I

t

/

/

N91-16051

AAE 241 Spacecraft Design Project

Group #5

Spring 1990

Marty

Stephan

Julie

Dave

Tim

Kim

https://ntrs.nasa.gov/search.jsp?R=19910006738 2019-05-09T14:35:50+00:00Z
INTRODUCTION

PULSE, Pluto Unmanned Long-Range Scientific Exp!c:-e:-,is a::

unmanned probe tha _ will dc _ =_y_v c c _l,_o T_ _ = _ _,_,_

weight, relatively ....low costing vehicle wl-=_'_, u_il _-_-_,___ _L,_.___: _-_=__-

the-shel f hardware but nct mat__ria!s t_._,=_=__..es _..-......... -

available after 1999.

PULSE will be launched within the first decade _f the

twenty-first century.

ORIGINAL PAG! iS

OF POOR QUALIT/
MISSION MANAGEMENT, PLANNING, AND COST

I.I INTRODUCTION

In the subsystem of mission management, planning, and cost

many selections were made. The mission type, trajectory: and

launch date were selected. The optimum de!ta-v and cost cf the

project were also calculated.

1.2 TYPE OF MISSION

A flyby was the type of mission selected. This se!ecticn

was made due to its low delta-v, short mission du_--ation: and

simplicity, all of which are directly related to this mission':{

low cost.

Simplicity was a main issue in selecting this missic.n cl_.:.=.

Since there have been no missions to Pluto and P!utc'__ d _c______"=_

from the Earth is very far, very little is known about P]ut- _.:-..f

Charon. Therefore, before a high-cost, elaborate mission _s_: ]:<

sent, scientists need more accurate information. A flyby mi:{:=_i: <

is the most efficient way to get the information that :_ no,def.

1.3 TRAJECTORY

The trajectory selected for this mission is a direct E]rth

to Pluto path. Again, simplicity was an important issue in the

selection process. The more complex a missic.n, the greg.te'_- th,=

opportunity for something to fail. So by using a direct 9_th,

simpli=ity is _ptimized. _?]_ .... -_[ i_
1.4 MISSION DELTA-V REQUIRED

The delta-v required for the PULSE mission is 8._06

kilometers per second from a parking orbit around Earth.

1.5 MISSION TIMEL!NE

The launch date was determined to be January 30, 2003.

arrival at Pluto was determined to be February 1 _01a _'_-

mission length is 16 005 years The launch date was _h<,_n _....

selecting the date with the optimum de!ta-v. To obtai:-_

selection of dates, data was input for the first of evevy

month of every year from the year 2000 to the year 201C.

1.1)

1.6 COSTING

The costing process of this mission was done in s+v__'-_.!

steps. First, for each subsystem, the direct labor hou_-s _=nf t].s

recurring labor hours were calculated This was done '_'" :e_._-_.

different formulas that used the mass of each subsystem an! t.h.---

number of spacecraft. The number of spacecraft c_=ueJ were _r-or.'_-.

three of which are flight ready and one which is used in _4-_

integrated ground test system.

Next, for each subsystem, an inheritance class had to be

defined• Class One is an off-the-shelf buy. Class Two is an

exact repeat of a subsystem. A Class Three inhe_citance is the

use of a previous subsystem with minor modifications. .-%_.,_ ..... =.=

Four inheritance is also a use of a previous subsystem m:_',*_,:[_-
° J,

A Wi.lOp

!

t

_';-_ IS

OF POOR QtJALITY
OF' POOR QUALITY'

major modifications. Finally, a Class Five inhe_citance i_ :::

entirely new subsystem. (Table !.I)

The next step was to convert labor hours into laho_c .-.e:ct.

Then the labor costs were converted into total costs. The

conversion factors were given in Fiscal Year I:_,,_..wh:_-_-., n ........ -

be converted to Fiscal Year 1988. This was done b,/ llSi:IJ a

consumer price index. The consumer price index fcz- =-_ ....i*-e:::_, i:-<

1977, with a base of 1967=100, was 181.5. The consume:- price

index for all items in 1988 with a base of 196v=100 was _-4

(Appendix !) .

.=inally, these conversions were made for each sub_._=_,_.v=__,.,___._"_

then added to obtain the total cost of the project. (Table 1.2!.

The total cost of the PULSE project is about 1.7 billion dcl!:r}.

1.7 EFFECTS ON SUBSYSTEMS

Many of the selections made affected the se!e__tic::.: zf tl:e

other subsystems. The selecting of a flyby affected the s::ie::c-:

instrument selection. Because the mission is a flyby, only

instruments which can be used quickly and at a distance :s__:!d ]-e

used. The power and propulsion subsystem was also affected. E?

utilizing a flyby instead of an orbiter or a lander, less f-::-I _:,:.-:

needed. These factors also affect the design of the

structure.

The length of the mission and the trajectory selected also

affected the other subsystems. Due to the length of the :?.issicn

16.005 years, science instruments and other materials whi:h

lifetimes exceed 16.005 years had to be selected. T]:e:--e
SUBSYSTEMINHERITANCE CLASS

Category

Structure

Thermal Control

Propulsion

Attitude & Articulation

Telecommunications

Antennas

Command & Data Handling

RTG Power

Line-Scan Imaging

Particle & Field Instruments

Remote Sensing Instruments

Inheritance

1

1

3

2

!

2

2

!
I.'z.

Costing for PULSE

Category

Structure

Thermal Control

Propulsion

Attitude & Articulation

Telecommunications

Antennas

Command & Data Handling

RTG Power

Line-Scan Imaging

Particle & Field Instruments

Remote Sensing Instruments

System Support & Ground Equipment

Launch + 30 Days Ops & Ground S/W

Image Data Development

Science Data Development

Program Management

Flight Operations

Data Analysis

Cost (FY 88 Do!"a:_]_ _,

59:_Q88,16 °- qs_

II:0 °_ q38 _

41 _ 927 _0 50 &- t ;

62,614,609 _7

64;098 191 33

13.04 _ 019 g_

24.,500 _,_ 53

170,4_4 _._-= _, J _ - 0

oo, I=, 302 <4

480., 062 53 = _

57:185,ge8_ ",o.

6,957,007 4_

II ,487 ,7_'3_ 40

1 _ 3g _, _g7 .4

458 -_ _

115,984,760.70

TOTAL 1,704,!92,542.00

.. . .... ,. ,._

OF POOR QU_kiTy
selections affect the amcunt of fuel needed ,and the desigr cf -..he

structure.

1.8 CONCLUSION

Within the mission management, planning, and cost suh_7:_t÷::

many important selections were made. The PULSE mission is a

flyby with a mission duration of 16.005 years. The launch £st_

is January 30, 2003. PULSE is scheduled to arrive at Plut_ .bn

February I, 2019. This mission requires an 8.606 delta-v from

parking orbit.
APPENDIX |

Fiscal Year '77 to Fiscal Year '88 Conversion:

(Total Cost)(FY88 dollars)/FY77 dollars : Total Cost for _::

Fiscal Year 19.78
REFERENCES

Prussing, John E. and Conway, Bruce A., Orbital Mech_nirs.

University of Illinois at urbana-Champaign, !98S.

U.S. Bureau of the Census, Statistical Abstracts cf the U.$.,

1979.

U.S. Bureau of Labor Statisics, Monthly Labor Review, J_nu_ryJune 1989.
01_ POOR Qu^u[Y

PULSE ATTITUDE AND CONTROL SYSTEMS (AACS

i. INTRODUCTION

Pulse is a three-axis stabilized spacecraft ut iizing s_iid

state sensors and reaction jets to provide control moments. The

control hardware utilizes advances in microprocee_ _ccu_-e_

capability, reliability and efficiency.

2. AACS FUNCTIONS

For the purposes of identifying

main mission phases are distinguished.

associated AACS tasks are listed below.

AACS requirements, three

These ohase_ and thai

GEOSTATIONARY EARTH ORBI T (GEO}

The launch vehicle and upper stage will insert PULSE :=t _

GEO. During this phase the deployment, of the. _oms_ th_

spacecraft attitude, and it's insertion into it'_ _nte,- D_net_-_

trajectory _ill all be controlled _ro_ the gr_,jn_ via tn_ _o_

gain antenna.

CRUISE PHASE

During the cruise phase oq the mission. L_ _

determination and control oe the spacecraft attitude w_l_

autonomous. The main spacecraft control requirement

maintaining the antenna pointing within one degree o_

O_

is tt-,a_: o-"
0_' _i'_O_ Q LIALiiY

the sD_cecraet prog,'esses along it's trajectory. Thie_ tas_ ca,-

be viewed as a continuous maneuve_ o e !o_ a_7gular rate o,- as

stabilization oe the spacecraet in a non-inertial reSerence

Frame.

ENCOUNTE_ PNASE

The accuracy required oE the AACS is much greater ac it no_

must comtrol the scanning of the scientific instr,jments. The

antenna pointing requirement must be maintained botm during ant

after the encounter while stored data from the science

instruments is transmitted to earth.

3. DESIGN OF AACS

THe primary movers ir desion of attitude determ_na_i_r, ape

control systems are ,-eliabiltt,x and low cost. Th_ emDha_i_ o =

current research in spacecraft attitude determination and cort,-c'

is in the area o_ control system_, where much og the fundamental

worF remains incomplete (Re_. p 714-715). Therefore, in the are_

have beem

duration.

of attitude determination,

alight tested

is maximized. Some oe

integ_atir,g gyros and servomotors

use o_ of _ the ehe!_ comDo,ne_t_ tm_*

on tnte_plane_ar_ missio_ ?= ]o,_n

the compone_nts, such as ,-at=

,,ill b_? d_rect ], Impi_m_nt_?i

alread,' unde _ developmen _ _!! oe uttii_d

develooing techrolog,

i_ sol ic _tat£" t=_,cb_olo_:. _ . ....

ir, teg-gteC i_to :" ight testec' _ttittJdc determir,-3=io, _ =.,,,_:- .=_n,:.
.... _j

_Ref. _). The rapid advances in mi_roprocesso_ technology tr,at

have taben place since the design o< the last in;ermla_e;ar

probes will also be ma_e use of. Modern microprocessors once

space hardened, will permit the implementation of _oqtro! law_

which greatly improve performance parameters of the AACS (Ref.

3). The computing power and memory _eoabiiit_ available _,_

permit utilizatiop of arti¢icia] intellige_ce _AI_ applications

such as expert systems. While their low processinc _e_e,

precludes their u_e in low level control loops they will be

useful in the areas of system checkouts and #rouble shooting

(Ref, 4). Previous missions have employed a fault recovery

ability which monitors the system and placed the spat#traCt ir_

safe mode ir_ the event of failure. However, ground cortrc_ _vas

necessary to reconfigure and reprogram the system beFcr_ t_÷ ,

mission could resume.

diagnose the fau!t_

rectify the failure.

Ar_ expert system wo.uid me able to, not or'l,_,

but to make and im_,lement O_._-tsic',-,= t:

ATTITUDE DETERMINATION

Figure 1 is an overview of senso_ types (Ref. _). mh_

,-el#rant criteria are that the sensors chosen must be ap_! icahlc

to, th,-ee-aKis, stabilized spacecraft im ecsentric orbit_ _r,_ r,_v_

at least meOi,jm aEcuraEy. The se_Tscr_ tD be utilized c_ c2L._T

are try,? Y_, S,__r,Sensor _YSS) a_d t_ So! _ State Det_: t:.... _?_

sta- t;-_cber.

