Aero9500 STK T2
Aero9500 STK T2
Aero9500 STK T2
Bytheendofthistutorialyoushould:
UnderstandthephysicalmeaningsoftheCOEandAOE
BeabletoPlaceasatelliteinanycustomorbitaroundtheearth
Becomfortablewithvariousorbittypes(andabletorecognisethem)
Setup
1. ClickCreateaScenario
2. Nameitappropriately(e.g.AERO9500_T2_COE_AOE)
3. Setstarttime:26Apr201602:00:00.000UTCG
4. Setstoptime:1Jun201602:00:00.000UTCG
5. Toaidvisualisationofdifferentorbits,youmaywishtoshowEarthInertialreferenceframe
inthe3Dview.Todothis,click3DViewProperties
3DView
Properties
6. NavigatetoVectorinthesidebar.
7. SelecttheAxestabalongthetopmenu.
8. ClickontheEarthInertialAxes,ShowLabelandDrawatPointcheckboxes.Selectavisible
colour(Brightcoloursworkwell).
9. SettheScaleto1.5(youmayhavetoscrolldowntoseethisoption),thenclickOk.
YoumayalsowishtodisplayareferenceplanesuchastheEarthCentredEquatorial(ECI)grid.This
gridistheXYplaneintheEarthInertialframe.Itisausefulreferenceforviewingorbital
inclinations,andfindingascendinganddescendingnodes.
10. Openthe3DViewPropertiesagain,andnavigatetoGrids.ClickShowunderECI
Coordinates,andchecktheShowRadialLines,andShowGridSpacingboxes.Itisagood
ideatochooseagreyorlightgreycolourforthegrid.ThenclickOK.
Nowthisissomewhatconfusing,sowellbeexploringeachelementtounderstandwhateach
elementactuallydefines.
1. Insertanewsatellite,butusetheDefinePropertiesoption.
2. Thisbringsupanewscreen,withyour6COEsbeingavailableontherighthandside.Change
theInclinationto1deg,leavingallothersastheirdefaultvalues.ClickApply.
Note:Whenchanging
valuesinboxesinSTK,
hittheenterkeyto
accepttheinputbefore
clickingApply.
Note2:Weareusinga
smallinclinationto
maketheorbitmore
visibleagainsttheECI
grid.
3. WewanttoseetheeffectsofchangingthevaluesoftheCOEs.Adjustyourscreensothat
youcanseethe3DViewandthepropertiespageofthesatellite.
4. YoushouldnowseeasatelliteinLEO(lowearthorbit),orbitingtheplanetintheecliptic
plane(orclosetoit).
5. ClicktheOrientfromtopbuttoninthe3DViewwindow
6. Changethesemimajoraxisto10,000km.HitEnterthenclickApply.Youshouldseethe
orbitincrease.
7. ClickOKtoclosethepropertiespageofSatellite1.
8. Insertanewsatellite,butthistimechangetheSemimajorAxisto25,000km(and
Inclinationto1deg),thenclickOK.
9. InsertathirdsatellitewithaSemimajoraxisof40,000km.
10. Youshouldobservethatallthreesatellitesareincircularorbitsclosetotheequatorial
frame,withdifferentradii.
11. Toclearuptheview,turnofftheECIgridandtheEarthInertialreferenceframe.Openthe
3DViewProperties,navigatetotheGridstab,anduncheckECICoordinates.Navigatetothe
Vectorstab>opentheAxestabanduncheckEarthInertialAxes.
Asyoucansee,increasingthesemimajoraxissimplyincreasesthesizeoftheorbit.
Eccentricity (shape) e
Theeccentricityofanorbitdetermineshowcircularitis.Aneccentricityof0isaperfectcircle,while
aneccentricityof1isaparabola.Wewillbelookingatellipticalorbits(0<e<1).
