The German Aeronomy Satellite AEROS

8/7/2019 The German Aeronomy Satellite AEROS

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NASA TECHNICAL TRANSLATION NASA TT F-15,104 ';

THE GERMAN AERONOMY SATELLITE AEROSU. Picker, E. Bachor, P. Soppa and .W. Trogus

Translation of:"Der deutsche AeronomiesatellitAEROS." Raumfahrtforschung, No. 2, 1973, pp. ^9-57

CO P Y" ^ ^ ^ . >/

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONWASHINGTON, D. C.. 205 6 , - . , SEPTEMBER 1973


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THE GERMAN AERONOMY SATELLITE AEROS .,{

Ulrich Picker*, Edgar Bachor*, Peter .Soppa* andWolfgang Trogus** .

1. Introduction .On December 16, 1972, the U. S. Air Force in Vandenberg, 749

California launched the third German Aeronomy satellite (AEROS)into a near-earth polar satellite trajectory (Figure 1). At12:25 Central European time, the new and improved first stage ofthe Scout launch rocket ignited. After burnout of the 4th stage,

o . , 'which is a solid fuel rocket like the others, a successful orbitinsertion was reported. The exact trajectory measurement by theNASA tracking station resulted in a perigee of 218 km, an apogee ,.of 865 km and an Inclination of 96.9. This was a sufficientapproximation to the nominal values. The propulsion system ofthe satellite which'-is provided to perform trajectory corrections,was therefore first not required to provide the nominal lifetimeof one-half a year or to guarantee sun synchronous conditionsby means of an inclination correction.

The development project came to a conclusion with thislaunch. The first feasability study was carried out in Januaryof 1968 by the GfW (Association for Space Research). . his was* Dipl.-Ing., Dornier-System GmbH, 799 Friedrichshafen,Postfach 6 4 8 ** Dipl.-Math., Dornier-System GmbH*** Numbers in the margin indicate 'pagination of original foreigntext.


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done at the request of the Federal Ministry for Sciences andResearch (BMWP). . .

2. Mission and operation. , - f 2.1 Experiment s _

AEROS is an aeronomy satellite. It is used to research theupper atmosphere in the range between 200 to 800 km. A largenumber of parameters is determined at the same time, so thattheir mutual relationships can be studied. For example, thedependence on altitude, geographic longitude and latitude, timeof day and season and the influence of sun radiation are. deter-mined. The satellite carries the following five experiments(see Figure 2):

- A mass spectrometer (MS) for- determining the partialdensities of ions and mutual particles in the mass range between1-4 4 (atomic weight);

- Counter voltage analyzer (GSA) for measuring temperatures,i.e., the energy distribution of ions and electrons. At thesame time, the total ion density is determined;

- Impedance probe (IP) for measuring the electron density;

- EUV-spectrometer for measuring the flux and the spectro-distribution of the solar extreme ultraviolet radiation in therange between 150 to 580 A" nd from 300 to 1080 A;

- Temperature measurement device for neutral particles.(NATE) for determining the temperature and total density of theneutral particles as well as the concentration of molecular


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nitrogen. .

The experiment mentioned last was built by Goddard" SpacePlight Center at NASA. The others were built by Germaninstitutes.. ' . "

In addition, a sixth "passive" experiment results in a NASAdetermination of the atmospheric braking of the satellite andtherefore, the short time density variations of the atmosphere.For this, it is necessary to continuously have a very exact tra-jectory determination. . .

2. 2 Measurement^p_rgr_am

2,21. ..Measurement profile, .

During the measurement phase, so-called "measurementorbits" alternate with "idling" orbits. In the measurementorbits, the experiments are operated according to one of thefour programs.. During the idling orbits, they are turned off. /Under normal conditions, no additional ground commands areneeded in addition to the tape interrogations.

t

Each orbit starts when the satellite emerges from theshadow, which is detected by the sun sensor. The EUV experimentcarries out,during the first 10 minutes after entering theshadow, a special shadow program during every 4th measurementorbit. The other experiments continue to measure in the shadow.

The counter voltage analyzer consists of two sensors,whichmeasure over the Northern and Southern hemispheres^ A switchingto a different sensor takes place about 24 and 156 North ofthe ecliptic, with an accuracy of 15..


