asS SATELLITE PROPULSION SYSTEM ANALYSIS · satellite with a 5-year lief per for minc. north-south...

AFRPL-TR-7M08

S SATELLITE PROPULSION SYSTEM ANALYSIS as

(^ R.D.KLOPOTEK.CAPTJSAF W.LS.UUKHUF.CAPTJSAF

TECHNICAL REPORT AFRPL-TR-7M08

SEPTEMBER 1971

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AIR FORCE ROCKET PROPULSION LABORATORY

DIRECTOR OF LABORATORIES

AIR FORCE SYSTEMS COMMAND Q ^ Q^j: UNITED STATES AIR FORCE T ' in! 'fi

EDWARDS, CALIFORNIA i

i

,,,

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UNCLASSIFIED Security Classific alion

DOCUMENT CONTROL DATA R&D (Security ctet-sillcation ol title, body ot abstract Itncl indexing utmottition nmsl be entered when the overall report is cttmsilled)

l ORIGINATING A c T l v l T v f Curporale aulhor;

Air Force Rocket Propulsion Laboratory Edwards, California

2«. REPORT SECURITY CLASSIFICATION

Unclassified 2b. GROUP

3 REPORT TITLE

Satellite Propulsion System Analysis

* DESCRIPTIVE NOTES fTVpo ol report and Inclusive dates)

Final (October 1970 to June 1971 5 AUTHORISI (First name, middle Initial, last name)

Raymond D. Klopotek, Capt, USAF Waiden L. S. Laukhuf, Capt, USAF

6 REPORT DATE

September 1971

7«. TOTAL NO OP PAGES

152 & 13 7h. NO OF HE FS

Ba. CONTRACT OR GRANT NO

b. PROJECT NO 3058

c Task No. 305801

9a. ORIGINATOR'S REPORT NUUBERISI

AFRPL-TR-71-108

9b. OTHER REPORT NO(S» (Any other numbers that may be assigned this report)

10 DISTRI BUTION STATEMENT

This document has been approved for public release and sale; its distribution is unlimited.

II SUPPLEMENTARY NOTES 12. SPONSORING MILITARY ACTIVITY

Department of the Air Force/AFSC Air Force Racket Propulsion Laboratory Edwards, California 93523

13. ABSTRACT

Propulsion systems were studied for post-1975 geosynchronous satellites that have stringent altitude and station maintenance requirements with mission durations of up to 10 years. Systems using catalytic monopropellant, nuclearthermal monopropellant, chemical bipropellants and electric thrusters were studied and ranked according to several analysis areas which included propulsion system weight, system reliability and costs. Areas requiring further technology development are recommended on the basis of system rankings.

DD FORM ,1473 151 INCLASSIFIFD Sf ( unlv Claisifi« «lion

- _ __

■iqui.îiiîîjimiBt'^.in- '^^^

I NChA.SSIl'IK'D Set-unty Classification

KEY «ORJS

C iuosy iiclironouä SaIr Hi I erf .Spiue Propulsion Systems Advanced Propulsion Systems Three-Axis Stabilized Satellite

1'roi)i;lsion System Weighl

XC:LASSIFIFD Security Classification

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AFRPL-TR-71-108

SATELLITE PROPULSION SYSTEM AXAI YSIS

Raymond D. Klopotck, Cap!, USAF

Waiden L.S. Laukhuf, Capl, USAI'

September 1 ')7i

This document has been approved for publu rrb ,i.s1

its distribution is unlimited.

AIR FORCE ROCKFT PHOPILSIOX ! .A I ■(; I; .\1 ( M D1KECTON OF I./M'Olx'ATOKM S AIH FORCFJSVS1 FMS COMMAND

UNITED STA'J Is Ail: I-()PfI" EDWARDS. CAl.ir Oh' MA

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FOREWORD

This rrporl summarizes work performed during a USAK in-house program under Project 3058, during the period October 1970 through June 1971.

The program was conducted by the Liquid Rocket Division of the Air Force Rocket Propulsion Laboratory. Captains Raymond D. Klopotek and Waiden L. S, Laukhuf were the project engineers.

The authors gratefully acknowledge the assistance of the following individuals in support of their Satellite Propulsion System Analysis: Dr. L. Quinn, Mr. P. Van Splinter, Mr. P. Erickson, Mr. M. Rogers, Mr. E. Barth and Captain D. Huxtable of the Air Force Rocket Propulsion Laboratory; Dr. D. Fritz of the Air Force Aero Propulsion Laboratory.

