Closed Loop Simulation and Testing of the STC Direct Broadcast Satellite Attitude and Orbit Control System

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CLOSED LOOP SIMULATION AND TESTING OF THE STC DIRECT BROADCAST SATELLITE ATTITUDE AND ORBIT CONTROL SYSTEM C. A. Benet* and T. G. Tracy* *

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ABSTRACT

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An Att itude Determ i nation and Control Systems

The testing of a sate llit e's Atti t ude and Urbit Control System is difficult to implement in a realistic manner on tile ground . A number of

(ADACS) test program should address the follo;lln,; points:

methods have been employed, however, to provide an environment which dynamically exercises ttle

Valida ti on of th e control a l gorithms .

con trol system. Une technique specifically de veloped t o test the Attitude and Urbit Control System of a geosynchronous three axis stabi li zed

the flight hardware for all modes of opera-

corrununications satellite is presented here .

tion.

Verification at the closed loop

This

operae~ons

at

technique allows the exercising of the control

system

ifl

a real time, closed -l oop configuration

Uemollstration that the system characteris tics meet key performance specifications .

on a fully integrated spacecraft and with a min imum of test equipment interLacing.

The test set-up can, in addition, provide a training ground for the control system uperaturs, a means to exercise tile operational procedur~s arIa a tool to develop ana verify the conringency

KEY\WKDS Attitude Control Communications Satellites

plans .

Closed Loop Testing Dynamic Simlllation

INTKODUCTlUN Co~nunications

Satellites are evolving towards

higher power, tighter beam pointing accuracy and greater autonomy,

resulting in new demands on

already sophisticated Attitude Determination ana

Control System (ADACS).

Testing of such systems,

required to ensure adequate perfurmance for all mission phases, is complicatea by the fact that satellites operate in an eIlvironment wtlich is dif -

ficult and expensive to duplicate on earth .

How-

ever, dynamic testing ot the flight hardware

call

Testing 01 a fu ll y integrated

subsy.t~~

is only one phase in a comprehensive test IJru gram, from "bench testing" at the ind~vidual cumponents tu the overall checkaut dt l~le spac~ ­ cratt level . A test prograllL sllould proviae the ultimate veritication at the control ld~S i[l a dynamic environme n t which incorporates the tligtll hardware. 1he testing perfurmed at the compone n t level, as well as the sirnulations run ~n support at tIle control law development, dn~ ~ntegrdl parts of the required conLprehenslve qualitication drld acceptance tests . This appruact1 enLphasizeb cur relation between the test data ana the perfunlL anCe requirenLents J thus estBolishinb continu~ry between the successive rest phases. llle test progrctllL is structured to encoI:Lpass all the IliUdeb of a missioIl sequence represeIlring rile actual spacecraft flight . when appropriate, each mode is tested under worst case conditions at space cratt mass properties, sensor and actuator characteristics, environmental extremes with boell pr~nl~ and back up equipment configuratiuns . Failure modes are simulated and switchover to redundant units is verified.

be accomplished by closing the control loops about a realistic satellite niodel operating in a

sitnulated environment . A test progrdln c&n be devised to verify the design and assure a tLnlely

correction of potential problems. There are several nlethods that can be used tor dynamic closed loop testing. These rlIay be classi tied into two general techniques; those using sta tionary targets ano depending on spacecraft motion for sensor stimulation, and those using movable tar~ets with a stationary spacecraft . The first

A tes t program WhiCh illustr~tes the testin~ phi losophy previously established is the Accep tan ce Testing of the ~atellite Television Corporation (STC) Direct Broadcast S~tellite (DBS) current l y

technique utilizes either a 3- axis girnballed table

under construction at kLA Astro tlectronics ~rL Princeton, ~ew Jersey , USA. The satellit~ is a three-axis stabilized spacecraft based url the kCA

or a gas bearing table . The second technique which utilizes serv~ driven earth sensor targets and chopped sun sensor and horizon sensor sources

Stabilite® control principle (Kef. 1).

was selected by KCA for testing of the STC/D~S Connunications Satelli t e . Although there are

T~ble

1

indicates the sequence of tests cuvering the Olt ferent modes of operat~on. For each test a briet description at its purpose ~s given as well as cl list ot t tle pertorhlance llleasurenLents . The ~Hltn e of t his table is to show, using a spec ~ t~c eX ample, the extent of a particular test prograu ••

unique advantages to either tecllnique of testing, the one chosen he r e is cost effective and permits

testing of the fully assembled spacecraft through all ~hases of integration including testing in a thermal-vacuum chamber.