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~~ _MS 7 A C E 3 SEP. S T A G E ^..SURMgUI

;TACE 1

Figure 1. Launch of the Scout,carrier rocket with the satellite. AEROS. . " .

During the subsequent idling orbits, the complete tape unitwith the recorded measurement data is played out if there iscontact with one of the ground stations. In.addition, thebatteries are recharged and, if necessary, an adjustment is madeto the axis position and spin rate.

When neither of the two tape storage units are empty, onemeasurement orbit is automatically suppressed. Changes in themeasurement orbit/idling orbit cycle can be brought about byground command..

2.22 Measurement programs

The following measurement programs were planned,- Normal program- Special program-. Reduced normal program- Modified special program


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R E T A R D I N G P O T E N T I A L A N A L Y Z E R

T E M P E R A T U R E EXPER I M EN T( N ASA )

A X I S C O N T R O L COIL

HYDRAZINE P OWE R P L A N T

YO-YO DESPIN MECHANISMMPEDANCE PROBE

"Figure 2. AEROS Configuration.Each program has; a duration of one complete revolution and

runs completely automatically. All programs can be turned onand off by ground command at arbitrary (contact) times. Allprogram commands are immediately carried out during a measurementorbit. Only the "special program on" command is delayed up tothe next following measurement orbit.

Normal program

The normal program brings about quasi-simultaneous measure-ments with all five experiments. The MS and the GSA measuresectors which are symmetric with respect to the incident flowvector, i.e., when the angle between the sensor axis and the


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incident flow vector is smallest. The measurement times thereforedepend on the spin duration. For fluctuations within the range(10 0.1 rpm) the experiments are controlled by the ion sensorsignal in such a way that the optimumlangular ranges are .coveredj[sebelow "Measurement Phase Control". .&

Within each experiment, there is a switching to various opera-tional modes which will not be described here.

Special program

In this case the MS and the GSA carry out rotational sweepmeasurements in a certain sequence. All measurement modes of thetwo experiments are. carried out in 163 revolutions. The otherthree experiments measure just like in the normal program.

Reduced normal program

This program is provided if there is a failure of the twobuffer storage units. During the third and fourth spin, themeasurements with the experiments MS and GSA are omitted. Theother experiments are not influenced (nor are the operationalsurve-n.1a.nce data).

' . Modified special program

The"measurement.data are interrogated and transmitted in amanner which is not synchronous with the spinning. Instead,it is controlled by a fixed rhythm. In addition, revolutionscan measurements are carried out just like in the special pro-gram. This program therefore provides the complete set of theexperimental data, just like the normal program and the specialprogram. There is a certain kind of "data reduction" on boardthe satellite for the normal and special programs. In the6 ' . . '


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modified special program, all the measured values are . transmittedto the ground and are processed there. /

2 . 3 Mi s_ s :L on analy3i .,&

2.31 Trajectory design and development

The requirements for the trajectory were the following.- Initial apogee between 800 and 1,000 km altitudes.- Initial perigee on the" day side of the earth- Apogee during the lifetime of the satellite shall not drop

below 600 km.- Perigee as low as possible, but under 250 km altitude.- Lifetime of the satellite must be -at least 1/2 a year.- High inclination- Sun-synchronous 3 hour/15 hour trajectory.- Accuracy of sun-synchronous conditions o within 1 hour

( 15) over 1/2 a year.r - The .apogee should be increased again after about 5 months.

These requirements were satisfied by the , trajectory shownin Figure 3- The main parameters after launched were the following (the nominal values are given in parentheses):

- Perigee h = 218 km (NominalV230 km)- Apogee hA = ^5 km (Nominal:800 km)- Inclination . i = 96.94 (Nominal; 97.2)- Argument of theperigee: . u> = 168 (Nominalf 160)- Longitude of theascending node n = 312 (Nominal;:310 )- Mean anomaly . M = 350.5 (Nominal: 0)


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APSOA.

Figure 3. Trajectory of the AEROS,(Epoch: . December 16, 1972, 11 hours, 32 minutes, world time)The parameters derived from this are:

. - Semi-major axis.- Eccentricity- Anomalistic revolution period- Lifetime prediction (GSFC),

6919.5 km0.046?95.47 min

without raising the trajectory 200 days

The"3 hour/15 hour trajectory" condition led to a launchtime of 11 hours, 25 minutes (world time) with only a 10 minute"launch window". However, any day of the year was available asa launch day. The sun synchronous condition is provided by theretrograde trajectory inclination. Apparently, this is reachedto within an accuracy of better than 7.

