BS ISO 27852-2016 Space systems Estimation of orbit lifetime《空间系统 轨道寿命估算》.pdf

1、BS ISO 27852:2016Space systems Estimation oforbit lifetimeBSI Standards PublicationWB11885_BSI_StandardCovs_2013_AW.indd 1 15/05/2013 15:06BS ISO 27852:2016 BRITISH STANDARDNational forewordThis British Standard is the UK implementation of ISO 27852:2016. Itsupersedes BS ISO 27852:2011 which is with

2、drawn.The UK participation in its preparation was entrusted to TechnicalCommittee ACE/68, Space systems and operations.A list of organizations represented on this committee can beobtained on request to its secretary.This publication does not purport to include all the necessaryprovisions of a contra

3、ct. Users are responsible for its correctapplication. The British Standards Institution 2016. Published by BSI StandardsLimited 2016ISBN 978 0 580 90375 5ICS 49.140Compliance with a British Standard cannot confer immunity fromlegal obligations.This British Standard was published under the authority

4、of theStandards Policy and Strategy Committee on 31 July 2016.Amendments issued since publicationDate Text affectedBS ISO 27852:2016 ISO 2016Space systems Estimation of orbit lifetimeSystmes spatiaux Estimation de la dure de vie en orbiteINTERNATIONAL STANDARDISO27852Second edition2016-07-01Referenc

5、e numberISO 27852:2016(E)BS ISO 27852:2016ISO 27852:2016(E)ii ISO 2016 All rights reservedCOPYRIGHT PROTECTED DOCUMENT ISO 2016, Published in SwitzerlandAll rights reserved. Unless otherwise specified, no part of this publication may be reproduced or utilized otherwise in any form or by any means, e

6、lectronic or mechanical, including photocopying, or posting on the internet or an intranet, without prior written permission. Permission can be requested from either ISO at the address below or ISOs member body in the country of the requester.ISO copyright officeCh. de Blandonnet 8 CP 401CH-1214 Ver

7、nier, Geneva, SwitzerlandTel. +41 22 749 01 11Fax +41 22 749 09 47copyrightiso.orgwww.iso.orgBS ISO 27852:2016ISO 27852:2016(E)Foreword ivIntroduction v1 Scope . 12 Normative References . 13 Terms, definitions, symbols and abbreviated terms . 13.1 Terms and definitions . 13.2 Symbols . 33.3 Abbrevia

8、ted terms . 44 Orbit lifetime estimation . 44.1 General requirements . 44.2 Definition of orbit lifetime estimation process . 45 Orbit lifetime estimation methods and applicability . 55.1 General . 55.2 Method 1: High-precision numerical integration 65.3 Method 2: Rapid semi-analytical orbit propaga

9、tion . 75.4 Method 3: Numerical table look-up, analysis and fit formula evaluations 75.5 Orbit lifetime sensitivity to sun-synchronous 75.6 Orbit lifetime statistical approach for high-eccentricity orbits (e.g. GTO) 76 Drag modelling 136.1 General 136.2 Atmospheric density modelling 136.3 Long-durat

10、ion solar flux and geomagnetic indices prediction .146.4 Approach 1: Monte Carlo random draw of solar flux and geomagnetic indices 156.5 Method 3: Equivalent constant solar flux and geomagnetic indices .196.6 Atmospheric density implications of thermospheric global cooling 237 Estimating ballistic c

11、oefficient (CDA/m) .237.1 General 237.2 Estimating aerodynamic force and SRP coefficients .247.2.1 Aerodynamic and solar radiation pressure coefficient estimation via a “panel model” 247.3 Estimating cross-sectional area with tumbling and stabilization modes .277.4 Estimating mass 28Annex A (informa

12、tive) Space population distribution 29Annex B (informative) 25-year lifetime predictions using random draw approach .32Annex C (informative) Solar radiation pressure and 3rd-body perturbations .37Annex D (informative) Sample code for drag coefficient estimation via panel model 39Bibliography .41 ISO

13、 2016 All rights reserved iiiContents PageBS ISO 27852:2016ISO 27852:2016(E)ForewordISO (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 t