,I I i

318

C. .\ . Be ller alld T. C; . Trac,' TABLE I Test

ADACS Test Program

Function Tested/Description

POND

Parking Orbit Nutation Damping. Acce leromet e rs are stimulated at a frequency and amplitud e repres e ntin g the nutational motion o f th e spacecraft.

HSA/SSA

Ho ri zo n Sensor/Sun Senso r. Attitude data collection. Attitude monitoring mod e in tra nsfe r orbit.

Ne a su rem e n t

5 /

Pe r formance

Th r uster firin g s. Abort Condi ti ons . Ove rri de Func tion.

Sun a n g l e deterrn i na t .on.

St~mu lati on

of both HSA and SSA by in f r a r ed and chopped li ght sources respec tivel y . T IMElJ THKUSTIl
Automatic a nd manual tllrust e r operation.

Th ru ster - on times. Thru ster pu l sew idt h selection. ~nable/disable funct i on .

SPM

Spin Precessio n Maneuver . S timu l at i on of th e Su n Se ns or Command Eye input.

Thruster firings . ~nab l e/ Disable tunction .

EART H ACQUISITION

Pit ch loo p l ogi c t est ing for va ri ous i nitia l attitude co nditions with respect to th e Ea rt h .

Pitch e rro r . Momentum wheel reac tion. Acq ui s i tion domain verification.

P ITCH OFFSET

Aut oma ti c and g r ound c ommanded pitch offd et capabi l ity .

Spacecraf t response. Overshoot auto/ma nual o fts e t operation.

ROLL OFFSET

Aut oma tic and g round commande d r oll offset capabi li ty

Spacecraft respo nse. Over shoo t. Auto/Hanua l Of fset Opera tion.

POID

Product Of Inertia Damping . Simu lation o f nut ational motion and chec k for damping capab i I it y .

Observe nutation damping . Tim e Constant he a s ur eme nt. Pit ch Ax i s Int ~ racti un .

MAGNJ::TIC CONTRUL

Ro ll /ya w co ntrol lo op . Va r y r o ll erro r and c heck fo r p r oper turn o n /ofl of the

Magnetic t orq ue r curren t. Timi n g of on/off ope ratio n. Thresho l d .

ma g net i c torquer .

!WLL CUNTkOL BACKUP

Koll co ntrol v.a thrusters Ap pl y dis t urbance torques large enough to force back up c o ntrol takeover .

Po intin g Verifica ti on . Threshold.

E/W STATIONKEEPING

~ast/West o rbit contrul maneuvers . Testi ng of var IOU S t h ruster combina ti ons .

Pointin g Verit i ca tl un . Abo rt ~ond i ti o ns . Thru s t e r o pe r ation . Transition to non - tllr uster mode .

l
North/South o rbit con tro l maneuver. Tes ti ng of varIOUS thrus t er combina tion s .

Pointin g Ve ri fica ti on . Back up Ti mer Abo rt. Gy r o OriEt Calib ra tion . Transition to nou - th ruster mod e .

ANOMALOUS OPERAT I O~ lJETECT ION MUDES

Testing of a ll abort logic combinations.

Safe mode thr esho ld leve l. Back - up tImer veritica ti on .

REDUNDANT EQU IPHE NT SWITCHING

Failure of fli gh t eq uipment simulation . Switch in g to re dundan t uni t.

Transi e nt measurements tarth recaptur e capabil it y . capab i l i ty.

APEMAC

Mome nt um Adjust Mode . Force whe e l o utsi de its speed dumain and ve rif y thruster firin g .

Th ru ster pulsewiath combination/s eq uen ce Po intin g verifica ti on .

LONG TERM ON-ORBIT OPERATION

No r ma l on- o rbit operation

Pointing verification. Automatic momentum un loaJing. Au t omat i c roll a nd p itch biasing.