This technical report has been reviewed and is approved.

JERKY N. MASON, Capt, USAF Chief, Subsystems Branch Liquid Rocket Division

11

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ABSTRACT

Propulsion systems were studied for post-1975 geosynchronous satellites that have stringent attitude and station maintenance requirements with mission durations of up to 10 years. Systems using catalytic mono- propeliant, nuclear-thermal monopropellant, chemical bipropellants and electric ihrusters were studied and ranked according to several analysis areas which included propulsion system weight, system reliability and costs. Areas requiring further technology development are recommended on the basis of system rankings.

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TABLE OF CONTENTS

Section Page

I INTRODUCTION I

II. APPROACH 3

III ANALYSIS 5

A. Post-1975 SYNCSAT Mission Model 5

1. Introduction 5

2. Satellite Geometry (,

3. Mission Requirements 9

B. SYNCSAT Propulsion Systems ] 5

1. Hydrazine Catalytic/Hydrazine Catalytic 17

2. Hydrazine Catalytic/Hydrazine Plenum \q

3. Hydrazine Catalytic/Hydrazine Electrolytic Ignition 21

4. Hydrazine Catalytic/Hydrazine Resistojet 23

5. Hydrazine Catalytic/Hydrazine Radioisotope. ... 26

6. Hydrazine Catalytic / DAR T 30

7. Hydrazine Catalytic/Cesium Bombardment Ion . . 32

8. Hydrazine Catalytic/Colloid 37

9. Hydrazine Catalytic/Mercury Pulsed Plasma ... 42

10. DO Radioisotopc/DO Radioisotope 4(,

11. DO Radioisotope/Colloid \y,

\1. Water Elect rolysis/Water Electrolysis 5Q

13. Water Elect rolysis/Colloid 53

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TABLE OF CONTENTS (CONT'D)

Section Page

m ANALYSIS (Cont)

B. SVNCSAT Propulsion Systems (Cont)

14. N O /N9H Bipropellant/Hydrazine Plenum . 56

15. C1F5/N H Bipropellant/Hydrazine Plenum . 56

16. DART/DART 58

C. Propulsion System Weight 60

D. System Costs 74

E. System Reliabilities 75

F. Plume Effects, Integration and Handling 85

IV CONCLUSIONS AND RECOMMENDATIONS 87

APPENDIX A - SATELLITE PROPULSION SYSTEM WEIGHT PROGRAM DESCRIPTION 91

APPENDIX B - SATELLITE PROPULSION SYSTEM WEIGHT PROGRAM USER MANUAL 101

REFERENCES 149

FORM 1473 151

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LIST OF TABLE'S

Table

I SATELLITE MODEL GEOMETRY

II GEOSYNCHRONOUS ORBIT PERTURBATIONS

III MISSION REQUIREMENTS

IV PROPULSION SYSTEM WEIGHT FOR BnOO-POl'ND SATELLITE WITH A 10-YEAR LIEF HERFOR Mli$; NORTH-SOUTH STATIONKEEPINC;

V PROPULSION SYSTEM WEIGHT B'OU 3()0()-POUND SATELLITE WITH A 10-YEAR LIFE NOT PERFOHMl. ' NORTH-SOUTH STATIONKEEPINC

VI PROPULSION SYSTEM WEIGHT FOR ÜOOO-POl'ND SATELLITE WITH A 5-YEAR LIEF PER FOR MINC. NORTH-SOUTH STATIONKEEPINC

VII PROPULSION SYSTEM WEIGHT FOR ZOOO-POPNU SATELLITE WITH A 5-YEAR LIEF iN'OT PERFOR Mi ;v , NORTH-SOUTH STATIONKEEPINC

VIII POWER PENALTY

IX THRUSTER COST

X TOTAL PROPULSION SVSIKM CuST

XI PROPULSION SYSTEM COM l$:i:.! !•: I 1,1 A 1. II.: ;'

XII VALVE-THRUSTER RELIABILITY

XIII SYSTEM RELIABILITY

XIV LARGE THRUSTER FEED SV ST PM HFL1AHLUTY RANKING

XV SMALL THRUSTER FEED SYSTEM U t.'Ll AIM LIT'i RANKING

XVI TOTAL LARGE THRUSTER SYS'l I: : 1 K K I,! A h I I ,i i i RANKING

VI i

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LIST OF TABLES (CONT'D)

Table Page

XV 11 TOTAL SMALL THRUSTER SYSTEM RELIABILITY RANKING 84

XV III PLUME EFFECTS 85

XIX COMPUTER INPUT VARIABLES 102

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17.