The Ya_

couples Oe,,xice (CCD) detector , Thi_ senso,- i= #as.i;,. ,mtmc_-ate_

into optico-inertial systems. It, addi t io,_ sensor_=, being

deveieped on this baseline can be radiation hardened, a_d -a_

utilize hybrid electroni__s to minimize weight and re¢_uce

dimensions. Finally it may be employed as a high sensiti,xitv sum

sensor to aim at sources be light much fainter than the sur (Ref.

6). I_ this capacity as a planet sensor it may b_ ,Jse_ tc_

generate erro_ signals to Orive the servomechanism which cont_-ol_

the instrument scamning platform.

The Sun Sensor provides only

pointing vector to the spacecraft. A

the star Canopus, near the south

the orientation be a 3.,]r_

star tracker whic_ t_acPe

ecliptic pole proviOes

additional input which uniquely fixes the space.raft attituOe.

Such sun-canopus systems have been flown on the mariner, su,-veyo,

and lunar orbiter missions (Ref. 1 pp.18_), mh÷ CCD st_- t_ac_e _

to be used features inherent geometric stability, low _c_Qe

ooeration and high reliability (_ee. 5 ). Because the, _ncui_"

displacements between the earth, sun and canopus are small a_o

the high gain antenna must be earth pointed. The optical s_ser_

must be pla_e on the antenn_ rim te avoid blocking t_ei,- eie_ o _

view.

Pate i,ntegrating gyros

integrated with the ootical

attitude measuremen_ system.

body c _ _h_ soacecra_t and or" th_ Eca _'

point _r_ ce _he sciemce instrumemt_ .

THe, gyros _il] be used re- shots term

a_nd the optical sensors will be

can be used off the shel _ and be

sensors int_ a_ ogt_co-_,_t,a]

The gyros wLl! be _i_:ev c. tb_

attitud= mecls ;,e,m ....

L_5_O fO- iO_ *erm m_2c_ "-9nE ""
and calibration of the gyros.

CONTROLHARDWARE

A high precision microprocessor implemented control system

accepts the angular displacement, rate and disturbing torque from

the sensors above. The control law produces time op_ima_

recovery from large angle errors and can obtain stable control

with disturbing accelerations approaching the control torque

The control law also incorporates fuel optimal slewing through

unlimited angles. Steady state limit cycles in the arc-second

region are attainable for precise control during the encounter

phase (Ref. 3>.

Fig 2 shows a block diagram of the control loop. The state

estimator generates a state vector consisting of angular ra'_ _

displacement and disturbance torque. The slew algori_n,_

optimizes fuel consumption. The control law controls timin_ _

jet firing.

For the PULSE mission it is required that the microprocessoralso generate the command input. This requires o_ boars_

calculation of the proper earth pointing angle at all stage_ _

the mission. Another difficulty may arise in controlling. Th_

scanning of the science from integrated gyro and acceleromet_ _-

data. A senarate planet sensor on the soon ol_tform m_ b_,

required to pro,xide

controls the platform.

a_ error signal to the servomotor _ _-_

Torquer Selection

Ther_ are, t_c types o_ torquers available fo_ a fie]_ 4r_r
0,-,,-,,-,,_, PAGE iS

OF F;_'JO__UALiTY

I

I

R

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0

0

0

0

L_

A

_U

°,lm

U.

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_J

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qI

I
OF ?OOt_ QuALI't"V

environn_ent: momentum exchange and mass expulsion. Gas jets are

the only viable alternative for missions of this duration (Ref.?>

estimates of spacecraft moment of inertia and an assumed impulse

bit of .005 s and a limit cycle deadband of 1 degree were used to

estimate total impulse required for maintaining antenna pointing

during cruise. This assumes that any maneuvering requirements

are negligible compared to the essentially continuous limit cycle

(Ref. 8)(Appendix A). The total impulse led to a trade study

among possible propellants. Cold gas, hydrazine and

bipropellants were the candidates. Bipropellants and augmented

hydrazine were eliminated because o? the required complexity°

Fig 4 shows a trade analysis for the propellants. This shows the

optimum propellant is hydrazine.

This analysis assumes a torque free environment. To chec_

the validity of this assumption an estimate _e the maximum solar

This torque was shown to be negligible whe_

control torque thus justifying the assumption

torque was made.

compared to the

(Appendix B).

Other possible errors are introduced into the analysis by

changes in thruster performance over time, propellant sloshing in

the tank, and inaccurate modeling of thrust pro?ile.
0_; ..... _r.;.. PAGE IS

OF FOOR QUALITY

Corrosive

_xhaust _ no

_dd_d

Comole_tv _o

using L=2m

_'AA

" •O0 1

yes

r_o

-hera= : deq_ee

Figure

_ug. 4ydrazi_=

_00

_.01

_/es

7=2000 _J-=

.015

yes

- = 5*E_
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r-C

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0 _'''_''' QU._LITY

OF pO0_

References

I. Wertz, J.A., Spacecraft Attitude Determination and Control_

_eidel 1978.

2. Amieux, Miller, Space Borne Attitude Measurement U_its,

Proceedings of the Ninth IFAC Symposium, 1982.

3. Dodds, S.J, An Adaptive High Precision, Gas Jet Satellite

Attitude Control System Proceedings of the Eight IFAE Symposium

1979.

4. Culbert, Boarnet, Savely AI Applications for Space _obotics

Journal, Mar. 88.

5. Connolly, et.al., Synopsis of Optical Attitude Senso_

Developed By ESA, Proceedings of the Ninth Annual IFAC Symposium

1982.

6. Flamebaum S. and Anstett, P., Multipurpose SuT_ Sense,- usin_

CCD Detector, Proceedings o # the Ninth I_AC Symposium. lq_.

?. AAE e41 Class Notes- Spring 1990.

8. Hove, L.E. , Design o_ Unmanned Automated Spacecraft fo _ The

Grand Tour, Spacecraft Systems, Pergamon Press_ !970.
0_." ,_-_..,.,, C_!_._,_i._T,{

3.0 Science

3.1 Mission Objectives

The primary objective for this unmanned, scientific st_idy of

Plutonian space is to expand upon our currert knowledge o_ the

Pluto-Charon system. This will be accomplished by obtai_i_g and

returning information concerning our three scientific ohject_'.,e_

which are listed and prioritized in Table 3.1. Each of the£e

objectives will be investigated through the use _ the, _ULS_

Experimental Package and the radio science equipment aboar'd the.

probe.

Table 3. 1 !

I

i

Scienti?ic Objectives o? the PULSE probe i

l. Investigate Plutonian Characteristics i

f

2. Investigate Satellite Characteristics i

......... i

3. Investigate Planetary and Interplanetary Particles _nd !

fields. !

_......... i

The investigation of each of these scienti_i_ objective_ is

the major concern of this mission. Since no probe has visited

Plutonian Space, little is known about the planet Pl_tc _r _ts

satellite Charon. However the scienti?ic community has c_dtlcted

recent studies concerning the Pluto-Charon system. The ;_n_ledge
OR'.G]_AI- pAGE IS

OF pOOR QU_,LITY

gained from these studies was one of the determining factors for

instrument selection aboard the PULSE probe. Althouoh the_e

studies have given us some new information, none of the

information can be considered conclusi, = until _ _!oser

investigation is conducted.

3.2 Science Objectives

3.2.1 Plutonian Characteristics

One characteristic of Pluto which must be investigated is

the atmosphere. Astronomers have found that Pluto does hay9 ._

dilute atmosphere which extends several hundred k_Io(neter_ abo,_

the planet's sur'?ace(Ref.2, p.45). This complex atmosphere _

believed to contain heavier molecules than methane which w_

previously believed to make up the entire atmosphere(_e_._.

p.326). Other atmospheric properties which must be _nvest_gat_d

include, measurements of temperatures and pressures at varicu_

altitudes and cloud characteristics (if present).

A second characteristic which needs inves_igat_or, _s th_

surface characteristics of the planet. Earth obse_-vat_ons have

shown the existence of polar ice caps at the poles of Plu_o wh!c_

are believed to be composed of methane ice (Ref.13, p29). This

possibility along with other surface features need investigation.

Other areas of interest include, mass, shape, density, orbit

characteristics and composition. By investigating these _reas,

we hope to gain improved knowledge of the planet Pluto.
ORIGiNaL PAGE IS

OF POOR QUALITY

3.2.2

Pluto is

satellite named

studied are

Charon Characteristics

believed to have only one orbitin_ _atur_'

Charon. The characteristics which re,_d t_ h_

relatively the same ones found in the previous

section. One difference is that the amount o? methane _ Charo_

is believed to be much less than on Pluto. Charon is believed to

be composed of water ice and not methane ice.

3._.3 Planetary and Interplanetary Particles and Fields

One interesting area which falls urder this category is the

gravitational and magnetospheric interactions of the Pluto-Char_,,

system. Charon is relatively large compared to Pluto. Tt _

because of this that the Pluto-Charon system was thought to b_

one planet which led to incorrect measurements. Ther_ is _

other planet-satellite system known so it seems very impgrta_t t_

study these interactions.

Other areas shall include investigation in; charged particle

environments, wave particle interaction,

rays.

The

located

measurements

solar wind and cosmic

instrumentation used in most of these measureme_ts _.s

on the probe's scientific boom _hich al!ew_ _

in the interplanetary environment as we!l ,is the

planetary environment.
OFPOORouAu 

3.3 Pulse Experimental Package

The Pulse Experimental Package(PEP) will consist _f ?_v

remote sensing instruments and four particle _d e_e!

instruments and radio science. Each of _hese instruments ;

listed in Table 3.2.

power specifications.

Also listed in this table are m_ee ar_

The total PEP weight is approximately qq_

kg and the approximate power they consume is _0 W. The select_o

of these instruments was based on their ability to i,nvestig_t

the scientific objectives.

3.3.1 REMOTE SENSING INSTRUMENTS

IMAGING SCIENCE SUBSYSTEM

The Imaging Science Subsystem(ISS) _as selected becaus9

has a much higher resolution (I024 x I024 pi×els) than a_y _f i _

predecessors(Ref.5, p.9). Many of the instrument's c_mpone_

are just improvements upon the camera systems of its ancestcr_

This instrument also offers data compression and storage _hic

will be necessary because of the large amount of data that _[

be obtained during our flyby of the Pluto-Charon system sin_

most of the investigation will be carried out at thi_ time. TI

data rates of the ISS are selectable.

to 350 kbps(Ref.5, p.[O).

The ISS offers the opportunity

system. The characteristics

investigated with the ISS.

to

of Pluto

They range ?tom 6.2 kb;

view the O_uto-Char

and Charon _,_i!!

We also will be able to investiga
NAC

Type:

Focal Length:

Focal Ratio:

Spectral Range:

Resolut ion:

Coverage:

WAC

Type:

Focal Length:

Focal Ratio:

Spectral Range:

Resolution:

Coverage:

Table 3.3

NAC and WAC Optics

Ritchey Chretien with three field correctors

2000 millimeters

f/10.5

200-1100 nanometers

The resolution per pixel will be six microradians

square.