STC Direct Broadcast Sate llit e ;\tliturie SENSOR STIMULATIONS

with a microprogramme r and/or custom built interface un its. Just to me nt ion a few:

EARTH SIMULATORS Control of the AlJACS - Simulation lntertace . There are generally two t ypes of earth sensor targe ts. For a simulation built around a flight simulator supporting the sensors, the earth target cons ists of a hot plate providing infrared radiation in the 14 to 16 ~ band. Depending upon the distance f r om tar get to sensor, the size of the ta r get aperture varies such that the simulator

subtended angle tended angle as The temperature surface heaters

is about 17 ", i.e., the earth subviewed from geosy nchronous orbit. of the target (aluminum disk with bonded to the back side) can vary

to simulate various earth radiance while the radi-

ance g radient across the face of the simulator is kept to a mlnlmum, usually w,th,n +1 0% . The front side of the target is coated to provide high emissivity . More sophis ticated definition of the target size can be achieved with a truncated conical section (cooled o r at room temperature) with th e cone geometrical characteristics selected to pre vent reflections from other external sources into

the field-of-view of the sensor.

Generation of the target or f li gh t slmulator commands - Storage of simulator state

variables. Conversion of data to e nginee ring units and

processing of the data for evaluation purposes. Storage of system telemetry data. Plotting of test results. The simulation is organized as shown in Figs . 3

and 4 for the hardware and software respectively . The satellite model is located in a Hewlett-Packard A-900 computer. Some of th e interfaces with the AUACS are provided through a multiprogrammer (HP 6942A) a nd a custom built interface unit.

Germanium lenses

are usually nlounted on the ea rth sensor to pro -

vide a collimated energy source for the sensing detectorSa

SPACECRAFT MODELLiNG ANI) SUFTwAkt IMPLEMENTATION

A second typ e of earth target is the servo-driven

Modelling of the spacecratt is designed

"hot plate" which can be mounted directly on the ea rth Sensor. Use of a col linl ati ng lens mounted

accurately simulate thl::: dynamics aud kinematics of the satellite due to control , internal and enviro nmental d i sturbance turques.

within the earth sensor permits the hot p l ate to

[0

be equivalent to the p r eviously described ea rth

target. The type used by RCA (Fig . 1), for the STC/DBS , consists of a small elliptical aperture (2 en, x 2 . 8 cm) driv" n by two stepper - motors. The

The tollowing torques are generated: The momentum

Btld

r eaction

wll~el

torques .

motor drive maximum steppi ng rate is 700 steps/ sec , resultin g in a maximum r ate of 4.2°/sec and

The torques due t o thruster tirings

5 .b' /sec for roll and pitch, respectively, rates

(including plume lmpingement etfec[s).

nluch higher than any test simulation requirement.

The range of the tar ge t motion (+10' in roll and !15' in pitch) is sufficient to Zover the need of the simulation. The r esolutio n of the target motion is O.OOb'/step and O.OOS'/step for roll and pitch respectively. Another kind of ear th simula tor is required for satellites which are spinning in transfer orbit . An infrar ed radiating eleme nt chopped at the proper frequency can simulate the spinn ing

The magnetic torquer e tt ect . The torques due to the flexible appendages (e.g ., solar array). The toraue generated by the slewing of the solar arrays. External environmental torques (e . g ., solar radiation pressure, gravi t y grad i ent) .

spacecraf t (Fig . 2) .

SU~

SHlULATOkS

The sun is a spheri cal source 1,392,400 km in diamete r l ocated 150,000,000 km f rom the tarth. Although impressive as these figures can be, from our van ta ge point we see it as a beam of light

with an angle of approximately 1/2 degree .

When

viewed from uuter space without atmospheric absorption the Sun ' s spectrum approximates a 6000 0 K

black body with opectra l lines supe rimposeo.

Such

a spec trum is ref erred to as the Ai r Mass Zero (Al'IO)

The test software is designed around a s~t ot tasks with proper priorities. Each sinlulation cyc le includes data acquisi(lon, integration, coo rdinate tran sformation , (arge( or flight Sim ulator commands displays, sto ra ge , e tc. The silllulation cycle tinle varies b~tw~erl 4U lO lUU msec, depending on the dynamics to be simulated and the capability ot the main con,putt:: r.