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DaysFigure 4. .Altitudes over the lifetime.

200 DaysFigure 5. Deviation from sun-synchronous conditions,

.360. .__

- . . . . _ ! _ ; JFigure, 6. Position of the sun in the trajectory system.

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.The trajectory also has the following characteristics:

- Shadow, time between 32 and 34 min (32.5 to 3%% 'of the /5revolution time).

- Migration of the perigee is about 3.5/day (retrograde).. - Migration of the nodes is 0.9856/day on the average.: In this way, sun-synchronous conditions are obtained.- Radiation from the sun on the trajectory plane is alwaysat an angle of somewhat less than 45.- The sun crosses the equator on the ascending leg and onthe descending leg at the same local time, that is, at15 hours and 3 hours, respectively.

.Figures 4 to 7 show a few parameters as a function of time.They are based on our own (pessimistic) trajectory, predictions.2.32 Mission phases .

The mission can be decomposed into the following segments:- Launch and ascent- Acquisition .

* - .- Preliminary measurement phase ' .- Main, measurement phase, interrupted by- Second acquisition - Reentry

Launch and 'ascent

The launch direction was approximately to the South (azimuth -188). The events during the launch are shown in Figure 1.

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Acquisition

During this phase, we wanted to achieve the followingC33, [5], [6]: . <

-. Trajectory measurement- Attitude determination- Decay of the nutation- Adjustment of a spin rate of 10 2.5 'PJ}> and later tothe nominal value of 10 0.1 'rpm.- Trajectory corrections if necessary- Turning into the sun (with a deflection of less than 5)- Switching of the subsistance to mission operational condi-tions.

Because the trajectory elements were close to nominal (seeabove), we do not have to carry out the planned correctionmaneuver. However, we will discuss this in more detail, becausethe AEROS differs considerably in this regard from other Scoutsatellites.

iThe statistical'-injection 'error^of the ';Sc'out rocket is quitelarge because it.is a solid fuel rocket. In the case of theAEROS launch, we had to plan with

f A h p ~ = ~ + : ' l 3 / - 3 0 " k m ~ ., A h A = 21 1 kmA i = 1,2

with a probability of 95% (2 sigma-values). This meant that thecombined trajectory requirements (altitudes, lifetime and sun-'synchronous condition) could not be satisfied with sufficientcertainty. This is why the AEROS has a special engine for carry-ing out trajectory corrections.

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, 5 0 Day's!--im --

Figure 7- Geographic longitudeof perigee and apogee.-

Because of the possible de-viations from the insertion peri-gee, it was appropriate to setthe perigee at a somewhat highervalue (240 instead of 230 km).The perigee altitude actuallyachieved was 218 km and justi-fied the use of this method.

The following procedure wasadapted for a possible correctionmaneuver. After turning the spin,axis, the first trajectory deter-mination and based on the derivedcalculations of the lifetime and

the sun-synchronous condition, a decision is made as to whichmaneuver (one only!) will be made. The.following are candidates.

- Change in the perigee altitude- C h a n g i n t h e a p o g e e a l t i t u d e- Change in the inclination

1The following have priority. '1 - At least a lifetime of 150 days.2 - Perigee altitude between 220 and 240 km.

- 3 --Inclination between 96.9 and 97.5.

The execution of the maneuvers requires extensive preparationso n board a n d also o n t h e ground. . ' - -

' - Trajectory and attitude determination..- Trajectory prediction- Determination of the nominal thrust direction.- Determination of the delayed command for the required axisrotation.

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- Transmission of the delayed command for attitude change tothe satellite and execution on board.- New attitude determination.- Transmission of the delayed command for engine firing andexecution on board. .- Trajectory determination . "- Axis rotation back to the sun (autonomous on board oncommand). . .We will not discuss the details here. For. this, see references

[3] and [6]. These items will again become acute towards the .end of. lifetime, when the apogee altitude is again raised (2ndacquisition).