1. InsertanewsatelliteusingtheDefinePropertiesoption.ChangeitsCOEstohaveSemi
majorAxis25,000km,Eccentricity0,Inclination0deg.ClickOK.
2. Inserttwomoresatelliteswitheccentricitiesof0.35and0.7.(Unlessspecified,leavethe
remainingCOEsastheirdefaultvalues).
3. Youshouldsee3satelliteswiththesamesemimajoraxis,butwithincreasingeccentricity.
Notethatthehighertheeccentricity,themoreellipticaltheorbitbecomes.Theorbit
becomesstretched,buthasthesamesemimajoraxis.
e=0.7
e=0
e=0.35
4. Animatethesimulation.Whatdoyounoticeabouttheorbitalperiodsofthe3satellites?
Whatdoyounoticeaboutthevelocityofthesatellitesneartheirrespectiveperigeesand
apogees?
5. CreateanewsatelliteSemimajorAxis30000km,eccentricity0.7.
6. Comparetheorbitshapetosatellite3.Notethattheorbitshapehasnotlinearlyincreased
insize,butratherproportionally(accordingtotheinterrelationbetweeneccentricityand
semimajoraxis).
7. OpenthePropertiespageofthe4thsatellitebydoubleclickingitintheObjectBrowser.
8. ThePropagatormodelbeingusedisaTwoBodymodel.Thismodeldoesnotallowparabolic
orhyperbolicorbits.
9. ChangetheCOEsofsatellite4tohavesemimajoraxis700,000km,andeccentricity0.99.
ClickOKandanimatetheorbit.Thisorbitisclosetoaparabolicorbit,althoughifyouzoom
out,youwillseetheapogee(aparabolicorbitwouldhaveapogeeapproachinginfinity).
10. Deleteallsatellites.
Inclination (tilt) i
Theinclinationofasatellitedeterminesitstiltrelativetothelocalcoordinatesystem.Aninclination
oflessthan90degreesisconsideredaprogradeorbit,andaninclinationofgreaterthan90degrees
isconsideredaretrogradeorbit.
1. TurntheECIgridandEarthInertialreferenceframebackon.
2. InsertanewsatellitewithSemimajoraxis25000km,andallotherCOEssetto0.
3. Openthepropertiespageofthesatelliteandchangetheinclinationto45deg.ClickApply
andwatchtheorbitchangeinthe3DView.
4. Noticethattheorbitnowcrossestheeclipticplaneintwopositions.Theseareknownasthe
AscendingandDescendingNodesandthevectorconnectingthemiscalledtheLineof
Nodes.
5. Changetheinclinationagainto95deg.NoticethattheorbitisrotatedabouttheLineof
Nodes.Theorbitnowgoesoverthenorthandsouthpoles.ThisisknownasaPolarOrbit.
6. Changetheinclinationagainto135deg.Noticenowthattheorbithastippedoverthe
vertical,anditsmotionisintheoppositedirectiontotheoriginal.ThisisaRetrogradeorbit.
7. Toviewalloftheseorbitstogether,insert3newsatelliteswiththefollowingCOEs.
8. Observethechangestotheorbitincreasinginclinationcausesincreasingtilt.
9. Noticethatallofthesatelliteshavethesamepositionfortheascendinganddescending
nodes,exceptforthe0deginclinationorbit,forwhichthenodesareundefined.
i=95o
i=45 o
i=135o
i=0 o
10. Animatethesimulation(Resetasnecessary).
11. Notehowthesatellitesrotatearoundtheplanetinthe3Dviewand2Dview.Anysatellite
withaninclinationofgreaterthan90degreesappearstobemovingbackwardsonthe2D
view.
12. Deletethesatellites
1. InsertanewsatelliteleavingtheCOEsastheirdefaultvalues.
2. YoushouldseethattheascendingnodeisalongtheEarthInertialXvector.TheRAANofthis
satelliteisthen0deg.