So la r Spec t rum.

As an example or software implementa tion, the KC1"\.

design is depicted in Figs. 5, band 7 . Ille Attitu de ~ynamics Test Simulator (nDIS) (~lg. b) (Photog raph 3) calculates tile spacec raft bOdy rate s and arlgles in r esponse to data trum t he

A high pressure Xenon lamp is an excellent art iticial source for solar simulation.

ADACS test set interface. lhe Homentum \,heel Assemb l y (HwA) speed and pivot angle, the thruster fire signals from the Command Logic

Processor (CLP), and the magnetic torquer coil COMPUTING REQUIREMtNTS

commands are processed and used to update the spacecraft dynamics. In turn, the simulation

provides stimuli for the spacecraf t sensors and The hea rt of the closed loop simulation is the

th e sensor tar gets .

central computer, residence of the spacecraft

model and its environment.

The modelling tech-

nique i s addressed in the sof tware portion of

this paper . Some o t her functions can be performed by the control computer itself or in conjunction

A set of dynamic equations is used, includin g the

lowest modal freauencies of the sola r arrays. At each simulation cycle all the dynamics variables are updated .

c:. .-\.

Belle! alle! T. C. Tracy

The cycle time of the AVTS computer is 100 milliseconds, 26 milliseconds of which are taken by the dynamics computation. The balance of the time is required by the input / output subroutines (to and from the multipro g rammer, access to disc,

Fig. 12 is the corresponding all software simulation. A good correlation exists between the two dynamic respons es for a capture time of ap-

proximately 12 minutes and a pitch angle oVer shoot of less than 12 degrees.

etc .) . No rth-South Stationkeeping Ma neuver Pre-Test Simulator Va lidation To be a true representation of the spacecraft in flight the test equipment's own dynamics should not affect the testing . In addition, the spacecraft modelling should be detailed enough to inc l ude all effects that fall within the bandwidth of the on-board controller .

This test is designed to demons trate the control system capab11ity to maintain the spacecraft attitude during a stationkeeping maneuver. In this mode the thrusters are used to simultaneously

change t he orbit inclination and to keep the sat e llit e attitude within specifications . The on-orbit spacecr&tt dynanlics is s i mulatea in

Spacecraft modelling verification can be accom plished by comparing specific output signatures with corresponding simu l ations run during the sub-

system design phase.

For instance, the following

cha racteristi cs can be monitored: spacecraft mo mentum change in r e sponse to specific thruster firi ngs, nutation frequency, individual stimula-

tion of the flexible modes by proper selection of the initial conditions (fre quency and damping racio). As for the simulator cilsracteristics which cannot be comp letely eliminated, it is imperative that they be measured and taken into account when mak ing comparisons between the tlight hardware test

results and the all software simulations .

For

instance, the loop delay can be measu red by monitoring rat e gyro r espo nses to thruster firings .

This delay can then be introduced in the sof twar e

the AVTS compu t er as previously described and the outputs of th e model are used to drive the earth sensor tar ge t and to simulate the rate integrat-

ing gyro signals. Hoth simulator data and space cratt telemetry are co l lected [or off-line test evaluation .

ri gs . 13 and 14 are representative

ot the type ot data collected during the teHt. The test sequence follows tile ditte r en t ph&ses ot a stationkeeping maneuver; gyro drift estinlation - distu rbance torque Calibration - thruster utt pulsing mode during which the Uelta - V nlaneuver is pertorlned - thruster on-pulsing mOCe (luring wInch transition to nornJal operation occurs . A aetailed

description of the STC/UHb Stationkeeping loop can be fOUIld in keterence 1. lhe attitude excursions during the torque calibration intervals are the results of estimator gain selection intended to

demonstrate the proper operation of the built - in

c lo se to th e actual operations of the satellite

ca libr ation logic. In normal on-orbit operation these transients can be minimized by se le cting tIle loop pa rameters ba sed Ofl pre-tlight aisturbance torque predictions or estinlates derived (runl p r evious maneuvers . Fig . IS shows ctle resulcs ut a silnilar maneuver conducted with an all - soitware simulation . Comparin g tile hardware and sottware sinlulation trac es demonstrates the ~roper ImlJle mentation ot the control algorithms in the tlight equ ipment .