The field of view will be 0.35 degrees square.

Refractor

250 millimeters

f14.0

350-1100 nanometers

The resolution per pixel will be 48 microradians

square.

The field of view will be 2.8 degrees square.

-Ref. 3, p. 8
C_!;_,_._L PAGE IS

OF POOR QUALITY

the Pluto-Charon interactions and determine other

be of interest in Plutonian Space.

This instrument, which is essentially the same am the ISI

that will be flown on the Cassini and CRAF missions _chedu!ed _,

be launched in t975 and

Narrow Angle Camera(NAC)

areas that ma_

cameras will have a spectral range which is exter_ded visible as

they well operate at a temperature slightly below too,

temperature. The components of these two came_a_ include a du_

cover, hood, optics, filter mechanism, shutter detector head an

radiator. The dust covers are a method of protection for th

optics which will be motor activated. The hood is designed t

also protect

parameters for

3.3(Ref.5, p.9).

The filter mechanism

Hubble Space Telescope.

had a maximum of seven

the optics and reduce the glare. The opt_ca

both the NAC and the WAC are listed _n Tab]

of the cameras was derived _rom th

Unlike Galileo's _ilter mechanism the

positions, Pulse's _ilter mechanism has

maximum of 36 positions. The two filter wheels of the NPC ar

the WAC contain 22 filters and 14 _ilters respectivel¥(_mf_

p.10).

The shutter technology oriented _rom shutters on Voyager ay

Galileo. It consists o_ a dual blade focal plane which m_

operate in either direction. The lower limit on exposure time j

.005 seconds and no limitation on the upper limit. One advanta_

of this system is that both shutters may be activat_

simultaneously (Ref.5, p.9).

Th

1996, is composed of two cameras,

and a Wide Angle Camera(WAC).
The detector head of the ISS contains the

Device(CCD), driver, thermal control unit and

circuits. This electronic module is common to both the NAC

the WAC. Other components o?

microcomputer 2) memory 3)power

5) image data multiplexer 6)

Charge Coupled

s;gna! cha_m

and

this module include: I) a

supplies 4) engineering _ens_r_

square root proce_or _) _ma_

memory 8) image data compression 9) bus intereace unit(Fee.5,

p.lO).

The radiator of the ISS is responsible for cooling the C_D

to temperatures approximately -80 degrees Celsius(_e?.5, p°).

NEAR INFRARED MAPPING SPECTROMETER

Th(e N_r In_r_rmd Spmctr_meter (NIMS >

i _truments that is aboard the spacecraft

i S One 0"_ 1⁄2k,_

Galileo. T_is

water vapor

respectively (Ref.8, p.207).

The objectives of NIMS fall into

objectives. NIMS will be used for

geological properties of both Pluto

instruments unique ability o? combining spectroscopy and image_y

in one instrument makes it a prime candidate ?or PEP. Another

reason for its selection is that it can monitor beth methane an_

which are believed to be present on Pluto and Charo_

the first t_o scie_nti?ic

both the investigation o_

and Charon. N!_S _i!i

accomplish this objective by investigating surface ?eature_ a_d

surface composition through surface mapping and in_rared spectral

investigations.

NIMS will also investigate atmospherical properties. Goals

of this investigation include information about atmospheric
ORIG!N,,_L ;:'AGE IS

OF POOR QUALITY

structure and composition. Also investigations about the

existence of clouds, cloud properties and temperatures at various

altitudes will also be conducted. Table 3.# lists a summary _

specifications for this instrument.

The NIMS will be placed on the scan platform. It is

protected against contamination by covers and heaters. ;t slso

has a passive radioactive cooler which will keep the _nstrument

at is operation temperature of 80 K(Ref.i_ p.2Ot).

PHOTOPOLARIMETER-RADIOMETER

Photopolarimeter-Radiometer(PPR) was

flown on the Galileo spacecraft. It

because of ability to measure its

p_|_ri_t|_n o_ mcattered sunlight in the spectra] region

also an instrumen _

wa_ selected Dr_mar_Iv

intensity and l_near

_her_

methane strongly absorbs radiation(Ref.19, p.l_8).

unique because of the combination o_ three separate

it may conduct; photometry, polarimetry and radiometry.

The objectives of this instrument is as described

measure the intensity and linear polarization _f

sunlight in the narrow spectral bands.

Another objective of the PPS is the measurement of therma

infrared radiation. This may only be investigated i_ clouds d_

exist in the Plutonian atmosphere since the radiation is believe

to be emitted primarily from cloud particles.

Some atmospheric properties well also be investigated, Thi

experiment is mostly concerned with the part_cl_ in th

atmosphere and their distribution.
Angular Resolution:

Angular Field:

Spectral Range:

Spectral Scan Time:

Telescope:

Spectrometer:

Detectors:

Signal-to-Noise:

Mass:

Power:

Date Rate:

Data Encoding:

Table 3.4

NIMSInstrument Characteristics

0.5 mrad x 0.5 mrad

I0 mrad (20 pixels) x 0.5 mrad (I pixel)

0.7 - 5.2 micrometers

4-1/3 seconds (20 pixels, 204 wavelengths)

23 cm diameter f/3.5 Ritchey - Chretien

wobbling secondary for spatial scan,

800 mm equivalent focal length

40 lines/mm plane-grating spectrometer,

f/3.5 Dall Kirkham collimator f = 400 rml,

f/1.86 wide-angle flat-field camera

f = 210 ran

InSb (15), Si (2), discrete elements,

quantum efficiencies = 70-80%, noise

equivalent power = 10 -14 watt,

D* = 3 x i013 cmJ_watt-i

i00:i (0.075 albedo surface at 3 micrometers)

18.o kg

12 W (average), 13 W (peak)

11.52 kbps

i0 bits

-Ref. i, p. 201
ORIGfNf, L ;_-",_._:_IS

OF POOR QUALITY

There are several different channels for the PPS the

"polarimetry channels are centered at #100, 6780, and _u50 i_Id

the photometry channels are centered at 6180_ 6330_ 6a60_ 9_80,

8300, 8410, and 8920 angstroms. When the instrument is used eor

radiomet-y the in_rared channels are centered below # micromet_r_

at 17, 21, 27.5, and 3?.5 micrometers, and above 42 misrometers."

(Ref.19, p.129)

There are

r_d|ometry m_d_.

two operational modes, a cycle mode and a

Th# ¢y_i_ mod@ rotatem the _|_e_ _h_e!

allowing each channel to transmit at least once every !8 seconds.

The radiometry mode rotates the infrared filter wheel bach a_d

forth.

The PPS weighs 4.8 kg and has both a replacement heater a_)d

a sunshade as safety features(Ref. 19, p.l_9).

ULTRAVIOLET SPECTROMETER

The ultraviolet spectrometer

the composition and structure of

satellite Charon.

was selected _o_- determining

the planet Pluto and it_

A secondary objective of this instrument is to determine tke

properties of the upper atmosphere. Although P!uto'_ itmcsphere

may not be as large as that of 3upiter, there is a poss_bi!ity of

molecular absorption features and auroral zone emissions that are

believed to be common among planets with large atmospheres.

Through airglow and occultation modes we hope to detern_i_e both

the atmospheric structure and the atmospheric composition.

This Galilean successor will consist of a 250 mm-aperture
C_c '_ .... _

_i_,_ _.i__ _°_Ji__

Cassegrain telescope, a 125 mm focal lergth Ebert-Fast_e

monochromator, three detectors and control logic. The UVS weighs

approximately 4 kg and consumes 5.33 W. The wavelengths covered

by the UVS range from 1100 to 1400 angstroms(Ref.19_ pp.130-'31).

The UVS also has flexibility. It may take data at a _i_ed

wavelength or it may change the wavelength every 0.0007 secured.

It is not limited to these two modes, however. Other variations

may be programmed into the microprocessor of the UVS

(Ref.19, p.131).

3.3.2 P_RTICLE AND FIELD INSTRUMENTS

MAGNETOMETERS

The magnetometers that were selected for PEP are actually

_elected because o_ their _billty to measure field_ _._n,_,_,g fr_,_

0.006 gamma to _0 G(Re_.4, p235). This wide range _ _ield

measurements will be needed to measure the fields in both the

Plutonian and interplanetary environments. The fact that th_

PULSE probe is three-axis stabilized, like V_yager, _l_o gives

reason for this selection.

The magnetometers that have been selected are two L_w =ield

Magnetometers(LFM) and two High Field Magnetometer_(HFM). This

redundancy makes the system reliable in the event that one of the

magnetometers does not function properly. The magnetometers

purpose is to study the planetary _nd interplanetary _art_cles
and fields.

l)

interactions.

2) Measure the magnetic field o? Pluto and Charon.

3) Measure interplanetary magnetic field_

4) Determine magnetospheric interactions _th s_lar

wind,

cosmic rays and plasma waves.

5) Use observations to make further abservatio_s.

6) Search for interaction between interplaneta_-_'

and interstellar media.

The LFM and the HFM are located on the particle and field

boom.

These objectives are described as ?o!!o_:

Investigate Pluto-Charon magneto_phe_ic

The placement of these magnetometers w_!! be

proportionately the same as the ones on the Voyager mission, s.

There will be one LFM located at the outboard end of the boom ar,d

be placed approximately at the _enter _f th_

will be located near the inbgard end o? th_

This placement ailo_ ?_,-

due to the _p_cecra_ '__

the other LFM will

boom. The two HFM

boom approximately one meter apart.

some measurement correction factors

magnetic field(Ref.4, p.247).

The range of the measurements as state earlier is _ai,-1"/

large. The LFM range is !8.8 gamma to ±0.50 G and the HFM ,-ange

is ±0.50 G to ±20 G with uncertainties of ±2.2 mil!igamma to

±12.2 gamma and ±12.2 gamma to ±488 gamma respectively. _his

total ±20 G range has a 12 bit digital reso!ution(_e? u, p._36_.

As the probe increases its distance from the sun, the data

rate will not vary greatly because of the data compaction modes

of the instrument(Re?.4, p_54}.

COSMIC RAY DETECTOR SYSTEM

Like the magnetometers of the PEP, The Cosmic _ay Detect,o_-
ORIGII'_AL PAGE IS

OF POOR QUAL/TY

System(CRS} selected for _EP has also flown on the Voyage_

missions. This instrument was selected because Earth-base_

observations show that something is blocking the light during

Pluto's occultation. There are beliefs that this "e×ti_ction

layer" is produced by particles which originated from cosmic

rays(Ref. 13, p.29_. Therefore the CRS investigation f_ay e_hanc_

our knowledge of both cosmic rays and the component_ of _he

Plutonian atmosphere.

The CRS objectives fall in the category of Rlanetary a,_d

interplanetary particles and fields. These objectives may 5_

almost exactly compared to those of the Voyager C_S o_je,:_v_s,

There only difference is the planet that is being tar_ete_.

Below is a list of the objectives of the Voyager missio,_ er_m the

Flight Science Office Science and Systems Handbook with t%_.

appropriate modifications for the Pluto mission.

an

3)

origin,

dynamics of

contribute

i) Measure the energy spectrum of electron_ 3-I_0 _e_?