Preliminary measurement phase and main measurement phase

During the preliminary measurement phase, the subsystems-ofthe satellite are tested and the experiments are turned on in se-quence and calibrated. Also preparations for the measurements aremade. This was done up to the 12th day after-launching.

The subsequent main measurement phase will result in themain scientific measurement data. The measurements made at thelowest altitudes, that is towards the end of the lifetime, arebelieved to have the greatest scientific value.

3. Basic problems in the conception

3.1 Mec_ha.noLcal_de_s_ign

As mentioned above, it was only the trajectory requirementsin conjunction with the insertion accuracies of the Scout carrierrocket which led to the incorporation of a trajectory correc- -..tion system in the satellite concept...- .Since there was then-a

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possibility of a trajectory correction device, the raising of theapogee.after about four months was only considered secondarily sothat scientific measurements could be repeated at the same alti-tude for changed seasonal conditions.

In addition to the low spin rate, which was a constraint forthe scientific experiments, it was the continuous orientationtowards the sun, a requirement for the EUV experiment, which in-fluenced the basic design of the satellite. The orientation of 'the spin axis towards the sun limited the position and dimensions .of a solar generator and this stimulated the conception of anactive attitude control system. Solar cell paddles could not be.used because the experimental sensors required a field of visi-bility of 27T. It would have been possible to control the axisand the spin using cold gas or hot gas systems. The mass spectre--meter with its large measurement range between the mass numbers1 to 44 posed a severe restriction here, however. This measure-ment range includes all conventional gases which are used in con-trol systems with thrust nozzles. For this reason, ,we preferreda magnetic control system, which at the same time meant that fueland gas tanks could be dropped. This fact is.a positive sidephenomenon'if we consider the greatly limited 'space limitations .of 'Scout satellites.

There are other special features of the satellite which canbe attributed to the sensitivity of the mass spectrometer. How-ever, from a different point of view, these new properties werenot very favorable. In order to avoid erroneous measurements, /53based on exhaust gas production of the satellite, it wasnecessary to avoid organic surface coatings on the outer surfaceof the satellite. The only exceptions to this were the adhesivesurfaces which.protruded between the solar, cells of the generator.The covering was made of a steel with low outgassing. It is madefrom one piece and only contains openings along the conical part

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and diametrally across from the MS sensor. These provide pressur.equilization during the launch and rapid degassing of the innerspace during the first days of the mission.

(/Detailed analytical models were required in order to be ableto evaluate disturbing effects of backward scattered outgassingparticles on the mass spectrometer [8],

The closed steel shell, of course, produced further problems.For example, it had a large weight and access to the interior wasrestricted. Also there was a technical difficulty of making the'surface have the desired absorption and the emission factor forthe thermal concept without the use of any organic paints.Finally, this problem was solved by bombarding the outer coveringwith glass splinters. .

As far as the requirements for surface properties, the MSexperiment was supported by the counter voltage analyzer (GSA),.which required an eigen potential of the satellite structure 'which would be as low as possible because this is used as areference voltage for the experimental measurements. The transition resistance from the satellite to the plasma in itssurroundings should be below one kOhm-cm2, which of course isprovided in an optimal way by a free metallic surface.

3 2 Teleoiranuni_cat ipns_

For scientific reasons, it was necessary t.o perform all ;measurements at the same time. This resulted in a maximum dataflow during a measurement orbit. One of the two tape storage,units (about 3 million bit) is sufficient to store data froman entire measurement orbit. In the case where two usefulground'contacts (at least 5 minutes of data, transmission) are

' ' . . - . ' . 1 5


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Figure 8. Geome tric conditionsfor the switching point of themagnetic control system.

separated in time so much that two measurement orbits must berecorded between them, the redundant tape storage unit is

' ' 'automatically used to record data during the second measur ementorbit. .Because the data production of the payload was irregu-lar, it was sometime s required to have intermediate storage forthe data before recording by the tape storage unit. In onepart of the measurement programs, all experiments measure in adense sequence during one spin revolution (6 seconds). Becauseof the high data rate, it was necessary to provide for datareduction on board in order to be able to store all the essen-tial data in one frame (6 seconds). In the GSA and MS expe ri-ments, the measurement range was specified in an angular rangewhich was symmetric with respect to the direction of the incidentflow (identical with the flight direction). This direc tionis indicated by the ion sensor when such measurements are carried1 6 . . .