I 'I.

.'11.

LIST OF ILLUSTRATIONS

Spari.i; raft Inertial Model 7

C'dUilytic Hydrazine Thruster |H

System Schematic for NJI. Catalytic - N 11 Catalytic. . . , 20

System Schematic for N-Jl. Catalytic - N,,H, GG Plenum . . 22 2 4 ■' 2 4

System Schematic for N^H, Catalytic - N^H . Electrolytic . . 24

N, II Rcsistojet Thruster 27 2 4

System Schematic for N?H. - N?ll . Resistojet 2.S

Nil . Radioisotope Thruster 29

System Schematic for N,H. Catalytic - N_H Radioisotope . 31 2 4 2 4

DART Thruster 33

System Schematic for INUH, Catalytic - DART 3-1

Cesium Ion Engine Schematic 3ci

Cesium Feed Systems 3^

System Schematic for N?ll. Catalytic - Cesium Ion }K

Cnlloid Engine Schematic Vi

Clolloid Thruster Concept -Id

System Schematic for N-.II. Catalytic - Colloid -13 2 4

Mercury Thruster Circuit Diagram -14

Pulsed Vacuum Arc Thruster (PVAT) 15

System Schematic for N,II. Catalytic - 1 Ig Pulsed Plasma. . 17

.System Schematic for DO - DO 1"

System Schematic for DO- Colloid 5 1

ix

LIST OF ILLUSTRATIONS (CONT'D)

Figure Page

Water Electrolysis Bipropellant Thruster 52

System Schematic for ILO - 11,0 54

System Schematic for HO - Colloid 55

System Schematic for NÔ./N-H. - N_H, GO Plenum .... 57

System Scheniatic for C1F /N H - N2H4 GG plenum • • ■ • 58

System Schematic for DART - DART M

Propulsion System Weight for 3000-Pound, I 0-Year Satellite Without Power Penalty (>9

30. Propulsion System Weight for 2000-Pound, 5-Year Satellite Without Power Penalty 70

31. Propulsion System Weight for 3000-Pouncl. 10-Year-Life Satellite with Power Penalty 71

32. Propulsion System Weight for 2000-Pounfl, 5-Year-Life Satellite with Power Penalty 72

3 3, Logic Diagram for Computer Program ')2

34. ACS Computer Program and Input Data Card Arrangement . 122

23.

:24.

25.

26.

27.

28.

29.

'" ■ ■■■ ' "--•»" I'

NOMENCLATlilU-

D

g

bit

t ac

AM

1c

XX

yy

zz

en

j

L

M

M e

M o

M sp

P

AP

projected area in a plane normal to the line (if hiuln, it"

= diameter of satellite renterbody, feel

- gravitational aeeele ration, il.l It/sii"

- impulse bit minimum, II) -sei

total impulse for altitude eonlrol, 11> -se»

= total impulse for attitude maint< name, lb -se»

= total impulse for north-south slaiionk«-« ping, II).-so

= total impulse for east-west stalionkeeping, ll)f-s.i

- total impulse for limit cycling, Ib^-si«

total impulse for solar pressure corrections, II) -.•>>■•

- satellite moment of inertia about x axis, slug-ft"

= satellite moment of inertia about y axis, slug-ft )

- satellite moment of inertia about /. axis, hlu^ ft )

polar moment of inertia, sluji-fl"

length of satellite centerbody, feet

mass of centerbody, lb m

mass of satellite at enfl of propulsion iiwuuuvi r, lb 11 ID

- mass of satellite at bi-ginning of propulsion maneuvn-, lb

mass of solar panels, lb Ml

orbital period, degrees/day

repositioning rate, deg rees/day

XI

- II ! I ■ , -■"■-" - T mm

-^—.— 1 - ..-.—"_.,_,— " ' •' '-—-—

NOML-NCLATURE (CONT'D)

1/1

ST

LT

SST

\- III, . I.T

V-THST

in

V o

AV

X

X

rt-p

CH

xsv

cv>

SI-

CI'.