2) Measure the energy spectra and elemental _mpo_it]or_

of all cosmic ray nuclei from W through c_ r_v_r

energy range from approximately 1-500 Me'?/nu_.

Provide information on the energy content,

acceleration process, life history _nd

cosmic rays in the gala:_/ and

to an understanding of th_

nucleosynthesis o_ elements in co_mi_ ray so_r_=es.

4) To provide information on the transport _

_osmic rays, Plutonian electrons a_d !cw

energy particles over an extemded

region of interplanetary space.

5) Measure the three-dimensional streaming patterms o _

the nuclei from H through Fe and elect, on_ over

an extended range.

&> Measure particle charge composition of the

magnetosphere of Pluto and Charon<Ref.17, p4._)

One may say that these objectives, inherited from the Voyager_:

are still of great importance to th_ scientific community.
OF pOOR QUALITY

The CRS is composed o? three system_; the _ig_ Energy

Telescope System, the Low Energy Telescope System and the

Electron Telescope System. These three system_ share s_c_e zc_,mcr_

electronics and are responsible for

nuclei charge and energy spectra may be determined

instruments for elements with atomic numbers from

the above objectives. The

by bhes, _

energy ranges o_ 1 MeV to 500 MeV for H and 2.5 MeV to, 500 MeV

for Fe. For isotopes the range of atomic numbers is I to _ _^_[th

an energy range of 2 MeV/nuc. to 75 MeV/nuc. Finally_ the raq_e

of atomic numbers of anisotro_ies is 1 to 26 _ith an ene,-gy range

of 1MeV to 150 MeV for H, 2.7 MeV to 500 MeV ?or me and _ t_ i0

MeV for electrons (Ref.4, p.365).

_LASMA INSTRUMENT

The Plasma instrument(PLS) that has been selected ,_as ?_c_-

aboard the Galileo Spacecraft. It wa_ selected because o? _t_

energy/unit charge and the decreased temporal resolutions £or

obtaining electron and positive ion spectra. The plasma

instruments of the Voyagers and the Pioneers don't even app,-oath

the values o? the PLS.

The objectives of this mission are also of the partt_!e ,_d

field type.

properties

intensities

particles.

The PLS is composed of the following:

I> T_c_ electrostatic analyzers that measure th_

energy/unit charge of electrons and positive t_n_

These objectives include measurements of the plasma

in solar wind, assessments of composition_ energy,

and three-dimensional distributioe o_ ]o_ enerQy
ORIGINAL PAGE IS

OF POOR QUALITY

2) Seven sensors that determine electron intensities.

3) Seven sensors that determine positive _

intensities.

4) Three mass spectrometers that determine the

compesition of ions(_ef.19, p.133).

The PLS capabilities range from I Vto 50,000 _' in 64

di_?erent passbands. The PLS also contains soe_t_are wk, i_

permits ground command alterations to the instrume,_ts comm_nd_.

The instrument weighs la kg and will be mounted or_ the _,_-_r_c_'

boom of the PULSE probe(Ref.19, pp.!33-135).

ENERGETIC PARTICLE DETECTOR

Another instrument selected from the Galiiean payload is th,=

Energetic Particle Detector(EPD>. It _as selected becaus_ c£ the?

need for measurement of high energy particles i_ th_

magnetospheres of Pluto, Charon and interplane_a_-y space.

Although the PULSE probe is three-axis stabilized_ _e shoulcl

still be able data about t_,e h_h

energy even without sweeping

to obtain a great deal of

electrons, protons and heavy ions

motions.

The EPD

Magnetospheric

is made up of two subsystems,

Measuring System<LEMMS) and

a Low Enecgy

a Cc,,_p(_'s _.t i c r '

Measuring System(CMS), formed by two separate telescopes(Re_.!_

p.136).

The LEMMS consists of two components. The first compo_e_t _s

an ion telescope _ith two solid-state detectors. One detectc_r_

the low _ield detector covers an energy range o_ 0.02 MeV to 3.4

MeV. The other detector will be used for the de_irition o _
O_,IG_NAL PAGE IS

OF poOR QUALITY

additional electron, protonp and alpha pa_ti_le cha_ne!s. The

second component of the LEMMS is a magnetic electron __pectrometer

with two detector pairs. These detector pairs span _ rar, ge of

0.015 MeV to O.eO MeV and 0,I0 MeV _o 1.0 MeV(Rer_!9, 2.136).

The CMS components will be used for the measuremer_t o _"

composition, energy spectra and pitch angle distrib,.:ti,._:_,s 9-r _e

high energy ions. These components

nine detectors(Re?.19, p.136).

The EPD weigh_ 9 kg and will also

boom(Re?.lg, p.6).

are the CMS be!esccpe a_d

be !ocated on the sc_ep,.-,_

PLASMA WAVE SUBSYSTEM

The last particle and field instrument is the _l_sma Wave

Subsystem(PWS). The PWS was selected because o? the _mport_nce

of plasma wave investigations.

These investigations include wave particle irtera_t_?_s =,_d

their e_fects on the Pluto-Charon system and measure_,_ c_-_

spectral characteristics of electric and magnetic fields _r_ tX,e

range of 5 Hz to 5.65 MHz.

the difference between

waves(Refl9, p.137).

There are two sensors of the PWS.

electric dipole antenna

elements mounted at the

We _ill also be able to distinguish

electrostatic and e!ectromagnet {c

The first is a 6.6 meter

which has two tapered graphite epo×,/

end of the magnetometer boom. Tine other

_ensor is a search co_l magnetic antenna. This antenm_ consists

of two high-permeability rods, 26.6 and _7.5 cm lonc. The !o_v

frequency search coil has a winding o_ 50,000 turns oe 0,07 mm
ORIGIMAL PAGE IS

OF pOOR QUALITY

diameter copper wire and a frequency range _f !0!4z to _.5 kHz,

This search coil must be mounted parallel to the electric

antenna. The high frequency antenna has a winding _? 2,000 t_rns

of 0.14 mm copper wire and a frenuency range oe ! Nz to 50 _Hz.

This search coil must be mounted perpendicular to the elect-to

antenna. There will also be a preamplifier mounted nea,- %h,?_

search coil to provide a Io_ impedance to the electronics(_e?.lg,

p.136).

The processing of the signal received from the set, sots ma>"

be processed by a low-frequency spectrum analyzer, a modicumfrequency spectrum analyzer, a high-frequency spectrum a_alyzer

and a wideband waveform receiver. The fastest measurements are

provide by the wide band waveform receiver(Ref.19, pp.!36-!3?).

3.4 CONCLUDING REMARKS

The Objectives in this subsystem report are by no means the

only investigations that will be conducted. There are _ndeed

some that _ere not mentioned and some that _i!l not ,nate,-_ize

until a probe visits Plutonian _pace. The purpose c,_ *h!_

mission is to observe as much as possible so as to enhance our

knowledge for further scientific investigations.
OR|G!NAL PAGE IS

OF POOR QUAUTY

3.5 REFERENCES

!. Aptaker, Irving M. "A Near-Infrared Mapping Spect_-ome_,

for Investigation of Jupiter and Its Satellites." _E?C_e_Os !AE

SP!E__Imagiog_Spectroscop__I_. Vol. 834. August 19_7, 196-_!_.

"Barren Planet Bares its Air."

3. Beynon, W.J.C. et al.

November 1977, 77-232.

4. Beynon, W.J.C. et al.

December 1977, 215-_76.
OR1G!;_P,L _.AGE IS

OF POOR QUALITY

5, "Cassini Mission: Saturn Orbiter Proposal

Package." Vol. VI. JPL D-6464. 10 Dctober 1989.

In?ormatioll

6. Corliss, William R.

D.C.: NASA, 1971.

1_ash ington

7, Eberhart, J, "Pluto's Atmosphere:

Science News. Vol. 135. 27 May 1989, 326.

More than Met_ a:-e.

8. Eberhart, J. "Pluto: Limits on Its Atmosphe,-e_ Ice o_ its

Moon." Science News. Vol. 132, No.!3. e6 September 19_7, 207.

9. Eicher, David J. "Pluto Emerges from the

Astronomy. Vol. 16, No q. September 1988, 52-53.

Shadow, s, "

10. "Galileo Represented

Week and_Space Technolog2.

Peak in Design Complexity."

10 October 1989, 77.

Aviat_o,_

11. "Getting Small." _@O_ifi___m_Ei_O. April !9_7, 68.

12. Nicks, Oran W. "A Review of the Mariner

SP-130. lq67.

TV _esults." _!A%A

13. "Passing Star Brings Pluto

_.i_9_!_.t. Novenber 1988, 29.

Icecaps to I_ight." N@_

14. "Pluto: Evidence ?or Polar Caps." Sc!ence_News. Vol. 13u,
ORIGINAL PAGE IS

OF POOR QUALITY

No. 9. 3 September 1987, 156.

15. "The Pioneer Mission to Jupiter." NASA SP-e68. !971.

16. "The Pluto that

September 1988, 43.

Came in ?tom the Cold."

17. "Voyager Neptune/Interstellar Mission: Flight Ss_e_ce

O_ice Science and Mission Handbook." J_L D-498, 1 October _oe?,

18. "Voyager to Jupiter and Saturn." NASA SP-U20. 1977.

19. Yeases, C.M. et al. "Galileo: Exploration o? Jup;ter's

System." NASA SP-479. 1985.
4.0 ............ Introduction to command, communication, and

control

4.1 ............ Antenna System

4.1.1 ............ High- gain Antenna

4.1.2 ............ HGA trade- offs

4.1.3 ............ Look at Optical Communication

4.1.4 ............ Size of HGA

4.1.4.1 ......... Deep Space Network

4.1.4.2 ......... Encoders and Modulators

4.1.5 ............ Amplifiers

4.1.6 ............ Radio- frequency Subsystem

4.2 ............ On- board Computers

4.2.1 ............ Lag in Technology

4.2.2 ............ Performance Characteristics

4.2.3 ............ Space Qualification of Computers

4.2.4 ............ Computer trade- offs

4.2.5 ............ Problem with Galileo

4.2.6 ............ Data Management System

4.3 ............ Conclusion
4.0 Introduction to command, control, and communication

The command, control, and communication subsystem has

several design requirements which include:

I) minimization of cost and weight

2) maximization of performance of reliability,

performance, and simplicity

3) use of off-the-shelf hardware

4) use of technology before 2000

5) application of AI, if applicable

6) sufficient life time to carry out the mission

The priority that overshadows all of them is cutting the

cost of the mission. As far as incorporating new technology into

PULSE, we are taking a conservative approach. Proven designs will

be chosen over new technology, except in the case where it would

be more cost effective to use the latter. When possible, past

deep space probes will be used as a prototype due to reliability

and cost requirements.

4.1 Antenna System

Reliability is the dominating factor when discussing

antennas. Voyager 2 and Galileo will be used as the prototype for

this subsystem due to the fact that proven techniques enhance

reliability and lower the overall cost of the vehicle. A

high-gain circular parabolic antenna will be used because this

shape optimizes the gain. A low-gain antenna will be included

mostly for communication when near earth for attitude articulation

and control reasons, since the high gain antenna can not be used

these ranges.
4.1.1 High-gain antenna

The high-gain antenna (HGA) meets all of the

requirements stated in the RFP. HGA's are the most cost efficient

antennas because they use off- the- shelf hardware. They are

reliable and their performance is well known because they were

used in many previous spacecraft and are based on already proven

technology. This antenna was chosen because it meets all of the

applicable requirements.