o n orbit . Of course, any semi -automatic procedure cu n also be implement ed , allowing routin e command -

East-West

simulati on to assess the phase margin of the con -

trol loops. The Mechanism of Running Tests Running tests can be either fu ll y under computer con trol, via test softwar~ called "module s " or under the operator's control, thus making testing

S ta tionke~plng ~laneuve r

in g to be generated by the compu ter and leavin g to the operator the freedom to intervene in response to the dynamics of th e s imulation . This latter approa ch is well suited for cr ying onorbit procedures, contingency and recovery plans,

although it may be argued that these issues fall beyond the purpose of simple s ubsystem testing . Testing at ReA is performed on a fully integrated spacecraft (Fig. 9), although some developmental tests were performed previously

on a system in-

stalled on a bench (Fig. lU) . rrom the list in Table I which briefly describes the overall test program, a few test s are now revi ewe d to show

A simi lar set ot data i::. sho""n for an l ast-west maneuver in which the same thru s ters used lur North - South con tr ul are used in all o n-~ulse j attitude co ntr ol mode while statlonkeepin~ thruster pulses are tIre d ill tile alung-t r ack

direction .

Figs. 16 and 17 show the p itCh, roll

and yaw ang l es, before, durlflg and alter ttle maneuver . Again, tu assess (he systenl perfonnance, Cl cOlnpar i son is made with an all - sottware silnula tion under identical maneuver condi ti ons (Fi g . l b) .

!10W

Closely the hardware testing verifies the all software desi g n simulation s . The testin g me thud just described has been suc -

Earth Capture

cessful l y implement ed on the SIC / UHS program .

This test evaluates the capability of the Contro l ~ystem to lock the satellite onto the Earth once

already mentioned in the intrOduction this technique is only one ot seve ral ways [0 dynamically stimulate a contro l s y stem. Amongst th~ ffi bn y

the system momentum is within a specified dynamic

range. The pitch earth capture uses the closed loop configuration of th e ADTS wherein th" realtime dynamic simulation drives the earth sensor

advantages of this method, the tollowing should be noted: Allows dynamic closed-loop testing of a rully integrated spacecraft at 011 levels of inte-

target to simulate the satellite mot ion. The wheel torque which would be applied to the spacecraf t in flight is determined in the AnTS from

gra tion. The same tests can be ~erturllit:!a ambient as well as in a (hermal VaCuum

t he direc t measurement of the moment um wheel

chamber.

speed variations . The domain of ea rth captura bili t y is fully tested by varying the spacecraft initial conditions (position, angular rates , r.lomentum, etc •.•. ) . Fig. 11 shows the satellite pitch and roll angles during the ea rt h capture .

A.

Makes the spacecrat t harness an integral

pa rt of the system under test . The interfaces are truly in their flight configuration.

d.t

STC

D irt'!1 BroadC
to thank all the RCA personnel involved for their many contribut i ons . A special reco gn ition go es to Ea r l Main from CUhSAT Laboratories lL-la r ksburg , Naryland) for h,S uutstandi n g su pp o rt .

S i multaneous testing of other spacec r aft subsystems (Powe r , Communications, Telemetry and Command) is possible . Interferences of the test equipment dynamics with the subsystem perfonnance measurements

are kept to a

KEFtR ENCLS

mInImUTTl .

The capital investment in terms of test equipment is relatively modest when compared with other closed - loop testing techniques.

1.

Benet, C. A., et a1 (June 11.)':' 5) Ine d,tt itu de DetermInation and Lon tr ol ~ubsys t em or ( Ile ~'lL

(Satellit e TeleVlSioll Lurporatlon) 0 1re Ct 1 0th i r Ln ~yll,pos iLJ Iil ! To ulous e , France. Broadcast Sate l litt:,

;\CKN UWLEDGNEI-iTS Ttle development of such a test facility is necessarily t he work ot a team . The authors would like

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