propt-llanl storam* tank pressure, psi

propellant storage lank radius, inches

n.oiiunt arm of thruster couple, feet

larm' ihrustcr Iced system reliability

small thruster feed system reliability

total large thruster system reliability

total small thruster system reliability

large thruster-valve combination reliability

small thruster-valve combination reliability

storage tank thickness, inches

mission duration, years

nominal orbital velocity, ft/sec

i hange in satellite velocity, ft/sec

i hange in velocity for repositioning, ft/sec

solar CP-CG offset, feet

for rectangular centerbody, the x-axis dimension, feet

freight of solar panel (distance parallel to centerbody, feet

for rectangular centerbody, the y-axis dimension, feet

length of solar panel (distance perpendicular to centerbody), feet

for rectangular centerbody, the z-axis dimension, feet

leadhand half-angle, degrees

Xll

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NOMFNCLATUUE (CONT'D)

nun

drift in satellite position cross-track, nm

drift in satellite position in-track, nm

satellite minimum achievable average angular rate, deg/set

propellant storage tank material density, lb/in.

yield stress of storage tank material, psi

i

xiii/xiv

SECTION I

INTRODUCTION

The task of evaluating future attitude control propulsion development

progrcimd for satiTlite applications is unwieldy clue to the proliferation of

systems concepts over the past 10 years. Since budgetary constraints

limit the amount of dollars for new technology efforts, a method for

selecting the most promising areas of future satellite propulsion work is

needed. Past evaluation methods have employed fragmented examinations

of various engine performance parameters such as specific impulse, pulse

centroid repeatability and minimum impulse bit (Reference 1). No recent

comparisons on a total system design basis for a specific mission have

been made. Only one other satellite system study has been undertaken at

the Air Force Rocket Propulsion Laboratory (AFRPL). This study was

completed on 20 May 1968 by Mr. E. C. Barth. It was entitled "Applications

of DART for Space Relay and Data Management Satellite. "

The present study uses an advanced geosynchronous mission model

having stringent attitude and station maintenance requirements to compare

16 satellite propulsion systems, in various phases of development, against

such important system design parameters as propulsion system weight,

system reliability and costs. These ranged from the conventional

monopropellaht hydrazine thrusters to more sophisticated electric ion

thruste rs.

I/.

SECTION II

APPHOACII

'I'lii' i oust nulion of a post-I 976 satilliie mission modrl was liascil

primarily upon existing model availability. From tin- st-vi-ral salrllitf

moflcls postulated for geosynchronous orbit, the Air Force's Spate and

Missile System Organization (SAMSO) model pertaining to the i lass of

satellites referred to as "SYNCSATS" was chosen as the framework for ilu.s

study (Reference 2). SYNCSATS pi ivide for a wide variety of eommeniai

uid military missions, including co imunication relays, navigation aids,

and meteorological and strategic reconnaissance.

So that a large variety of satellite propulsion systems could be readily

evaluated, a computer program (sec Appendix A) using the SYNCSAT

mission model was developed to calculate total propulsiun system weight,

mopellant tank sizing and mission total impulse requirements. The pro-

gram is also designed to size and weigh the satellite centerboclv and solar

panels and to compute the available on-board electrical power. Some of

'he SYNCSAT parameters which may be varied are the satellite life, initial

M ross weight, initial angular momentum, centerbody bulk density and

i-eposilioning rale.

In conjunction with the weight computer program, a reliability and

i ost study for each propulsion system design was undertaken. Reliability

data were extracted from a recent Jet Propulsion Laboratory (JPL) report

(Reference; 3) which arrived at quantitative satellite propulsion component

reliabilities based on a review of existing reliability studies and reported

component reliability and failure rate values. Reliabilities for noncyclic

i omponents were based on a 1-year mission duration. Improving reliabil-

ily figures through the use of redundancy was not assessed in this study.

II is to be noted that a quantitative ranking of uie components is difficult

.-iince reliability numbers for propulsion system components do not have

3

tho »'xliMi.sivt- statistical failure rate data typical of electronic components.

Dt-vclopment tost data for advanced propulsion systems are very

difficult to obtain. Moreover, a significant portion of the development cost

is expended for flight qualification. This dichotomy between development

and system enyineering groups compounds the total cost estimate. Instead

of expending many hours in an attempt to acquire every bit of cost data, a

rough lost estimate for existiny propulsion systems was undertaken by

using available figures from reported development and flight qualified sys-

tems. Postulated propulsion system costs were then extrapolated from

these existing estimates. Although the anticipated monetary inflationary

rate will alter cost estimates for post-1975 propulsion systems, the figures

used for this study are based on 1971 dollars. In addition to a quantitative

evaluation of system weigh»., reliability and cost, other tradeoff areas were

qualitatively evaluated. These included plume effects, integration problems,

design flexibility and ground handling requirements.