14、echnical 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 clos

15、ely 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 n

16、eeded 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 pate

17、nt 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

18、 in this document is information given for the convenience of users and does not constitute an endorsement.For an explanation on the meaning of ISO specific terms and expressions related to conformity assessment, as well as information about ISOs adherence to the WTO principles in the Technical Barr

19、iers to Trade (TBT) see the following URL: Foreword - Supplementary informationThe committee responsible for this document is ISO/TC 20, Aircraft and space vehicles, Subcommittee SC 14, Space systems and operations. This second edition cancels and replaces the first edition (ISO 27852:2011), which h

20、as been technically revised.iv ISO 2016 All rights reservedBS ISO 27852:2016ISO 27852:2016(E)IntroductionThis International Standard is a supporting document to ISO 24113 and the GEO and LEO disposal standards that are derived from ISO 24113. The purpose of this International Standard is to provide

21、a common consensus approach to determining orbit lifetime, one that is sufficiently precise and easily implemented for the purpose of demonstrating compliance with ISO 24113. This project offers standardized guidance and analysis methods to estimate orbital lifetime for all LEO-crossing orbit classe

22、s. ISO 2016 All rights reserved vBS ISO 27852:2016BS ISO 27852:2016Space systems Estimation of orbit lifetime1 ScopeThis International Standard describes a process for the estimation of orbit lifetime for spacecraft, launch vehicles, upper stages and associated debris in LEO-crossing orbits.This Int

23、ernational Standard also clarifies the following:a) modelling approaches and resources for solar and geomagnetic activity modelling;b) resources for atmosphere model selection;c) approaches for spacecraft ballistic coefficient estimation.2 Normative ReferencesThe following documents, in whole or in

24、part, are normatively referenced in this document and are indispensable for its application. For dated references, only the edition cited applies. For undated references, the latest edition of the referenced document (including any amendments) applies.ISO 24113, Space systems Space debris mitigation

25、 requirements3 Terms, definitions, symbols and abbreviated terms3.1 Terms and definitionsFor the purposes of this document, the terms and definitions given in ISO 24113 and the following apply.3.1.1orbit lifetimeelapsed time between the orbiting spacecrafts initial or reference position and orbit de

26、mise/reentryNote 1 to entry: An example of the orbiting spacecrafts reference position is the post-mission orbit.Note 2 to entry: The orbits decay is typically represented by the reduction in perigee and apogee altitudes (or radii) as shown in Figure 1.INTERNATIONAL STANDARD ISO 27852:2016(E) ISO 20

27、16 All rights reserved 1BS ISO 27852:2016ISO 27852:2016(E)Figure 1 Sample of orbit lifetime decay profile3.1.2earth equatorial radiusequatorial radius of the EarthNote 1 to entry: The equatorial radius of the Earth is taken as 6 378,137 km and this radius is used as the reference for the Earths surf

28、ace from which the orbit regions are defined.3.1.3high area-to-massHAMRspace objects are considered to be high area-to-mass (or HAMR) objects if the ratio of area to mass exceeds 0,1 m2/kg3.1.4LEO-crossing orbitlow-earth orbit, defined as an orbit with perigee altitude of 2 000 km or lessNote 1 to e

29、ntry: As can be seen in Figure A.1, orbits having this definition encompass the majority of the high spatial density spike of spacecraft and space debris.3.1.5long-duration orbit lifetime predictionorbit lifetime prediction spanning two solar cycles or more (e.g. 25-year orbit lifetime)3.1.6mission

30、phaseperiod of a mission during which specified communications characteristics are fixed. Note 1 to entry: The transition between two consecutive mission phases may cause an interruption of the communications services.3.1.7post-mission orbit lifetimeduration of the orbit after completion of all miss

31、ion phasesNote 1 to entry: The disposal phase duration is a component of post-mission duration.2 ISO 2016 All rights reservedBS ISO 27852:2016ISO 27852:2016(E)3.1.8space objectman-made object in outer space3.1.9orbitpath followed by a space object3.1.10solar cycle11-year solar cycle based on the 13-

32、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 wavelengthNote 1 to entry: Historical records back to the earliest recorded data (1945) are shown in Figure 2.Note 2 to entry: For reference,

33、 the 25-year post-mission orbit lifetime constraint specified in ISO 24113 is overlaid onto the historical data; it can be seen that multiple solar cycles are encapsulated by this long time duration.Figure 2 Solar cycle (11-year duration)3.2 Symbolsa orbit semi-major axisA spacecraft cross-sectional