3. ChangetheRAANto20degandhitApply.Watchthechangeintheorbitinthe3DView
window.Thepositionoftheascendingnodehasnowshifted20degEastwards.
RAAN=0deg RAAN=20deg
4. ChangetheRAANto40,60,and90degusingtheApplybuttontoseeeachchange.For
RAAN=90deg,theascendingnodeshouldcrosstheEarthInertialYvector.
5. Inserttwomoresatellites,andchangetheRAANofyourthreesatellitestothefollowing,
leavingallotherCOEsastheirdefaultvalues.
8. Deletethesatellites.
1. Insertanewsatellite,withthesemimajoraxis25000km,eccentricity0.7,leavingallother
COEsastheirdefaultvalue.
2. Notethelocationoftheascendinganddescendingnodes,andthelocationoftheperigee.
3. Changetheargumentofperigeeto30,60,and90deg,observingthechangestotheorbit.
4. Notehowthe3satelliteshavetheexactsameorbitshapetheirorientationismerely
twistedwithincreasingargumentofperigee.Theorbitsarestillinthesameorbitalplane,
andtheRAANisunaffected.Theonlydifferencewillbethealtitudeofthesatelliteatthe
nodecrossings.
Argumentof
o
perigee=40
Argumentof
Argumentof perigee=0o
o
perigee=80
True Anomaly
Thetrueanomalyisthespacecraftsangularpositionasmeasuredfromtheperigee.Itdefinesthe
positionalongitsorbit.
1. Insert4newsatellites,withthefollowingproperties:
2. Youshouldobservethatall4orbitsareentirelyidentical.Theonlydifferenceisinthe
satellitestartingpositionalargerinitialtrueanomalysimplyincreasestheanglefrom
whichthesatellitebeginsitssimulation.
o
Trueanomaly=80
o
Trueanomaly=40
o
Trueanomaly=180 o
Trueanomaly=0
3. WearenowgoingtocreateaReportoftheorbitalelementsoveraperiodoftime.
4. Keepsatellite1,anddeletetherest.
5. RightclickonthesatelliteandselectReportandGraphManager
6. SpecifytheTimePropertiesfrom26Aprto10May.
7. ExpandtheInstalledStylesfolder.Thishasacollectionofpredesignedreportstyles.
8. SelecttheReport versionofClassicalOrbitElements.DoubleclickorpressGenerate.
9. Thisgeneratesareportlistingthe6COEsatdifferenttimestepsoverthespecifiedtimes.
10. YoushouldscrollacrossuntilyouseetheTrueAnomalycolumn.Notethatitincreaseswith
timeuntil360o,andthenresetsto0o,designatingthesatellitespositionalongtheorbitat
thattime.
11. Deleteallsatellites
Argument of Latitude
Whenanorbitisperfectlycircular,thereisnodefinedperigee.Sincetheascendingnodeis
measuredfromtheperigee,anewdefinitionisprovided.Theargumentoflatitudeistheangular
positionofthespacecraftasmeasuredfromtheascendingnode.
1. Insertanewsatellite,withthefollowingproperties
COE Satellite1
Semimajoraxis 25000km
Eccentricity 0
Inclination 28.5deg
ArgumentofPerigee 0deg
RAAN 0deg
TrueAnomaly 20deg
2. Thesatelliteshouldbeinaninclinedcircularorbit,withitsstartingpoint20degreespastthe
ascendingnode.
3. NowchangetheArgumentofPerigeeto50degandobservethechanges.Thistwiststhe
orbitsothatthesatelliteisnow20+50=70degreesfromtheascendingnode.
4. UsethedropdownmenutochangetheelementtypefromTrueAnomalytoArgumentof
Longitude.
5. Thiswillautomaticallychangevalueintheboxto70degrees,indicatingthatthesatelliteis
70degreesfromtheascendingnode.
6. IfyounowchangetheArgumentofPerigee(tosay,0deg),itwillhavenoeffectonthe
ArgumentofLatitude.