SECTION HI

ANALYSIS

A. POST-1975 SYNCSATS MISSION MODEL

1. Introduction

One of the most useful satellite orbits is the "earth-synchronous" or

"geosynchronous" orbit, i.e., a circular orbit in the equatorial plane

with an orbital period of one sidereal day. A satellite placed in such an

orbit will (ideally) remain fixed in the sky, relative to an observer on

the earth. The orbital characteristics for a "geosynchronous" orbit

are:

Semi-major axis, a = 22,808. 5 nm

Eccentricity, ( = 0

Inclination, i = 0

Period, P = 24 hours

The class of satellites having the above orbital parameters are

referred to as "SYNCSATS, " and cover a wide variety of useful missions,

both commercial and military. Many of these missions will require

extremely close pointing accuracy and/or precise stationkeeping. Such

requirements, coupled with a long mission duration, lax the capabilities

of current propulsion technology.

A three-axis active attitude control system (ACS) was chosen for the

SYNCSAT in preference to spin-stabilized or gravity gradient systems.

A fully stabilized satellite presents the most demanding propulsion

requirements, offers a significantly higher on-board electrical power

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capacity through the use of a one-degree-of-£reedom, sun-oriented

solar array, and can meet the close stationkeeping and tight attitude

control specifications. Furthermore, the three-axis active control system

does not demand the rapid pulsing capability of a spin-stabilized spacecraft

ami thus can utilize a wider range of propulsion concepts.

Z. Satellite Geometry

An inertial model of a SYNCSAT spacecraft was formulated to permit

propulsion system sizing and power allocation. The model does not

represent any specific design in this family of prospective SYNCSATS, but

merely a consistent set of dimensions and inertias. Geometrically, the

spacecraft centerbody may assume a cylindrical, spherical or rectangular

shape. The one-degree-of-freedom, articulated solar array takes the

form of two rectangular solar panels symmetrically deployed on either

side of the centerbody, which contains the remaining equipment. The

spacecraft configuration is shown in Figure 1. This figure shows the

centerbody as a cylinder. The spacecraft model incorporates bt-th high-

thrust engines of 5-pound thrust and low thrusters of less than 1-pound

thrust.

For the two solar panels, an ideal specific weight is required.

Combining this figure with a percent life degradation factor yields an array

specific weight. A specific surface area must also be assumed.

To arrive at dimensional and inertial characteristics for the model, it

was first necessary to size the solar array. The assumed available on-

board power can be given as a function of initial gross weight according to

the equation:

On-Board Power (w) = -400 + 1.25 x Initial Gross Weight (lb ] 0 m (III-l

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Once the on-board electric power is known, then the weight and size of

the two rectangular solar panels to supply this power are calculated.

From the centerbody weight, density and L/D configuration, the dimen-

sions of the centerbody are obtained. Next, calculated moments of

inertia for the three principal axes are developed.

For this study, the following design information was used. The

initial gross weights were assumed to be 2000 lb and 3000 lb . The u mm centerbody assumed a cylindrical shape with an L/D of 2:1. Existing

spacecraft indicate that a bulk density of 20 lb /ft is representative m r

for the centerbody. The solar panels were taken to be square with an

ideal specific weight of 88 lb /kw. The percent life degradation was

80 percent and f 3 solar panels were assumed to have a specific surface 2

area of 100 ft /kw.

For the two initial gross weights assumed. Table I presents space-

craft geometry data.

