GEO geosynchronous earth orbit
GTO geosynchronous transfer orbit
IADC Inter-Agency Space Debris Coordination Committee
ISO International Organization for Standardization
LEO low-earth orbit
RAAN orbit right ascension of the ascending node (the angle between the vernal equinox and the
orbit ascending node, measured CCW in the equatorial plane, looking in the –Z direction)
STSC Scientific and Technical Subcommittee of the Committee
UNCOPUOS United Nations Committee on the Peaceful Uses of Outer Space
3 Orbit lifetime estimation
3.1 General requirements
The orbital lifetime of LEO-crossing mission-related objects shall be estimated using the processes specified
in this International Standard. In addition to any user-imposed constraints, the post-mission portion of the
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ISO 27852:2011(E)
resulting orbit lifetime estimate shall then be constrained to a maximum of 25 years as per IADC
recommendations using a combination of the following:
a) initial orbit selection;
b) satellite vehicle design;
c) spacecraft launch and early orbit concepts of operation which minimize LEO-crossing objects;
d) satellite ballistic parameter modifications at EOL;
e) satellite deorbit manoeuvres.
3.2 Definition of orbit lifetime estimation process
The orbit lifetime estimation process is represented generically in Figure 3.
Atmosphere Model
Method 1 or 2: Orbit
Anticipated Long-
Integration and
Duration Solar &
Propagation Package

Geomagnetic Activity

Orbit Lifetime
Estimate

Spacecraft Orbit Initial
Conditions at Epoch

Method 3: Graphical or
Tabular Look-up

Spacecraft Ballistic
Characteristics, Attitude
Rules


[4]
Figure 3 — Orbit lifetime estimation process
4 Orbit lifetime estimation methods and applicability
4.1 General
[4]
There are three basic analysis methods used to estimate orbit lifetime , as depicted in Figure 3.
Determination of the method used to estimate orbital lifetime for a specific space object shall be based upon
the orbit type and perturbations experienced by the satellite as shown in Table 1.
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ISO 27852:2011(E)
Table 1 — Applicable method with mandated conservative margins of error
and required perturbation modelling
Orbit apogee Special orbit: Conservative margin applied to each method:
altitude
Sun- High Method 1: Method 2: Method 3a: Method 3b:
sync? area-to-
km Numerical Semi-analytic Table look-up Graph,
mass?
integration equation fit
Apogee < 2 000 No No No margin req'd 5 % margin 10 % margin 25 % margin
Apogee < 2 000 No Yes No margin; use SRP 5 % margin; use 10 % margin IFF N/A
SRP
C ≈ 1,7
r
Yes No No margin req'd 5 % margin N/A N/A
Apogee < 2 000
Yes Yes No margin req'd; use 5 % margin; use N/A N/A
Apogee < 2 000
SRP SRP
Apogee > 2000 Either Either No margin req'd; use 5 % margin; use N/A N/A
3Bdy+SRP 3 Bdy + SRP
N/A = not applicable
3Bdy = third-body perturbations