7. Deletethesatellites
Longitude of perigee
Whenanorbitisequatorial(thatis,inclinationof0oor180o),thereexistsnoascendingnode.As
such,thereexistsnoRAAN.
Unfortunately,STKdoesnotofferthelongitudeofperigeeasanoptionwheninclinationissetto0o
or180o.Instead,bydefault,theargumentofperigeeandRAANCOEswillhavethesameeffectonan
orbit,angularlyreferencedfromEarthInertialX.
Totestthis:
1. ShowtheEarthInertialVectorsifyouhavenotalready.
2. Insert2satelliteswiththefollowingproperties:
3. Youshouldobservethattheorbitsareentirelyidentical.Furthermore,fortheseparticular
instances,argumentofperigeeandRAANhaveanidenticalandcumulativeeffectonthe
orbitforexample,inSTK,anorbitwithinclinationof0o,argumentofperigeeof20oanda
RAANof20owillbeidenticaltoanorbitwithinclinationof0o,argumentofperigeeof10o
andaRAANof30o.However,keepinmindthatthisisinSTKonlyduetoitslackofthe
longitudeofperigeeoption.
4. Deletethesatellites
True longitude
Thetruelongitudeonlyexistswhentheorbitisbothequatorial(I=0oor180o)ANDcircular(e=0).
Thetruelongitudeistheangularfromthe vectortothespacecraft.Theascendingnodeand
perigeebothdonotexist.
However,muchlikethelongitudeofperigee,STKdoesnotofferaproperoptionbankforinputtinga
truelongitude.Changingthetrueanomaly,RAAN,andargumentofperigeewillallappeartohave
thesameeffectontheorbit.Feelfreetoexplorethislimitationinyourowntime.
Orbit Types
Nowthatweunderstandthevariousdefiningcomponentsofanorbit,wewillbeexaminingdifferent
orbittypes.
1. Basedonthefollowingtable,insertatleast1satelliteineachtypeoforbit,namingthe
satelliteappropriately:
OrbitType Semimajoraxis Inclination Eccentricity Argumentof
(km) perigee
LEO 6700 28.5o,39o,51oor57o 0 Yourchoice
Semisynchronous 26610 55o 0 Yourchoice
Molniya 26571 63.4o 0.7 270o
Sunsynchronous 65007300 95o 0 Yourchoice
2. Animatethesimulation,observingthedifferentorbits.Payparticularattentiontodiffering
orbitalperiods,andinthecaseoftheMolniyaorbit,whereithangsover.
3. Feelfreetoinsertasmanysatellitesasyouwish,toobservetheirdifferentorbits.
4. Onceyouarefinishedexaminingtheorbits,eitherdeletethesatellites,orhidethem(to
hideaparticularobject,clickonthecheckboxnexttoitsnameintheObjectBrowser).
3 2
Where =398600km /s forearth.
2. Now,asiderealdayisactuallyonly23.9345hours.Calculatethesemimajoraxisfora
satellitetohaveasingleorbitinonesiderealday.
3. Beforeanimating,whatdoyounoticeabouttheinitialorbitsinrelationtoeachother?
Changethetimestepto600seconds,thenanimatethesimulationovertheentirescenariotime.
4. Whathappenstothesatellitewitha24hourorbitalperiod(GEO1)?Whatdoesthisimply
aboutgeosynchronousorbits?
Uploadyouranswerstothequestionstomoodlebeforetheendofthetutorial.
Summary
Inthistutorial,youhavebeenexposedtoseveralmoreadvancedoptionsforthe3Dviewer.You
haveexploredvariousCOEs,AOEs,andshouldnowbecomfortablewiththedifferencebetween
differentorbittypes.Youshouldbeabletoinsertanewcustomsatellite,andaccessitsCOEreport.
Additionally,youshouldnowunderstandwhygeostationary/synchronousorbitsarebasedonthe
siderealday,andnotthesolarday.