4.1.2 HGA trade- offs

The most important trade- off in HGA's is the powergain tradeoff. Gain is increased as the antenna size is

increased, this also result in a higher weight. If more power is

needed the weight also increases because the weight of the RTG's

must be greater. This is accompanied by the requirement of

minimizing the weight of the antenna. The maximum of power- gain

trade- off occurs when the product results in minimum weight.

4.1.3 A Look at Laser Communication

Optical communication could result in 47 bps from 50 AU

from a mass of one kilogram. There are many reasons that this

technology cannot be justified given the requirements from the

RFP. Optical communication is in the high- risk department as of

now because it has not been deep space tested yet. Plans for

testing are planned but it is doubtful optical communication will

be ready for deep space missions before the year 2000. This

antenna would also require that a 20 m receiving antenna be put in

orbit, since optical communications have a severe limiting factor

of weather dependence.
4.1.4 Size of High-gain Antenna

The size of the high-gain antenna is going to be 2.5

meters in diameter. This is the maximum size that the launch

vehicle will allow. This is smaller than either Voyager or

Galileo, which are 3.66 and 4.8 meters in diameter consecutively.

This decrease in size can be accounted for in several different

ways including increase of gain in the antenna, improvements in

the Deep Space Network (DSN), and improvements in the encoding and

decoding of data.

4.1.4.1 DSN

The DSN applies the technique of antenna arraying.

includes many large antennas from all over the world.

LOCATION

GOLDSTONE

DISH X-BAND

SIZE REC'V

34m YES

70m YES

34m YES

It

V.L.A. 27x 52m YES

CANBERRA 34 m YES

70 m YES

34m YES

USUDA 64m NO

PARKES 64m YES

MADR ID 34m YES

70m YES

34m YES
Possible improvements to this network include changing

the Usuda antenna so it is capable of X- band reception.

Increasing the size of the 64 m antennas to 70 m. Adding a 34 m

antenna at the Parkes and Usuda location would add i.i db each.

General Electric has suggested that the masers be replaced by

high- electron- mobility transistors, which would cost a third as

much to operate and a quarter of the implimentation cost. These

improvements could led to 3-4 db increase in gain.

4.1.4.2 Encoders and Modulators

The effectiveness of digital satillite communications

systems (DSCS) will increase when well chosen modulation and

noise- immune encoding methods are used. The PSK-4-CC was found

to to be a good method. Both the frequency effectiveness and

energy can be increased. Power gains may reach 5 db and specific

rates can increase by a factor of 1.5. From a costing side,

increasing the efficiency of the encoder is less expensive than

increasing antenna size or transmitted power, or increasing the

receiver noise sensitivity.

4.1.5 Amplifier

The amplifier used will very from the one in Voyager 2,

but will be similar to the one used for the generic Mariner Mark 2

(MM2) design. This design includes the use of gallium arsenide

field-effect transistors in the amplifier to produce an output of

5.6 W. This value could be raised to about i0 W with only minor

modifications. This application of solid state electronics would

cost less than half that of the system used in the Voyagers which
featured traveling-wave-tube-based amplifiers.

4.1.6 Radio-frequency Subsystem

PULSE's high-gain antenna will maintain communication

with Earth in only X- band, as in the case of CRAF. S- band

communication was used in the Voyagers because not all ground

stations could not handle X- band when they were launched. Now,

all stations except the Japan based antenna are capable of X- band

communication. X- band offers better range and range- rate

measurements, and greater immunity to charged particle

interference. Using only one band simpifies the ground system and

lowers the operational costs.

4.2 On- board Computers

Radiation- hardened versions of widely available

microprocessors and integrated- circuit chips supported by wellknown software development tools. Handling of scientific data

during and after the mission must make use of the latest

technology.

4.2.1 Lag in Technology

The computer industry is one of the most rapidly

developing industries. There has been a problem with computer

systems in past spacecraft due to the lag in technology because of

this rapid development. This is difficult to avoid because of the

time delay between deciding on a system and the actual launch

date.

4.2.2 Performance Characteristics
The PULSE probe will be outdated by the time it is

launched, as in the case of all spacecraft, but on- board

computers need to be selected about five years in advance to

develop, test, and integrate the spacecraft subsystems. A

schedule and summary of major features of the PULSE computer

system are listed below.

Launch date 2003

Year computer selection made 1993

Year commercially available 1990

Difference in launch and avalable 13

Microprocessor

Performance

RAM

32 bit

4 MIPS

4000 kbytes possible

4.2.3 Space Qualification of Computers

The problem with spacecraft computers is that they must

be able to withstand radiation and the bombardment of high-energy

particles, and operate in a highly reliable manner. NASA,

Defense, and the Department of Energy are working to develop and

deploy space qualified computers.

There are several space qualified computers. Sandia

National Laboratory is developing a set of advanced 32- bit and

16- bit microprocessors called the SA 3300 family. The

microprocessor and its associated computer hardware should be

available in about four years. There is also a generic version of

the 32- bit processor RH32 which will be fully developed soon.

4.2.4 Computer trade- offs
Because of size, weight, and power limitations onboard computers must be small in size, lightweight, and have low

power requirements. Selecting more advanced computers for the

spacecraft can result in higher development costs, but the overall

result is lower overall life- cycle costs of space missions

through lower software development and maintenance costs. This

can be further decreased when a universal higher level languages

are approved for space programs. The Department of Defense

approved Ada recently. The advantage for this standardization is

lower cost, lower development risks, shorter delivery schedules

and ease of maintenance. To date, assembly language source coding

has been used for spacecraft data processing. Sufficient support

software should be available by the time PULSE is launched. The

emphasis will turn from hardware to software to control the

spacecraft. By putting all the sophisticated logic in software,

much less hardware is needed and designers have the flexibility of

reprogrammability.

4.2.5 Problem with Galileo

NASA used a RCA 1802 8- bit microprocessor which caused

problems due to the limited capabilities. Its relative low speed

and its limited memory increased cost because of problems with

writing efficiency and maintainable software. The 32- bit

processor in PULSE will allow expanded mission objectives such as

acquiring and relaying more pictures faster, and allowing more

autonomous operations. While scientific objectives could be

reached with a less modern computer, lower cost and risks

encourage its use.
4.2.6 Data Management Systems (DMS)

The DMS must regulate power management, command and

telemetry, thermal regulation, and antenna control.

centralization of the DMS ensures command prioritization and

synchronization of resources. Using separate microprocessors and

spares can result in power, weight, and code complexity to provide

the necessary redundancy. The DMS may make use of a internally

redundant Intel 80386 for data processing and automatic control

purposes. The only problem is that it is not radiation hardened

yet and may not be by the year 2000. If it is not a back-up

option would be a 32- bit radiation hardened mi_roprom_ssor

combined with a direct memory access chip that simplifies software

which is being developed by JPL.

The DMS will be similar to the ESA probe ISPM include

a Central Terminal Unit (CTU), Remote Terminal Unit (RTU), Command

Decoder, and data storage ( a tape recorder or hard drive ). The

CTU controls the automatic functions and operations. The main

tasks will be performed on the Intel 80388 microcomputer. The

f

software governing articulation and control is based on the Ada

language. The CTU contains a fault detector which will switch to

redundant units when problems arise. The command detector that

will be used is the NASA standard which is upgraded from the one

used in Galileo.

4.3 Conclusion

The most important features of this subsystem is the

2.5 m high- gain antenna which will communicate with the Deep

Space Network at a distance of around 33 AUs with x- band uplink
and downlink and the centralized Data Management System which

utilizes the Intel 80386 computer, and the Ada language for

software applications.
A. Bsrstta and F. Longoni. "Automatic controls on board planetary

probes", Dept. of space Instrumentation and Systems,Laben,Via E.

Bassini, 15 Milano, Italy.

Banket, V.L.. "Effective Methods of Modulation and Encoding for

SAtellite Systems with FDMA," Scripta Technica, Inc.,1989.

Bartok, Carol D.. "Catching the whispers from Uranus," Aerospace

America, May 1986, pp44-48.

"Communications," Aerospace America, Dec. 1986, pp 44-45.

Draper, Ronald F.. "Affordable solar system exploration,"

Aerospace America, Oct. 198S, pp 28-31.

Feldmann,Robert J.. "Feasibility analysis of an air-to-satellite

laser communication link," US Government work.

GAO, Information Management and Technology Division report, AAE241

notes.

"NASA's software engineering with Ada," Aerospace America,

Feb. 1989, pp 8-10.

"Software," Aerospace America, Dec. 1987, p 43.

Theis, Douglas. " Spacecraft Data Management Hardware

State-of-the-art," AIAA, Inc., 1984.

Zyuko, A.G.. "Methods for Increasing the Effectiveness of

Satellite Communication Systems," Scripta Technica, Inc. ,1989.
5. STRUCTURE

5.1 Requirements to be met by the structure:

The structure has the objective to support all other

subsystems and carry them out to Pluto safely. It has to protect

them from destruction or damage and also from influences which

might affect the performance of those subsystems. In this context

the following requirements were derived from the RFP.

use no materials available after 1990

lifetime long enough, with a safety margin

weight and cost optimization

stress reliability

stress simplicity

stress low cost

nothing should preclude other missions

interface to the launch vehicle

if necessary, on orbit assembly should be minimized

5.2 Shape and Configuration:

5.2.1 Grouping:

The structure of PULSE has to support all subsystems and meet

all the different requirements from those systems. In order to

comply with conflicting requirements, groups of subsystems with

similar requirements have to be placed together. This subsystem

grouping yielded 4 major areas with different necessary attributes:

The main body :

Requirements: provide thermal environment

support mass

radiation shielding

micrometeoroid protection

withstand launch forces

Subsystems: Communication electronics

Control electronics

Data storage

Gyroscopes

Power conditioning equipment

Fuel pumps and lines
To meet these requirements the subsystems have to be

encased in a shell which will protect the inside from

micrometeoroids, radiation, will not yield due to the launch

forces and provide a sufficient insulation against heat

loss. Conflicting requirements are here low cost and low

weight against high protection and strength. Desirable is

also good damping of vibrations during take off to protect

the electronics from mechanical damage.

The science boom :

Requirements: negligible magnetic and electric interference

support mass

provide thermal environment

micrometeoroid protection

Subsystems: magnetic field instruments

particle detectors

The predominant point in this group is, that the s_srme

instruments have to be able to measure an as much as

possible undisturbed environment. To keep disturbance by the

electronics on board the probe as low as possible, those

instruments have to be away from the spacecraft. Even though

micrometeoroid protection is necessary, shielding is not

feasible since that would shield off the fields to be

measured also. The same applies for the heating. On one hand

the electronics needs to be kept at an operating

temperature, but on the other hand, heaters would create a

disturbance. For these reasons, the instruments have to

provide these measures themselves.