SRP = solar radiation pressure

Method 1, certainly the highest fidelity model, utilizes a numerical integrator with a detailed gravity model,
third-body effects, solar radiation pressure, and a detailed satellite ballistic coefficient model. Method 2 utilizes
[5],[6]
a definition of mean orbital elements , semi-analytic orbit theory and average satellite ballistic coefficient to
permit a very rapid integration of the equations of motion, while still retaining reasonable accuracy.
methods 3a and 3b are simply a table lookup, graphical analysis or evaluation of equations fit to pre-computed
orbit lifetime estimation data obtained via the extensive and repetitive application of Methods 1 and/or 2.
4.2 Method 1 — High-precision numerical integration
Method 1 is the direct numerical integration of all accelerations in Cartesian space, with the ability to
incorporate a detailed gravity model (e.g. using a larger spherical harmonics model to address resonance
effects), third-body effects, solar radiation pressure, vehicle attitude rules or aero-torque-driven attitude
torques, and a detailed satellite ballistic coefficient model based on the variation of the angle-of-attack with
respect to the relative wind. Atmospheric rotation at the earth's rotational rate is also easily incorporated in this
approach. The only negative aspects to such simulations are the following.
a) They run much slower than Method 2.
b) Many of the detailed data inputs required to make this method realize its full accuracy potential are simply
unavailable.
c) Any gains in orbit lifetime prediction accuracy are frequently overwhelmed by inherent inaccuracies of
atmospheric modelling and associated inaccuracies of long-term solar activity predictions/estimates.
However, to analyse a few select cases where such detailed model inputs are known, this is undoubtedly the
most accurate method. At a minimum, Method 1 orbit lifetime estimations shall account for J and J
2 3
perturbations and drag using an accepted atmosphere model and an average ballistic coefficient. In the case
of high apogee orbits (e.g. geosynchronous transfer orbits), sun and moon third-body perturbations shall also
be modelled.
4.3 Method 2 — Rapid semi-analytical orbit propagation
[5],[6]
Method 2 analysis tools utilize semi-analytic propagation of mean orbit elements influenced by gravity
zonals J and J and selected atmosphere models. The primary advantage of this approach over direct
2 3
numerical integration of the equations of motion (Method 1) is that long-duration orbit lifetime cases can be
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ISO 27852:2011(E)
quickly analysed (e.g. 1 s versus 1 700 s CPU time for a 30 year orbit lifetime case). While incorporation of an
attitude-dependent ballistic coefficient is possible for this method, an average ballistic coefficient is typically
used. At a minimum, Method 2 orbit lifetime estimations shall account for J and J perturbations and drag
2 3
using an accepted atmosphere model and an average ballistic coefficient. In the case of high-apogee orbits
(e.g. GTO), sun and moon third-body perturbations shall also be modelled.
4.4 Methods 3 — Numerical table look-up, analysis and fit equation evaluations
In these methods, one uses tables, graphs and equations representing data that was generated by
exhaustively using Methods 1 and 2 (see 4.2 and 4.3). The graphs and equations provided in this International
Standard can help the analyst crudely estimate orbit lifetime for their particular case of interest; the electronic
1)
access to tabular look-up provided via this International Standard permits the analyst to estimate orbit
lifetime for their particular case of interest via interpolation of Method 1 or Method 2 gridded data; all such
Methods 3 data in this paper were generated using Method 2 approaches. At a minimum, method 3 orbit
lifetime products shall be derived from Method 1 or Method 2 analysis products meeting the requirements
stated above. When using this method, the analyst shall impose at least a 10 % margin of error to account for
table look-up interpolation errors. When using graphs and equations, the analyst shall impose a 25 % margin
of error.
4.5 Orbit lifetime sensitivity to sun-synchronous and high-eccentricity orbits
For sun-synchronous orbits, orbit lifetime has some sensitivity to the initial value of RAAN due to the density
variations with the local sun angle. Results from numerous orbit lifetime estimations show that orbits with
[4]
6:00 am local time have longer lifetime than orbits with 12:00 noon local time by about 5,5 % . This maximum
difference (500 days) translates into a 5 % error, which can be corrected by knowing the local time of the orbit.
As a result, Method 1 or Method 2 analyses of the actual sun-synchronous orbit condition shall be used when
estimating the lifetime of sun-synchronous orbits.
For high-eccentricity orbits, it has been found difficult to iterate to lifetime threshold constraints due to the
coupling in eccentricity between the third-body perturbations and the drag decay. Due to this convergence
difficulty, only Method 1 or Method 2 analyses shall be used when determining initial conditions that achieve a
specified lifetime threshold for such orbits.
5 Atmospheric density modelling
5.1 General
The three biggest factors in orbit lifetime estimation are the following:
a) selection of an appropriate atmosphere model to incorporate into the orbit acceleration formulation;
b) selection of appropriate atmosphere model inputs;
c) determination of a space object's ballistic coefficient.
These three aspects are discussed in 5.2 to 5.4 and Clause 6.
5.2 Atmospheric drag models
There are a wide variety of atmosphere models available to the orbit analyst. The background, technical basis,
utility and functionality of these atmosphere models are described in detail in References [7] to [16]. This
International Standard does not presume to dictate which atmosphere model the analyst shall use. However, it
is worth noting that in general, the heritage, expertise and especially the observational data that went into