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out. However, the pulses from the ion sensor cannot be useddirectly for controlling the interrogation. In the polar regions,we can expect so-called ion winds which are so strong that theywould falsify any indication of true flight direction. Becauseof the trajectory geometry, the pulse frequency of the-- sensorchanges during one earth revolution (independent of the spindisturbance). This change is also not linear over time. In thepolar regions, it is so large that the limits of-adjustment fortelecommunications are reached.

For this reason, an average pulse frequency was formed overthe equatorial latitudes, and it was maintained over the poles.A deviation is only permitted if it is to be expected because ofa change in the spin rate during a revolution [9].

3 - 3 Att_itud.e_cntrol_and_at_titude_ measurement^As already indicated, we decided to use a magnetic control

system because of the perturbation problems caused by themass spectrometer as well as because of space limitations. Thissystem can orient the., spin axis with respect to the sun within5, and can continuously control the spin rate'to 10 rpm ig.'

In addition, the satellite must be oriented during the acqui-sition phase in such a way that a midcourse maneuver can becarried out using the engines which are aligned with the directionof the spin axis. The motions of the satellite which produceattitude changes are caused by servo-moments, which are based onthe interaction between the earth's magnetic field and magneticdipoles which are produced on board the .satellite at certain timesby means of coil systems. The axis orientation with respect tothe sun and the control of the spin rate are completely autonomousfunctions on board the satellite. Arbitrary changes in the axis

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orientation must be commanded through ground stations. Thesemaneuvers require preliminary calculations of the satelliteattitude in order to activate the coils,.using delayed.commands,at the right time and for a certain duration with the correctpolarization. . / ' ' .

Such maneuvers assume that so-called switching points existalong the orbit, i.e. sections along the trajectory which offerspecial relationships between the coil geometry on board thesatellite and the direction of the earth's magnetic field (Figure8).

The attitude measurements required for control jare based onseveral sensors which are oriented towards the sun, towards jthee a r t h T _ s _ horizon.Ia^d_tow_ard_s_t'he' _ d l r e _ c t _ i q n _ _ o _ f _ f c h e earth'js _magneti_cjfield. The measurement data are processed on board the satellitearid are also transmitted by telemetry for scientific evaluation.

In the same issue of "Raumfahrtforschung" CSpaceflight Research)there is a detailed report on the attitude measurement andattitude control system of the AEROS satellite. This is why wewill not discuss the details here.

3. Energ_y_supp_ly_

The small margin between supply and'energy requirement ledto a particular concept of energy supply. The area of the solargenerator could not be increased. It was necessary to optimizethe conversion losses, for the main part. The power productionof the solar generator is used in an optimum way because it canbe continuously operated at the point of maximum power. If thereis an energy requirement, a specially designed control systemcontrols the generator to the point of maximum power along aU/I characteristic which varies primarily with temperature.

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The satellite is in theearth's shadow during 1/3.ofa revolution and must thereforedraw power from batteries.AEROS had two batteri es of '.-''varying types: ' he Ag-Znbattery has a high capacityfor use during the acquisitiontimes, for which the satel^lite attitude results ininsufficient solar irradiationfor the generator. The so-called mission batt ery, aNi-Cd type can be'cycled morefrequently.

Figure 9. Internal structureand arrangement of the modules.

The energy supply system also had to solve the problem ofdegassing, because the Ag-Zn cells are not gas tight. This led to.a denser battery housing with an overpressure, safety valve.

t _ .

4. Description of the subsystems

4.1 Mec_han:Lcal_de_si_gn

The structure (Figures 9-12) ,shows~~th'at~tn"ere~Ts'~a d'i vi s ion of the satellite into an outer and inner structure. The innerstructure consists of three vertical equipment platforms whichare supported by three cross members. The three collar supportsmeet in a form of a star at the adapter for the launch rocket.The fuel tank is suppo rted in the center of the star by means ofa ring, and the ring constitutes a support between the threesandwich plate s. The internal structure supports the entire

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r'1 "-.v?-~*,-.-%>-;..*'X'-:-^-" ^-* ; - , * . . * . 'e?.-.*-*&a {-;-* j 'ji'y t\;:-~~*. r$?d ffiy&*~


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Figure 11. Prototype on theshake table.