The science platform:

Requirements:

Subsystems:

Pointability and good field of view

support mass

micrometeoroid protection

provide thermal environment

pointability

Science instruments

spectrometer)

(cameras,infrared

Other science instruments require less shielding than

the field and particle instruments. For this reason they can

be mounted on the main body and micrometeoroid protection

and heating can be supplied by the structure. In addition to

the control electronic housed in the main body these

instruments needs to be pointable and they have to have a

good field of vision. This is accomplished by separating

them from the main body and mounting them on a movable

platform on top of the main body. To ensure the

micrometeoroid protection, a steel canopy is placed over the

platform. Steel has been chosen to maximize the protection

since the science instruments are the essential parts of this

mission. During the cruise phase it will be closed and only
when PULSE approaches Pluto it tilts open. The platform will

be turnable by 360 degrees and tiltable by +- 15 degrees.

These values ensure that a large area can be scanned by the

mounted instruments.

The power boom:

Requirements: micrometeoroid protection

allow heat radiation

support mass

Subsystem: RTG

RTG's radiate a large amount of unwanted radiation

which would have a negative influence on the performance of

electronic equipment, this radiation has to be kept away

from those instruments. It would require heavy shielding to

protect the computers which would interfere with the

requirement of low weight. It also would affect the

necessary heat radiation of the RTG's. Thus the RTG's have

to be moved away from the main body. This yields now two

booms which can be spaced by 180 degrees to enhance symmetry

and maximize the distance between the sensitive science

instrumentation and the high radiation of the RTG's. The

spacecraft body also functions as a shield. The science

platform will not be operational during the cruise phase.

During the flyby, the open steel canopy will be tilted in

the direction to the RTG's to provide shielding.

Other subsystems:

The remaining subsystems are the antenna, the

propulsion tanks and the startracker and sun sensor. The

predominant requirement for the antenna is, that it has to be

pointed to Earth at all times. Additionally the antenna is

required to function as an adapter interface with the launch

vehicle. This yields, that the antenna is firmly mounted on

the main body to provide the necessary support. Thus the

whole body of the spacecraft will be pointed at earth.

The propellant tanks will be bought from stock and

placed next to the main body on both sides of the boom

structure. This will limit the volume needed for the main

body and thus decrease the weight. There will be four

propellant tanks and the their steel body will provide a

sufficient protection against micrometeoroids.

The startracker and the sun sensor need a good field of

vision to be able to scan a large area. This is accomplished

by placing them on the rim of the parabolic antenna. Both

have similar pointing requirements, and since the difference

in angles to the sun and the earth is maximal 12 degrees in

the periphery of our sun system the instruments have to

provide only a small correction to their pointing. Here they

also have a large angle available where no obstacles block

their field of vision.
¢,,,.<,.e'<< t,_ _/,,<,,,,D,iy
©

I=,

r-T.1

%

L_ ....
5.2.2 Shape determination

The main driver when determining the shape of the main body,

is the prevention of heat loss to space. An important variable

there is the surface. The smaller the surface, the smaller the heat

loss. Therefore I considered shapes which allow me to have a large

volume but also have a small surface area. Obviously the sphere has

the highest volume to surface ratio (V/S ratio) but production and

interface problems make the sphere less desirable to be used on

PULSE. I then considered the cylinder. It has a smaller V/S ratio,

but provides two flat interface surfaces. Looking at the amount of

equipment to be mounted inside the hull it is apparent, that this

is not enough. Adapters need to be installed to fit the instruments

to the curved surfaces. This would increase the weight of the

structure and complicate the manufacturing. From these

considerations I propose a regular octagon as the shape of the main

body. It has still a high V/S ratio but has flat sides so the

instruments can easily be mounted.

From the volume required I derived the design sizes. This yielded

a diameter of 0.5 m and a height of 0.8 m.

5.2.3. Configuration:

Due to the requirements of having both RTG's and highly

sensitive particle and field instruments on the same craft, it is

necessary to separate them as far as possible. For this reason

booms need to be employed. I propose two booms, one carrying the

two RTG's and the other all the particle and field sensors. This

enables a 180 degrees separation which gives the maximum separation

distance. This way the main body also acts as a shield in between.

Since even the on board electronics interfere with those sensors,

the science boom needs to be considerably longer than the power

boom. Only 3 m are necessary for the power boom this allows the

downward folded boom to fit in the launch vehicle in it full

length. The science boom, which requires a length of 10.6 m needs

to be partially retractable. This retraction technique can be

directly inheritated from the Galileo spacecraft.

The antenna will be firmly mounted on top of the main body so

that its center section can support the adapter to the launch

vehicle. I also considered making the antenna pointable. This would

decrease the attitude correction maneuvers and thus reduce the

necessary amount of propellant. Added weight and complexity due to

the pointing mechanism and compatibility problems with the launch

vehicle discard this option. A pointing mechanism would not be able

to provide a stiff support when placing the adapter on the antenna.

A complex design is necessary to comply with both, the pointability

and the stiffness during launch. Placing the adapter on the other

side of the craft requires a very large adapter because it has to

give room to the booms and using the booms is not feasible because

they, as the pointing mechanism are not stiff enough to firmly

support the probe during launch.

Since the remote sensing instruments need to be pointed at

the object of interest and the antenna needs to be pointed at

earth, a pointing mechanism is necessary for the science platform

which will house the remote sensing equipment. These can than be
e,cj

e,cj

_D

[- | !
pointed independently from the main body. During the cruise phase

these instruments are not used and to protect them a steel canopy

is placed over them. This canopy will tilt open when the

instruments are operational.

5.3. Material selection:

To perform the material selection I gathered as much

information from different sources as possible and incorporated

them into the following table.

PROPERTIES:

Property A1 Be

Density

Yield str.

machinability

weldability

handling

cost

corrosion

resistance

Mg Ti Kevlar Steel Unit

2.8 1.85 1.74 4.5 1.9 7.87 g/cmA3

500 415 103 830 1600 1800 MPa

ex. poor ex. good poor good

good poor ex. good none ok

ex. poor ok ex. poor ex.

low high low mod. high low

ex. ok poor ex. ok ex.

I then awarded points for their properties on the scale of 0

through i00 according to the desirability of the properties.

POINTS:

Property A1 Be Mq Ti Kevlar Steel weiqht

Density 72 81.5 82.6 55 81 21.3 0.55

Yield str. 25 20.75 5.15 41.5 80 90 0.i

machinability i00 40 i00 80 40 80 0.i

weldability 80 40 i00 80 0 60 0.075

handling I00 40 60 i00 40 i00 0.05

cost i00 0 I00 60 0 i00 0.i

corrosion I00 60 40 i00 60 i00 0.025

resistance

Sum : 577 282.2 487.7 516.5 301 551.3 1
The final evaluation is based on the points received and a

weighing factor which allows to stress more important properties

over less important ones.

EVALUATION:

Property A1 Be Mg Ti Kevlar Steel weiqht

Density 39.6 44.82 45.43 30.25 44.55 11.715 0.55

Yield str. 2.5 2.075 0.515 4.15 8 9 0.i

machinability i0 4 10 8 4 8 0.i

weldability 6 3 7.5 6 0 4.5 0.075

handling 5 2 3 5 2 5 0.05

cost i0 0 10 6 0 i0 0.I

corrosion 2.5 1.5 i 2.5 1.5 2.5 0.025

resistance

Sum -

Selectiom made:

75.6 57.4 77.44 61.9

Magnesium

60.05 50.715

Legend: Points synonym

100

8O

60

40

20

0

ex. or low

good

ok or mod.

poor

bad

none or high

Formulas used: For density : Points = i00 - density/10

=> density = 0 -> i00 Points

=> density =i0 -> 0 Points

For yield strength : Points = Ys / 20

=> Ys = 2000 -> i00 Points

=> Ys = 0 -> 0 Points
5.4. Calculation of required wall thickness for

micrometeoroid protection.

Material proposed:

Magnesium

Constants:

meteoroid mass,M :

meteoroid velocity,V :

meteoroid density,roh :

mat. constant for A1 :

mat. constant for Mg,K :

Density of Mg,RMG :

Yield strength,YS :

0.i g

25 km/s

0.5 g/cm" 3

0.06 (from reference)

0.08 (estimated)

1.74 g/cm*3

22000 ibf/in^2

Derived Values:

Variable:

meteoroid diameter,D : 0.725566 cm

(spherical meteoroid shape assumed)

first sheet thickness,Tl : 0.072556 cm

(TI/D=0.1 requ. by Formula)

spacing,S : 2 cm

Formula : (for double sheet penetration)

t = K*roh^0.15*M'.35*V/S^0.*(70000/YS)

t = 1.015542 cm

Summary :

First sheet thickness,Tl : 0.072556 cm

Second sheet thickness,t : 1.015542 cm

Spacing,S : 2 cm

Protects from 0.i g micrometeoroid at average speed.

Design sizes :

First sheet thickness,Tl :

Second sheet thickness,t :

Spacing,S :

0.2 cm

0.9 cm

2 cm
5.5. Mass estimation from desiqn and sheet thickness:

Constants:

First sheet thickness,tl :

Second sheet thickness,t :

Lid thickness,tl :

Density of Mg,roh :

Area of spar,Asp :

Variables:

0.2 cm

0.9 cm

1 cm

1.74 g/cmA 3

4.1 cm^2

Height,h :

Diameter,d :

80 cm

50 cm

Formulas:

Panel length,s : s = d/2 * (2-2"0.5)'0.5

S = 19.13417 cm

Panel area,Ap : Ap = 8 * s * (tl+t)

Ap = 168.3807 cm^2

Spar area,As : As = 8 * Asp

As = 32.8 cm^2

tot. cross sect.

area,Ac :

Ac = As + Ap

Ac = 201.1807 cmA2

Lid area,Al : A1 = DA2 * 2".5 / 2

A1 = 1767.766 cm^2

Lid volume,Vl : Vl = 2 * A1 * tl

Vl = 3535.533 cm^3

Trunk volume,Vt : Vt = Ac * h

Vt = 16094.45 cm^3

total Volume,V : V = Vt + Vl

V = 19629.99 cm^3

Total weight of the main body structure:

M : 34.16 kg
5.6. Production techniques required:

The magnesium side panels can be bought from stock, cut and

welded to the spars. The magnesium spars need to be extruded. The

main body lids and the base of the sclence platform have to be

casted. The steel canopy has to be produced by dee_ drawing and

then weld the second sheet onto it to enhance the mlcrometeoroid

protection. The boom struts can be bought from stock and then

assembled.

All these techiques are well known and readily avalible today. Any

new developements can be incorporated at a later point to improve

the performance of the craft.
5.7 Re_erences:

it

2.

3.

4.

5.

6.

7.

8.

Scientific Satellites, William R. Corliss, NASA 1967

Spreading Spectrum of Reinforced Fibres, Alan S. Brown,

Aerospace America Januar 1989, pp 14-18

Selection and Use of Engeneering Materials, F.A.A Crane,

J.A. Charles, Butterworth 1984

Materials in Space Technology, Thompson, Gatland, ILIFFE

1963

Metallurgical Assesment of Spacecraft Parts and

Materials, Barrie D. Dunn, Hallsted Press 1989

Engeneering Materials: Properties and Selection, Kenneth

Budinski, Reston Pub. Inc. 1979

Space Materials Handbook, J.B. Rittenhouse, J.B.

