ISO/TR 19473:2015

TECHNICAL ISO/TR
REPORT 19473
First edition
2015-09-01
Space systems — Best practices
for orbit elements at payload — LV
separation
Systèmes spatiaux — Meilleures pratiques pour les éléments en orbite
à charge utile — Séparation LV
Reference number
©
ISO 2015
© ISO 2015, Published in Switzerland
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Contents Page
Foreword .iv
Introduction .v
1 Scope . 1
2 Symbols and abbreviated terms . 1
2.1 Abbreviated terms . 1
2.2 Symbols . 2
3 Orbit elements and calculation conditions . 3
3.1 Orbit elements . 3
3.2 Data source . 5
3.2.1 General. 5
3.2.2 Guideline of correction about the external measurements . 5
3.2.3 External measurement data accuracy . 5
3.3 Coordinate systems and time systems . 5
3.3.1 Coordinate systems . 5
3.3.2 Time systems . 7
4 Calculation method of Keplerian elements . 7
4.1 Calculation method of orbit elements . 7
4.2 Transformation of other orbit elements . 9
4.2.1 Parameters of orbit size and shape . . 9
4.2.2 Parameters of orbit orientation .10
4.2.3 Parameters of satellite location .10
5 Calculation method of orbit elements error .10
Bibliography .12
Foreword
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The committee responsible for this document is ISO/TC 20, Aircraft and space vehicles, Subcommittee
SC 14, Space systems and operations.
iv © ISO 2015 – All rights reserved

Introduction
This Technical Report will provide a recommendatory method for post-launch assessment of the orbit
elements precision at separation, which is conductive to improving international communication effect
and reducing the risks from errors resulting from miscommunication. It can estimate the orbit elements
precision at separation, provide the reference for fuel capacity design of launch vehicle and spacecraft,
and then help to reduce the manufactory costs of rocket and payload.
TECHNICAL REPORT ISO/TR 19473:2015(E)
Space systems — Best practices for orbit elements at
payload — LV separation
1 Scope
This Technical Report provides the best practices for orbit elements at payload-LV separation. It
includes orbit elements and calculation conditions, calculation method of orbit elements and their
errors at elliptical orbit insertion of various payloads. The fit between the actual and expected values of
orbit elements can be used as a criterion of commercial launch.
There are many different sets of orbit elements. Each is best suited for a particular application. The
traditionally used set of orbital elements is called the set of Keplerian elements. This Technical Report
gives the calculation method of Keplerian elements and the transformation method of all the other orbit
elements, in order to satisfy different user’s need.
Affected by terrestrial gravitational perturbation, lunisolar gravitation perturbation and other factors,
orbit elements change slowly after orbit injection. Orbit elements calculation methods after separation
are not included in this Technical Report.
The technical communication and specific progress for orbit elements is relatively easy to be agreed on
by applying this Technical Report, which can contribute to avoiding possible disputes.
2 Symbols and abbreviated terms
2.1 Abbreviated terms
BIPM Bureau International des Poids et Mesures
CTP Conventional Terrestrial Pole
GAST Greenwich Apparent Sidereal Time
GMST Greenwich Mean Sidereal Time
GCRF Geocentric Celestial Reference Frame
GPS Global Positioning System
IERS International Earth Rotation and Reference System Service
IRM International Reference Meridian
ITRF International Terrestrial Reference Frame
ITRS International Terrestrial Reference System
LGEIF Launch Geocentric Equatorial Inertial Frame
LV Launch Vehicle
PZ90 Acronym of Russian Parametry Zemli 1990
SI International System of Units
TAI International Atomic Time
TCG Geocentric Coordinate Time
TDT Terrestrial Dynamical Time
THF Topocentric Horizon Frame
UT1 Universal Time
UTC Coordinated Universal Time
WGS84 World Geodetic System, 1984
2.2 Symbols
a semimajor axis
a earth semimajor axis of terrestrial ellipsoid IERS used in ITRS
e
b semiminor axis
E eccentric anomaly
e eccentricity
GM earth gravitational parameter used in ITRS
e
h apogee altitude
a
h perigee altitude
p
i inclination
M mean anomaly
n mean motion of satellite
p semilatus rectum
r apogee radius
a
r perigee radius
p
S GAST at the time of payload – LV separation
T period
t time interval between the launch moment and the perigee passing
p
t time interval between the launch moment and the payload – LV separation
SEP
u argument of latitude
V velocity
V , V , V projection of velocity in LGEIF
x y z
x, y, z projection of position in LGEIF
α flattening of the earth
e
2 © ISO 2015 – All rights reserved

θ true anomaly
φ geocentric latitude at launch point
e0
λ longitude at launch point
λ longitude of the ascending node in LGEIF
N
ω argument of perigee
ω angular velocity of the earth
e
Ω right ascension of the ascending node
3 Orbit elements and calculation conditions
3.1 Orbit elements
Six independent orbit elements describe the orbit of a satellite. Two elements describe orbit size and
shape, three elements describe orbit orientation, and one element describes orbit location.
Orbit size and shape parameters include the following:
a) semimajor axis;
b) eccentricity;
c) semiminor axis;
d) semilatus rectum;
e) perigee radius;
f) apogee radius;
g) perigee altitude;
h) apogee altitude;
i) period;
j) mean motion.
Orbit orientation parameters include the following:
a) inclination;
b) right ascension of the ascending node;
c) argument of perigee;
d) longitude of the ascending node.
Satellite location parameters include the following:
a) true anomaly;
b) eccentric anomaly;
c) mean anomaly;
d) time past perigee;
e) time past ascending node;
f) argument of latitude.
The orbit elements are shown in Figure 1.
Figure 1 — Orbit elements
The orbit ellipse geometry is shown in Figure 2.
Satellite
Semilatus rectum
r
Semiminor axis
b
Radius
p
True anomaly
θ
Apogee
C P
Perigee
Empty focus
A
Geometric center
Focus
ae
r
a
r
p
Apogee radius
Perigee radius
a
Semimajor axis
Figure 2 — Ellipse geometry
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