spectrometer sensor had to bemade. Analysis and tests /then led to 'a rigid -connectionbetween the sensor and theinner structure using a three-legged support construction anda very soft connection with theouter shell. The entire outerstructure (shell plus solarcell support) is connected. withthe inner structure with dampingelements. It is connected tothe adapter, the ends of the -.collar support and to the in- .strument pl ate at 2 points. 'Both structural parts arecoupled with a large damping.The damping behavior can bevaried over a wide range by

exchanging the damping element s or by a continuous variation of thetensioning moments . \ In this way we were able to reduce theoscillation amplification of the entire structure compared withthe level introduced at-the adapter of the 4th rocket stage tobelow a factor of four. It was very important to save weight andto use materials with small amounts of outgassing. In addition,it was necessary to avoid ferromagnetic materials. Often theseparameters were contradictory, which was especially the case inthe steel outer shell. The collar support adapter and otherstructural pa rts, as well as most of the component housingpiec es, were made of magnesium. For example titanium screws wereused based on magnetic considerations.

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Figure 12. Preparation forspace simulation test,

The despin from l65(Scout'4th stage) to 10 + 2.5 rpjn: isperformed by a normal yo-yosystem. The fine' tuning to10 1% i r p m is provided by amagnetic spin control system.

Two different configura-tions were considered forbalancing the system. Thiswas 'the launch configurationand the mission configuration.The mission configuration be-came applicable about one weekafter launch, as soon as the.experimental shields had beendropped after a sufficientdegassing phase. These twoconfigurations could not bebalanced using only fixedbalance masses - o f a structure'^

and a sufficiently accurate attitude control could not be providedeither. On the other hand, the arrangement of the sensorsxalongthe periphery was not arbitrary because among other things, theplanned data interrogation of the experiments was prescribed andreference to the incident flow vector (flight direction) of thesatellite at a constant spin rate. Also the mutual angular posi-tion of the sensors along the circumference of the cylinder isprescribed by the constant spin rate of the satellite. This couldonly be solved by using an additional balance mass, which isjettisoned together with the coverings of the mass spectrometera n d NATE.- .

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TABLE 1. PRIMARY MECHANICAL DATA

Dimensions .-f* 'Altitude (maximum, without IS antenna): 7 0.1 mm.Diameter-(without telemetry antenna) 914.0mmMass: Launch ' 126.8 kg

Mission (tank filled) : 125.2 kgCenter of gravity (Mission configuration)

'Xs =' 0.18 0.01 mm "|. Y, = -0.86 0,01 mm

' ' 2i =_ 389.5 i: 0,5 mmPrincipal moment of inertia and moment ratios(Mission configuration, tank filled).

' i, = i 2 . 3 k g m 2~ i / o ; (around spin axis)|,/I2 =. .1.152 ',\il\3 = 1.106 '

Dynamic imbalance (Mission configuration) . .~lxy: 276.2 kgcm2Ix*: -2.4 kgcm2' Iv2: -7,7 kgcm2

Origin of the coordinate systems:Penetration point of the spin axis through the adapterseparation plane (Figure 2).

4.2 P_ropuls_ion_sy_st_em

The system (Figure 13) uses hydrazine as fuel. The spheri-cal tank contains a membrane which separates the hydrazinin the lower part from the propellant gas nitrogen. There arefueling valves and lines for each of the media. It is possibleto fill and monitor the pressure of each from the outside whenthe satellite is completely integrated [7].

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Figure 13. Structure of theAEROS propulsion system. Figure 14. Block diagram ofenergy supply.For safety reasons, the fuel was; only loaded at the launch

site in order to not have to handle the dangerous hydrazine..more often than necessary, the system was filled with alcohol orwith gas only for test purposes.

The thrusters consist of a two-seat magnetic valve. After* it opens by remote control, the hydrazine enters the combustion

chamber. .The combustion chamber is lined with a catalyzerwhere the hydrazine coming in is decomposed exothermally. Thereis an electrical resistance heating system and additional instal-lation in the combustion chamber and in the valve.

o4.3 Energy supp_ly_4. 31'Requirements .