Singletary NASA SP-3051

NUSAT I - The first GAS can ejected satellite, AAS paper

86-293
Propulsion

Numerous factors must be considered in selecting propellants

and propulsion systems for space missions. One of the more general

characteristics is performance, in terms of both specific impulse

and hardware mass. Final selection must depend on tradeoffs

between several of the major competing selection criteria: for

example performance, reliability and cost.

The first decision to make was what launch vehicle the Pulse

probe would be launched on. After evaluation of all of the United

States vehicles and some International launch vehicles, it was

found that the four best choices for this mission were the U.s.

Space Shuttle, the Ariane IV, the Titan IV Centaur G Prime, and the

Titan IV IUS. This primary trade study was based on the mass that

each vehicle could be place into a geostationary transfer orbit.

The United States Space shuttle was ruled out because of the higher

cost for a non-expendable launch vehicle.

After this preliminary study a more in depth study was

performed on the Ariane IV and the Titan IV configurations. Using

the equations from Conway (Ref. 4), a comparison was made between

the three launch vehicles on the basis of payload ratio, propellant

mass and total mass, given a delta-v and a payload mass (Figures

6.1, 6.2, 6.3, 6.4, 6.5). The conclusion reached was that the

Ariane IV launch vehicle was the best selection in all comparisons.

The Launch Specifications for the Ariane IV are given in the

appendix.

The fuel used for each stage of the Ariane vehicle will be the
specified fuel in the launch specifications in the appendix. In

these specifications one will find that the diameter of the upper

stage is 2.59 meters in diameter which is sufficient for the

largest diameter of our spacecraft which allow the antenna to fit

in uncollapsed.
Fig.6.1

_ubsystem Masses

System and components

Science

Telecommunications

C:ontroI

Receiver

Amplifier

Data handling

Data storage

Spacecraft control

Computer and sequencer

Sun sensors

C:_lopus tracker

Gyros

ISc_n control ,qnclplanet -,,en,_or

Electrical power

RTG's

Conditioning and control

("abling

Structure and mechanical

Bu;

Parabolicantenna

Temper,q.turecontrol

"rrajecto_ correction 13roputsion

Total ;pa.cecra.ii weight

Launch vehicle adapter

Toto.I injected weight

Number ol

component;

(l+redundanc't')

1

Weight (kg)

i 94.9 75.46

22.73

14.55

3.54

1636

18.18

38.17

10.91

5.45

5.45

5.45

10.91

121.41

44.4

45.45

31.82

290.4fi

150

9.1

11,36

120

620.4

50

670.4
Fig.6.2 Launch Sr_:_ficado_

Variables Ariane IV Tit=_n IV C_ntaur El.prime Titan IV IUS

thrust1 IN] 204318.20 72715000.00 72715000.00

thrust2 [N] 40227.30 23636.40 23626.40

th rust3 [N] 3181.80 7500.00 13840,90

thrust (total) [N] 247727.30 72746136.40 72752467.30

cl [kin/s] 3038,00 2989.00 2989.00

c2 [kin/s] 3136.00 3136.00 3136.00

c3 [kin/s] 3528,00 3528.00 2842.00

c (total) [kin/s] 9702.00 9653.00 8967.00

R1 2.56 2.46 3.45

R2 2.13 2.19 2.88

R3 2.90 2.95 1,99

R (total) 7.58 7.61 8.32

Ms1 [kg] 786.02 779,84 1210.00

Ms2 [kg] 334.52 366.74 372.91

Ms3 [kg] 125.11 129.82 57.84

Ms (total) [kg] 1245.64 1276.39 1640.75

Mpl [kg] 10510.00 10420.00 16180.00

Mp2 [kg] 3161.00 3465.00 3524.00

Mp3 [kg] 1672.00 1735.00 773.20

Mp (total) [kg] 15343.00 15620.00 20477.20

Mo [kg] 17260.00 17570.00 22790.00

lambda 1 0.53 0.57 0.31

lambda 2 0.71 0,66 0.39

lambda 3 0.37 0.36 0.81

lain bd a (total) 1.61 1.59 1.50

1,7

1.0 t

1.4

1.2

0.._ _

0.7

0.¢3 _-

°°I 0.4

0..._

0.2

)

C C_._.

Fig.6.3 Payload Mass Rat[o

Artar_ IV"

,,_,,,, ,

T_tar_ IV Cet_taur 0 Prttt_

. ^ .

<:_:" ;,'.,_ ;_ <4

b::,.t¢.,_;%7-..G-. ;

Ttta_ IV /US

._'5<

._"_." ..y.:_-_._

Lamt_ai _ Lamt_a_ _ Lamb_a:_ _ Lambda<totalJ
Fig.6.4 Propellant Mass

20

1,6

14

12

IC

8

4

2

o

L

F

t

I

I

.4r!a.,,'_ b/

T', _' I_S

Fig.6.5

24

2O

14

B -

4 -

2 -

0 ®

Total Mass of Entire System

Tlten _ I&G
Power System

The operational capabilities of a space vehicle is dependent

upon an adequate supply of power. This power is necessary for

communications, guidance, control, and operation of sensors or

scientific instrumentation.

When trying to select a power source for the PULSE probe there

were 12 factors which I took into consideration: l)Duration

2)Mission 3)Availability 4)Reliability 5)Weight 6)Compatibi!ity

7)Environment 8)Power level 9)Area 10)Cost ll)Volume 12)Hazard.

Since the mission duration of our probe is about 16 years the

selection of power source was limited to nuclear power, either from

decay of an isotope or a nuclear reactor. Batteries were also

considered for storing the electrical energy provided by the power

source. The approach taken consisted of listing the 12 factors and

rating the sources from 1 to 10(highest) on the quality of

performance related to each of the 12 factors as shown in figure

6.5.

The results from this trade study eliminated the nuclear

reactor as a power source but showed that batteries should be

further considered as energy storage devices for the RTGs. B_.it

when looking at the predicted power to weight ratio of both the

RTG(12 W/kg) and the Ni-Cd battery (I0 W-Hr/kg) in the year 2000

the choice was that the RTGs were the only power source that was

going to be used on the PULSE probe (Ref. I0, pp.l-45).

The next step in developing the power system was finding out

how much power the power system would have to put out at peak
operating loads. Figure 6.6 shows a list of the subsystems and

the power that each subsystem requires at peak level. Figure 6.7

shows the percentage of power each subsystem requires of the total

power. A total power system requirement of 372.94 W is needed upon

arrival at Pluto.

The isotope selected for this mission is Pu 238, with a half

life of 87 years. This isotope has been proven by earlier space

missions and often exceeded its original design life requirements.

Some studies have used a design lifetime of I0 years for the RTG

and found that the RTG has a 20% reduction in power at the end of

the projected I0 year life (Ref. I0, pp.l-48).

The PULSE probe's RTGs will have to supply power for at least

16 years. This results in a 70% reduction in 16 years which shows

that at launch the PULSE probe will have 529.7 W of power that

would diminish to the amount needed at Pluto (See appendix for

these calculations). No safety margin is needed with these figures

because the Pu 238 RTG "has operated considerably longer than their

original design life requirements" (Ref. I0, pp.l-44). From the

total power needed at launch a calculation was made to determine

the mass of RTG needed. The mass of RTG needed is 44.40 kg, which

would require 23 slices of fuel cells in the Modular Isotopic

Thermoelectric Generator (Ref. 12, pp.340) (See appendix for

calculations). The RTG fuel capsule is designed to withstand

intact reentry should there be a mission failure or abort.

The electrical power from the RTG will go to the Power

Conditioning Unit which will regulate the voltage and convert the

DC power into whatever form it needs to be in for the applied
loads. This will depend upon the voltages needed by the

instruments and if they are powered by AC or DC voltage (Figure

6.8).
Fig.6.6

Power Supply Determination

IDuration

iMission

Availability

ReIio.biIity

Weight

iC:omp,_tibiIi_

iEnvironment

lPower level

Area

Co_t

Volume

_H,qzard

 otal

ReactoY RTG

6

2

6

6

4

6

8

10

4

4

2

8

,10

_8

10

8

10

8

8

8

6

8

6 8

Battew

4

6

8

11]

8

8

8

6

6

10

8

10
Fig.6.7 Power Systems

System function

tScience

Telecommunications

C'ontrol

Receiver

Amplifier

Data samplin, encoding,

_nd decoding

Data storage

Spacecraft control

Sequencing 8nd command

Sun ,;el'l,_OrS

Canopus tracker

Gyros"

Electronics

Hes.ter,'_

Total sx._temrequirement._

C'onver;ion Io;s (20%)

Fotal power requirement

Power required

_t pe_k level;

(w tts)

78.78

110

5

10

70

20

5

78

10

3

10

15

4O

44

310.78

62.16

372.94

Fig.6.8 Power Subsyztems

25.10%

14. 16% _ _,,_

li!i iii!ii!iiiiii'iiii! iii!i!!!i!:ii

ii!iii!_ii?i:._ti [:!!ii:i !:: !i :_.Dl_!k:.:._!):-i?ii!:,:.::?::iilY !?i:i:::i!i!!i::i:i:i:::: _

iiii2i}i;i i}i:,i)i}i!i i}i:i!iil :il)!;i?i: !iii?il ;i;i;iii?_i_ i

',;:s:;:i::.:i:iiiiil}:!:i:i:):i,i,i:4:;.i:!iiiiii!:.:i[Y

:::::::::::::::::::::::::::::::

..... -_ 35.39%

i
Fig.6.8

Power System

Eflefqy Source

Solar Photons

Heat From DecaM_ng Radioisotope

Heat From Nuclear Reactor

C:hemicpl Enerqy

P (Wa_s)

Enerqy Converter

_olar Cell

Thermoelectric Elements

Thermionic Converter

Fuel Ceil

} Batter_'

V

(I-e)P (watts)

Waste-Heat Radiator

Solar Cell Surtace,'_

Fins On Radioisotopic Generator

Satellite Skin. Where Conduction AndS_r

Radiation Heat Transfer Paths are

Provided

eP

(wstts)

i Power-Conditioning Unit

i Itoltaffe Regulators DC-A C Cont_rters

> DO-AOlnverter._

L Batlery Fuses Etc. (-harge t::ontrol

Waste Heat Radiated

To £pace Environment
Propulsion

Appendix I

Isp = specific impulse

V = delta V needed

= structural coefficient

ML = mass of the payload (spacecraft)

Thrust = thrust given by each of the stages

c = exhaust velocities

f(a) = function used for Newton's Approximation

fprime(a) = derivative of f(a)

a = Lagrange multiplier

R = mass ratio

MSP = mass of structure and propellant of that stage

M = mass of that stage plus payload weight

Ms = mass of the structure of that stage

Mp = mass of the propellant on that stage

Mo = total mass of the launch vehicle and spacecraft

:,',= payload ratio

Massflow = massflow of that stage

Burntime = burntime of that stage

Base units:

sec _ IT

Normal units:

m

N _ kg ..................

2

sec

Constants:

kg HIM

km { 1000m

This shows only one launch vehicle.

A chart with all the values is in the text

Isp := 310 sec Isp := 320 sec

1 2

m - Ii

kg

ib - ................

2.2

ibf - 4.4 N

This process was done 3 times

Isp := 360 sec

3

m km

g := 9.8 ....................... V := 8.974 .................