1) See www.CelesTrak.com.
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ISO 27852:2011(E)
creating each atmosphere model play a key role in that model's ability to predict atmospheric density, which is,
in turn, a key factor in estimating orbit lifetime. Many of the early atmosphere models were low fidelity and
were created on the basis of only one, or perhaps even just a part of one, solar cycle's worth of data.
The advantage of some of these early models is that they typically run much faster than the latest high-fidelity
models (Table 2), without a significant loss of accuracy. However, the use of atmosphere models that were
designed to fit a select altitude range (e.g. the “exponential” atmosphere model depicted in Table 2) or models
that do not accommodate solar activity variations should be avoided as they miss too much of the atmospheric
density variations to be sufficiently accurate.
There are some early models (e.g. Jacchia 1971 shown in Figure 5) which accommodate solar activity
variations and also run very fast; these models can work well for long-duration orbit lifetime studies where
numerous cases are to be examined. Conversely, use of the more recent atmosphere models are encouraged
because they have substantially more atmospheric drag data incorporated as the foundation of their
underlying assumptions. A crude comparison of a sampling of atmosphere models for a single test case is
shown in Figures 4 and 5, illustrating the range of temperatures and densities exhibited by the various models.
Although this International Standard does not presume to direct which atmosphere model the analyst should
use, the reader is encouraged to seek atmosphere model guidance from existing and upcoming ISO
International Standards and CIRA Working Group (e.g. CIRA-2008) recommendations. Models worthy of
[11] [12] [13] [14]
consideration include, but are not limited to, the NRLMSISE-00 , JB2006 , JB2008 , GRAM-07 ,
[15] [16]
DTM-2000 and GOST-2004 models.
Table 2 — Comparison of normalized density evaluation run times
Atmosphere model 0 < Altitude < 5 000 km 0 < Altitude < 1 000 km
Exponential 1,00 1,00
Atml962 1,43 1,51
Atm1976 1,54 1,54
Jacchia1971 13,68 17,31
MSIS 2000 141,08 222,81
JB2006 683,85 584,47

Figure 4 — Temperature comparison by atmosphere model
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ISO 27852:2011(E)
Log(Density) vs Altitude
0.00
0 200 400 600 800 1000 1200 1400 1600 1800 2000
-5.00
AtmExp log(Rho)
Atm62 log(Rho)
Atm76 log(Rho)
Jac71 log(Rho)
-10.00
MSIS00 log(Rho)
JB2006 log(Rho)
-15.00
-20.00
Altitude (km)

Figure 5 — Comparison of a small sampling of atmosphere models
5.3 Long-duration solar flux and geomagnetic indices prediction
Utilization of the higher-fidelity atmosphere models mentioned in 5.2 requires the orbit analyst to specify the
solar and geomagnetic indices required by such models. Care shall be taken to obtain the proper indices
required by each model; subtle differences can exist in the interpretation of similarly named indices when used
by different atmosphere models (e.g. centrally-averaged vs. backward-averaged F ).
10,7
Key issues associated with any prediction of solar and geomagnetic index modelling approach are the
following.
a) F predictions should reflect the mean solar cycle as accurately as possible.
10,7
b) Large daily F and A index variations about the mean value induce non-linear variations in
10,7 p
atmospheric density, and the selected prediction approach should account for this fact; i.e. one should
account for the highly non-linear aspects of solar storms versus quiet periods.
c) The frequency of occurrence across the day-to-day index values is highest near the lowest prediction
boundary (see Figure 7).
d) F cycle timing/phase are always imprecise and should be accounted for; the resultant time bias that
10,7
such a prediction error would introduce can yield large F prediction errors of 100 % or more.
10,7
e) Although still under review, the long-time duration currently being advocated by the IADC (i.e., 25 years)
would require that the solar/geomagnetic modelling approach provide at least that many years (i.e. 25) of
predictive capability.
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Log(Density in kg/m^3)

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ISO 27852:2011(E)
f) Predicted F values should be adjusted to correct for earth-sun distance variations.
10,7
g) Some atmosphere models (e.g. JB2006 and JB2008), due to the newly invented indices adopted thereby,
preclude the use of historical indices for long-term orbit lifetime studies, while currently also precluding
use of any predictive forecasting model(s) for those indices until such time as those become publicly
available.
Accounting for these constraints, the user shall adopt one of the two approaches:
[17][18]
⎯ Approach 1: Utilize Monte Carlo sampling of historical data mapped to a common solar cycle
period;
[19]
⎯ Approach 2: Utilize a predicted F solar activity profile generated by a model such as is detailed in
10,7
Figure 6, coupled with a stochastic or similar generation of corresponding F and A values, e.g.
10,7 p
Reference [20].
Since Approach 2 is a well-known and common approach, the focus of the remainder of this subclause is
[4]
devoted to the Monte Carlo “Random Draw” approach . Be aware (see Figure 2) that there are already more
than five solar cycles of observed solar and geomagnetic data to choose from. Processing of this data maps
each coupled and correlated triad of datum (F , F , a
...