The primary sources based on solar cells carried a special.control system ("Maximum Power Point Tracking" = MPPT) whTch'is usedto exploit the output power in an optimum way. Shadow times


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TABLE II. TECHNICAL DATA OP THE PROPULSION SYSTEM

Tank volume: 7.4 LitersFuel . 4.7 kg hydrazine,//Propellant gas NitrogenNumber of engines 2Maximum operational pressure 42 ataMinimum operational pressure 14 ataThrust range per engine: 1.86 kp (42 ata),0.77 kp (14 ata)Possible veloc ity change 80 m/secfor 125 kg satellitemassHeating cone valve 3-0 Watts

Combustion chamber 1.5 Watts

(1/3 of the revolution time) must be bridged by means of achargeable battery system. . . .

The secondary energy for the electrical users must be madeavailable in the form of direct voltage, of 28 V 5/6.and16 V2JL

4.32 Functional descr iptio n

Figure 14 shows the block diagram of the energy supply systemwhich satifies the require ments ment ipne d. It is characterizedby two parallel energy paths.

- Solar,generator - direct, converter - load- Solar generator, - charge control - battery system- Discharge control - load

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The energy flow in these two branches is determined bymeans of a "higher level control". The MPPT control s-ystem is themost important component of this device. With it, it is possibleto draw the maximum power from the solar generator. ,,,The maximumpower varies with temperature and irradiation conditions. In thisway, it is possible to minimize the number of solar cells, whichrepresents a substantial cost savings.

The energy flow itself is characterized by two states:

. a) Consumer power < solar generator power.

In this case, the users are directly fed by the direct con-verter, and the output voltage is controlled by the overvoltagecontrol system. The MPPT control circuit adjusts the inputimpedance of the charge control in such a way that there is a loadmatching between the total load and the solar generator. In thisway the entire excess energy is directed to the battery system..

b) Consumer power > solar generator power

In this case, the charge controller is blocked during the sunphase and the input impedance of the direct converter is adjustedbased on power. The entire solar generator power then flows tothe users. The missing power is taken from the battery systemusing the discharge controller, and the output voltage is con-trolled by the undervoltage control system..

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TABLE III TECHNICAL DATA FOR ENERGY SUPPLY SYSTEM

- Solar generator- 1300 1 ft cm n/p solar cells, 2 x 2 cmoutput .power over temperature: 60-80 WWeight: 1.3 kg (without structure) ' '- Direct converterOutput power: 40 W, efficiency: 0.93- Charge controllerInput power: 80 W, efficiency 0.93- Discharge controllerOutput power: 40 W, efficiency 0.91*- 16 V converterOutput power: 11 W, efficiency: 0.91- Ag-Zn battery: 22 cellsCapacity: 10 Ah, weight: 5. 5 kg- Ni-Cd battery: 29 cells + 1 couloumeter cellCapacity: 3 Ah, weight: 6.95 kg

The same holds true in the shade except that in this case thesolar generator and the charge controller do-'not operate. Thefour direct voltage converters adapt the varying voltage levels.The direct converter and charge controller operate according tothe "step-up" principle. The discharge controller and the16JV^converter use the "step-down" principles.

The battery system consists of a Ag-Zn and Ni-Cd battery. Thestate of charging is monitored by the .battery logic. The batteryconnected with the system is easily selected autonomously bythe battery logic. Of course, it is possible to intervene hereusing radio commands.. Finally, we would like to make a fewspecial features of this energy supply system:

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"Maximum Power Point Tracker" (MPPT)Up to now, such a control system has only been used' once in a

satellite in the United States. The MPPT used in the AEROS and.-''which operates according to the "Characteristic principle"(Dornier Patent) is remarkably simple and accurate. The corres-ponding test results and flight results support this statement.

Rechargeable silver-zinc battery.There is an unusual degree of rechargeability in this battery

compared with other missions. The battery is primarily used during the acquisition phase, but it could also be used as a"stand-by redundancy" system for the cyclically Ni-Cd battery.

4 .. 4 T e l _ e c _ o r n m u n i _ c a _ t i ^ o n s_ystem_

Telecommunication system processes analogue and digital datafrom ex periments and operational monitoring.systems using thePCM ethod. It is used to receive the ground commands corres

ponding to the tone digital command system. The structure of the telecommunication system is shown in /5

Figure 15. The telemetry encoder processes the data of theexperiments and from the operational surveillance during themeasurement orbit. Part of the scientific data is first inter-mediately stored in buffer storage units, in order to smooth thediscontinuous and high data rate of the experiments' mass spectro-meter and counter' voltage analyzer. The data processing andthe control of the data interrogation from experiments 'are per-formed according to four possible programs (four possible formats).