34、 area with respect to the relative windApearth daily geomagnetic index ballistic coefficient of spacecraft = CD A/mCDspacecraft drag coefficientCRspacecraft reflectivity coefficient ISO 2016 All rights reserved 3BS ISO 27852:2016ISO 27852:2016(E)e orbit eccentricityF10,7solar radio flux observed dai

35、ly at 2 800 MHz (10,7 cm) in solar flux units (10-22W m-2Hz-1)F10,7Bar solar radio flux at 2 800 MHz (10,7 cm), averaged over three solar rotationsHaapogee altitude = a (1 + e) ReHpperigee altitude = a (1 e) Rem mass of spacecraftReequatorial radius of the Earth3.3 Abbreviated terms3Bdy third-body (

36、perturbations)CAD computer-aided designGEO geosynchronous earth orbitGTO geosynchronous transfer orbitHAMR high area-to-mass ratioIADC Inter-Agency Space Debris Coordination CommitteeISO International Organization for StandardizationLEO low earth orbitN/A not applicableRAAN orbit right ascension of

37、the ascending node (angle between vernal equinox and orbit ascending node, measured CCW in equatorial plane, looking inZ direction)SRP solar radiation pressureSTSC Scientific and Technical Subcommittee of the CommitteeUNCOPUOS United Nations Committee on the Peaceful Uses of Outer Space4 Orbit lifet

38、ime estimation4.1 General requirementsThe 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 resulting orbit lifetime estimate shall the

39、n be constrained to a maximum of 25 years per ISO 24113 using a combination of (a) initial orbit selection, (b) spacecraft vehicle design, (c) spacecraft launch and early orbit concepts of operation which minimize LEO-crossing objects, (d) spacecraft ballistic parameter modifications at EOL, and (e)

40、 spacecraft deorbit maneuvers.4.2 Definition of orbit lifetime estimation processThe orbit lifetime estimation process is represented generically in Figure 3.4 ISO 2016 All rights reservedBS ISO 27852:2016ISO 27852:2016(E)Figure 3 Orbit lifetime estimation process45 Orbit lifetime estimation methods

41、 and applicability5.1 GeneralThere are three basic analysis methods used to estimate orbit lifetime,3as 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 spacecraft a

42、s shown in Table 1. ISO 2016 All rights reserved 5BS ISO 27852:2016ISO 27852:2016(E)Table 1 Applicable method with mandated conservative margins of error (in percent) and required perturbation modellingSpecial orbit Conservative margin applied to each methodOrbit apogee altitude, kmSun- sync?High ar

43、ea- to- mass?Method 1:Numerical integrationMethod 2:Semi- analyticMethod 3:Table look-upMethod 3Graph, formula fitApogee 2 000 km Either Either No margin reqd; use 3Bdy+SRP5 % margin; use 3Bdy+ SRPN/A N/AN/A = not applicable3Bdy = third-body perturbationsSRP = solar radiation pressureMethod 1, certa

44、inly the highest fidelity model, utilizes a numerical integrator with a detailed gravity model, third-body effects, solar radiation pressure, and a detailed spacecraft ballistic coefficient model. Method 2 utilizes a definition of mean orbital elements,4 5semi-analytic orbit theory and average space

45、craft ballistic coefficient to permit the very rapid integration of the equations of motion while still retaining reasonable accuracy. Method 3 is simply a table lookup, graphical analysis or evaluation of formulae that have been fit to pre-computed orbit lifetime estimation data obtained via the ex

46、tensive and repetitive application of Methods 1 and/or 2. It is worth noting that all methods (1 through 3) shall include at gravity zonals J2and J3at a minimum.5.2 Method 1: High-precision numerical integrationMethod 1 is the direct numerical integration of all accelerations in Cartesian space, wit

47、h 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 spacecraft ballistic coefficient model based

48、 on the variation of the angle-of-attack, with respect to the relative wind. Atmospheric rotation at the Earths rotational rate is also easily incorporated in this approach. The only negative aspects to such simulations are (a) they run much slower than Method 2, (b) many of the detailed data inputs

49、 required to make this method realize its full accuracy potential are simply unavailable, and (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 J2and J3perturb