2 sec

sec

ML := 670.40 kg

Assuming structural coefficients to be the same for Titan and Ariane

(Actual Ariane values)

£ := .0696 g "= °0957 g := .I008

1 2 3

thrust -= 899000 ibf

1

thrust := 177000 ibf

2

thrust := 14000 ibf

3
Equations:

± := 1 ..3

c := Isp g

i i

Iteration using Newton's approximation

f(a) := V -

,,q............... i

%!,,

1

lac - 1

C ln! i sac 1

i i a c a I

L i i j

C

i

3 1 -1'

......3..,..0..3...8.......1.Q._...:._ lencJ_ h i ...t ime...... 1 _

..... 3....I,3.,6,,i, ,I._L..:,...i e,ngth ..... t im_ ........ '.

1 3 1 -I _

i,.,.3.,.,. 5,.?..8.,.L.....I.0 ............. l.._.ng.th,,.... .....t ime ......

fprime(a) := "i_'._''

i

j := 0 ..20

X := until

j+l

C

i

C "]

km

ac - 1 .................. I

i sec]

X := .43

0

km 1

.oooi ................. ,ix -

sec ..,

-!

a .99999999999 I

............ ::J!...................

......

...... l

-1

n " = size(x)

km

a c - 1 .................

1 sec

R1 "= .......................................................................

a c E

1 1

:= X

n

km

ac - 1 ................

2 sec

R2 := ......................................................................

a c

2 2

5=0.4

km

a c - 1 ............

3 sec

R3 := ..............................................................

a c

3 3

R1 = 2.557 R2 = 2.128 R3 = 2.898
ML - R3 ML

MSP3 := ....................................................

R3 £ - 1

3

MSP3 = 1.798 I0

MSP3 + ML - R2MSP3 - R2ML

MSP2 := ........................................................................................................................................................

R2s - I

2

3

MSP2 = 3.495 I0

MSP2 + MSP3 + ML - R1MSP2 - R1MSP3 - R1ML

MSPI := .............................................................................................................

RI£ - 1

1

4

MSPI = 1.129 10

M := MSP3 + ML M := MSP3 + MSP2 + ML M

03 02 01

3 3

M = 2.468 10 mass M = 5.963 I0 -mass M

03 02 01

:= MSP3 + MSP2 + MSPI + M

4

= 1.726 10 mass

Ms := _ MSPI Ms := _ MSP2

1 1 2 2

Ms := _ MSP3

3 1

Ms = 786.017_mass

1

Ms = 334.515mass

2

Ms = 125,111 mass

3

Mp := MSPI - Ms Mp := MSP2 - Ms

1 1 2 2

4 3

Mp = 1.051 I0 mass Mp = 3.161 I0 mass

1 2

Mp := MSP3 - Ms

3 3

3

Mp = 1. 672 I0 mass

3

M

0

:= MSPI + MSP2 + MSP3 + ML

M

0

4

= I .726 I0 mass
M

O2

.= ................ ,, := ., :_

! M - M 2 M - M 3

0 02 02 03

M

03 ML

M - ML

O3

'. = 0.528 >, = 0.706

1 2

>' = 0 . 373

3

thrust

i I

Massf!ow :=

i c I000

i

Burntime

±

Mp

i

Massflow

i

Massflow

±

1 1 -i i

_L._.Q 21_ma,.%s.._t.im_e___/

1 -! i

1 -1:

Burntime

i

3 1

8.07 I.Q_,___i_L__,

4 I _

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_9 10 tSm.L

:!

= 1

|

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i

:lk

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_am'_ 3

4J

t ,

i _:_ i_' i

_2

1 = MM-7B I=_*_

" = Y'_,,:J _ i,:*=:i

_ * 1:'7.3 t:x,a

1 x MM- Z'_ IZ_._Q

24.V, _61m_

EFINO/S£P

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_3_-

' __-"4_ar " rcu I (31

ORIGI?,!AL PAGE IS

OF POOR QUALITY
Power

20% decrease in power over I0 years (Ref. I0, pp.l-48)

N(t) = percentage of power after t years

No = percentage of power at launch

k = decay constant

t = time

N(t) = No e-_t

.80 = I e -k(I°l

k = -in(.80)/10

k = 0.022314

N(t) = I e "I°'°2n_'_''°°5_

N(t) = 0.69967

This is a 30% decrease over 16 years

Total power needed/70% = Power at launch/100%

372.94/70% = Power at launch/100%

Power at Launch = 529.69 W

Assuming (12w/kg) power to weight ratio predicted for the year 2000

(Ref. I0, pp.l-45)

529.69 W/12W/kg= 44.40 kg of RTG at launch

MITG Generator give 23.5W/slice (Ref. 12, pp.340)

529.69 W / 23.5W/slice = 22.54 slices approximately 23 slices
I. Advanced Propulsion concepts, Proceedings of the Sixth AFOSR

Symposium, Air Force Office of Scientific Research, Vol. I, 1971,

(629.42S468P 1971)

2. AIAA/NASA Space Systems Technology Conference, A Collection of

Technical Papers, AIAA/NASA Space Systems Technology Conference,

June 5-7, 1984, (629.44A:IIC)

3. Bennett, Gary L., and Lombardo, James J., and Rock, Bernard J.,

Development and Use of Nuclear Power Sources for Space

Applications, The Journal of the Astronautical Sciences, VoI.XXIX,

No.4, pp 321-342, Oct-Dec., 1981.

4. Conway, Bruce A. and Prussing, John E., AAE 306, Orbital

Mechanics, Fall 1989, University of Illinois, (Class Notes).

5. Corliss, William R., Spacecraft Power, National Aeronautics and

Space Administration, Washington, D.C. 20546, (629.474C81S).

6. Harvego, E.A. and Sulmeisters, T.K., A Comparison of Propulsion

Systems for Potential Space Mission Applications, ASME Winter

Annual Meeting (December 13-18, 1987), (621AS54).

7. Hove, L.E, Design of an Unmanned Automated Spacecraft for the

Grand Tour, Spacecraft Systems, 1970, (629.4In82p).
8. Jones, Christopher P. and Risa, Thomas H., The Voyager

Spacecraft System Design, AIAA Paper 81-0911, AIAA 1981 Annual

Meeting and Technical Display, (629.4am35a).

9. Koepke, Andy., AAE 241 Spacecraft Propulsion System, Spring

1990, University of Illinois, (class Notes).

I0. Mondt, J.F. and Rock, B.J., Thermal-to-Electric Conversion,

Jet Propulsion Laboratory.

II. Palaszewski, B. and Engelbrecht, C., Lightweight Spacecraft

Propulsion system Selection, AIAA Paper 1987 AIAA/SAE/ASME/ASEE

23rd Joint Propulsion Conference, AIAA Paper 87-2022, (629.4Am35a).

12. Schock, A., Modular Isotopic Thermoelectric Generator,

Fairchild space and Electronics Company, Germantown, Maryland

20767.

13. Space Launch Vehicles, John F. Kennedy Space Center, Rev. Nov.

1983, (DOC.NASI.20:KSCI35-81).

14. Smith, Bruce A., U.S. Planetary Launch Surge, Aviation Week

& Space Technology, Oct. 9, 1989, (629.105AU Vol 131 No 15).

15. Sutton, George P., Rocket Propulsion Elements, Fifth Edition,

1986, (ISBN 0-471-80027-9).
16. Transactions of the Symposia on Space Nuclear Power Systems,

Vol. 2, Second symposium, Jan 14-16, 1985, (629.4753SY68S).

17. United States. Congress. Senate. Committee on Commerce,

Science, and Transportation, "Commercial Space Launches," 1984,

(MFICHEDOC.YI.I+5:98-656).

18. Unmanned Exploration of the Solar System, Volume 19, Advances

in the Astronautical Sciences, Proceedings of the American

Astronautical Society, Symposium on Unmanned Exploration of the

Solar System, Feb 8-10, 1965, (TL787.A6A2).

19. U.S. and International Launch Vehicles, Aviation Week & Space

Technology, March 20, 1989, (629.105AU).
The spacecraft PULSE uses much off-the-shelf hardware from Voyager

and other planned probes. New technology is only applied if it

would include a more reliable and less costly trade-offs, as in

the case of onboard computers. PULSE willyield quality science at

low cost by using incorporation of off-the-shelf products,

choosing radiation-hardened version of widely available

microprocessor and integrated-circuit chips supported by efficient

software. In general, proven techniques were used throughout the

entire design.
• Jj

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• J,

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¢_

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0

I

i_

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m--

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a wl|ep
Costing for PULSE

Category

Structure

Thermal Control

Propulsion

Attitude & Articulation

Telecommunications

Antennas

Command & Data Handling

RTG Power

Line-Scan Imaging

Particle & Field Instruments

Remote Sensing Instruments

System Support & Ground Equipment

Launch + 30 Days Ops & Ground S/W

Image Data Development

Science Data Development

Program Management

Flight Operations

Data Analysis

"r °

Cost (FY 88 Dollars)

59,988,162.98

ii, 037,938.33

412,927,670.50

62,614,609.37

64,098,191.33

13,043,018.66

24,500,108.53

37,386,446.55

170,454,335.10

71,222,537.72

29,154,302.64

280,062,535.20

57,185,698.78

6,957,007.47

11,487,733.40

17,365,267 .83

258,722,216.60

115,984,760.70

TOTAL I, 704,192,542.00
Costing for PULSE

Category

Structure

Thermal Control

Propulsion

Attitude & Articulation

Telecommunications

Antennas

Command & Data Handling

RTG Power

Line-Scan Imaging

Particle & Field Instruments

Remote Sensing Instruments

System Support & Ground Equipment

Launch + 30 Days Ops & Ground S/W

Image Data Development

Science Data Development

Program Management

Flight Operations

Data Analysis

TOTAL

"r •

Cost (FY 88 Dollars)

59,988,162.98

ii, 037,938.33

412,927,670.50

62,614,609.37

64,098,191.33

13,043,018.66

24,500,108.53

37,386,446.55

170,454,335.10

71,222,537.72

29,154,302.64

280,062,535.20

57,185,698.78

6,957,007.47

11,487,733.40

17,365,267.83

258,722,216.60

115,984,760.70

1,704,192,542.00
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Instrument

ISS

MAG

NIMS

PPR

WS

PLS

EDP

PWS

CRS

Table 3.2

Weights and Power for PEP Instrumentation

Power (W) Mass (kq)

20 28*

2.2 5.6

13 18

4.5 4.8

5.33 4

I0" 12

i0" 9

8.4* 6

5.35 7.5

* Values are estimates

f

bb

I
Table 3.2

Weights and Power for PEP Instrumentation

Instrument Power (W) Mass (kg)

ISS 20 28*

MAG 2.2 5.6

NIMS 13 18

PPR 4.5 4.8

WS 5.33 4

PLS i0" 12

EDP i0" 9

PWS 8.4* 6

CRS 5.35 7.5

* Values are estimates
Table 3.2

Weights and Power for PEPInstrumentation

Instrument Power (W) Mass (kq) _

ISS 20 28*

MAG 2.2 5.6

13 18

PPR 4.5 4.8

VVS 5.33 4

PLS i0" 12

EDP i0" 9

8.4* 6

CRS 5.35 7.5

* Values are estimates

D,O. -6_eu,_.'-7