The synchronization of the data frame with the spin required2 8 '


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for certain measurement modes of the MS and the GSA based onthe ion sensor pulse is a task assigned to the time clock.This device also contains the on board clock, revolution" counterand orbit counter. ./

The digital control is performed in such a way that the phasecomparison between the ion sensor pulse and the data frame pulseis carried out. Phase deviations are transformed into thecorresponding bit rate changes, in order to correct the phase,displacement. . .

Because of the fact that the bit rate is not constant, thetape velocity of the tape storage units is also synchronized tothe bit rate in order to avoid disturbances of the storage rates.Because of the dynamic behavior of the tape storage units, thecontrol behavior for spin synchronization is also designedaccording to the properties of the tape storage unit. The fre-quency jumps during the control are limited to 0.3% in absolutefrequency range starting with an average value of 5?.

As from the Figure 15, only one data format is produced in thetelemetry encoder,, which is used for real time transmission aswell as for storage in.the tape storage units. Since two tapestorage units are provided, not only does redundancy appear butthere is also the possibility of bridging conditions when thereis unfavorable coverage by ground contact. In this case, thereis an increased storage capacity. The recording-to-transmissionratio of the tape storage units, is 1:25. '

The high frequency transmission system makes it possible tosimultaneously transmit RT and TT data. Als'o it is possible totransmit RT data alone. The telemetry encoder considered as adata producing unit, is only turned on during the measurement orbit

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I - a

q(1) Q6 '-HH AJM 0} < u ; - ia. a>x a,w o,

Telemetrytransmitter

o ; o ; < UOKI'

f '_!._. Command. ....Figure 15. Block diagram fortelecommunications.

During the measurement orbit, thedata from the encoder are simul- .taneously stored on the tapestorage unit and they, are trans-mitted to the ground as real timedata using HS. 'In.addition, ifnecessary, it is possible to playout the redundant device RT and TTtransmission during ground contactduring the measurement orbit inaddition to recording on one of thetape storage units. During theidling orbit, it is possible toturn on the encoder in order toperform housekeeping monitoringfunctions. In the case wherethere is simultaneous RT and.TT

e data transmission, (RT = 512 bps,TT = 12,800 bps) both signals

are phase modulated in the transmitter on to the same HF carrier(137-29 MHz) after suitable filtering.. By filtering with lowpass filters, the frequency spectrum of the RT data has an upperlimit of 2 kHZ in order to avoid cross talk into the TT spectrum.

tNo additional uncoupling filters are necessary.

Because of the increased ;b~Tt rate during taped data trans-mission, the transmitter HF power is increased from 150 mW to 1.5 by adding a power stagej considering the level balance of thedata transmission.

In the command link, the command decoder was modified fromAZUR. The threshold, the number of correct required "execute"words (Tone-Digital-Command-System) and the decoder logic were'changed.- The command length has proven itself well in orbit.

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The command signals which are not directly routed on aretransformed in the command distributor into commands for con-trolling the addressed devices. The command distributor con-tains program generators and devices for the' elayed executionof. commands. Also it can adjust the time duration of .these.commands. These-so-called delayed commands are required for arbi-trary attitude maneuvers which have to be.carried out atpoints along the orbit where the required magnetic conditionsexist but where there is no contact with the ground.5. Ground system and operation.

The German ground system is used for monitoring and operatingthe AEROS. (See contribution of M. Gass in this issue).Only the trajectory measurement over the entire mission and

the monitoring during the acquisition phase (n) were assigned toNASA.

When there is contact with the ground station, the fullmagnetic tape is played in about 4 minutes. As a minimum require-ment, about 30$ of all! earth revolutions should be measured. Thisamounts to 5 measurement orbits per day.

A monitoring program is used for status analysis of thesatellite systems and for error in identification in real timeduring - t h e satellite passes. This program provides theoperations engineers of the GSOC rapid decision aids. It isalso described in more detail in this issue (contribution ofE. Velten).

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Documents

The German Aeronomy Satellite AEROS