TABLE I. SATELLITE MODEL GEOMETRY

Initial Gross Weight (lb ) 2000 3000

On-board Electric Power (kw) 2. 10 3.35

Centerbody Diameter (feet) 3.833 4.375

Centerbody Length (feet) 7.666 8.75

Centerbody Weight (lb ) 1769.0 2631.5

Solar Array Area (ft ) 210. 0 234.8

Side of One Solar Panel (feet) 10.25 12.94

Maximum Projected Area (ft ) 239.4 373.3

Moments of Inertia (slug-ft ) 1 378.3 889.3 (Solar Panels Deployed) r

yy i zz

382.2

800.6

779.0

1796.8

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3. Mission Requirements

With the model spacecraft established, a consistent set of maneuver

and control requirements were taken from Reference 2. The propulsive

functions involved are of four types:

Initial positioning

Attitude maintenance

Station maintenance

Repositioning

Propulsion requirements for the foregoing maneuver and control functions

are now described in detail.

a. Initial Positioning (Injection Error Corrections)

The initial positioning errors are primarily caused by the launch

vehicle. Upon separating from the booster, the spacecraft will have a

residual rate (tumble) in each axis which must be nulled. It will also have

a terminal velocity and position error to be corrected. Eccentricity

and inclination errors need to be reduced only if the resulting oscillation

is greater than the allowable deadband.

Based upon existing booster performance, a AV allowance of

50 ft/sec for position and velocity error correction, and a 1 cleg/sec

residual rate correction in each axis would be nominal. For this study,

the inclination and eccentricity errors were assumed tolerable, and the

total impulse required to correct for the initial tumble in all three axes

was taken to be a constant value of 23 Ib^/sec for both spacecraft

weights.

At titiulf Maintenance

Altitude maintenance comprises the limit cycling within some

prescribe'l fii-adljancl and the correction of disturbance torques. The

primary c uitribiitor of disturbance torques is solar pressure. Assuming

unit reflectivity and normal incidence, the total corrective impulse which

must be supplied to the spacecraft in t years is (Reference 2); i m /

5.91 in v

j

Ai X lb -sec (111-2)

wne re: Ai total satellite projected area in a plane normal to the line of sight to the sun (ft )

where the summation is carried out for the two axes involved. Appropriate

values for the solar CP-CG offset, X, for different spacecraft were taken

as in Reference 2:

so that \ A V CP-CG OFFSET ^ Moment Arm

0. 35 in all cases.

The impulse involved in limit cycling depends upon a number of

factors. Maximum propellant consumption occurs in a symmetric

(undisturbed I limit cycle. Although symmetric limit cycling is not truly

representative for tins type of spacecraft, the conservatism implicit in

such an assumption floes not significantly distort the results and greatly

simplifies the calculations. The primary parameters in the total impulse

10

requirements are the half-anyle of the deadband, i, and the size pi the

minimum impulse bit, I, ... The total limit cycle impulse delivered in 1 bit ' t years is (Reference 2): m '

llc bit m V (—H lb -sec (111-3) A v j; k f k

where

r - moment arm of thruster couple (feet)

j = polar moment of inertia (slug-ft )

down to some minimum achievable averaue aninilar rate, 0 . , at which " 0 min point tlie rate limit impulse is (Reference 2):

0 t

Ic . mm mm

m S(-f-) lb -sec (III-4) n Deadband angles typically range from _J_ 0, 125 degree in coarse mode

control to _^ 0. 100 degree in fine mode control. This study assumed a

minimum achievable average angular rate of 2 x 10 deg/sec and the

deadband angle of _|_ 0. 125 degree.

For satellites with relatively large surface areas and long mission

durations, the effects of micrometeoroid bombardment must lie assessed.

Using probability theory based on the possible case, Aerospace

(Reference 2) has shown that the predictable impulse for micrometeoroid

impact correction is negligible. The unpredictable impulse requirement

due to a large and improbable impact must be provided in the "contingency"

impulse.

1 I

OthiM- rli.slui-bances, such as torques imparted by the friction in

moving telescopes or antennae, gravity gradient and earth magnetic field

torques, and coupling of translation thrust into the attitude control axes

caused by thruster misalignment with respect either to the spacecraft

center of mass or to each other, could not be accurately estimated

without a more detailed and sophisticated model. It was therefore decided

to apply a generous contingency of 50 percent to the total of solar and

limit cycle impulse allocations. Thus, the attitude control total impulse

vas assumed to be (Reference 2):

= 1.5 ac ulc

Ib.-sec (III-5)

c. Station Maintenance

The bulk of the spacecraft's propulsion requirement is for station-

keeping. A real earth SYNCSAT tends to drift from its initial position

radially, longitudinally and latitudinally (cross-track). These drifts are

caused by the triaxdality (asphericity) of the earth and by the gravitational

perturbations due to the sun and the moon.

At synchronous altitude, the observable angular deviation clue 'o radial

drift is negligible in any foreseeable mission. Thus, only in-track

(east-west) and cross-track (north-south) stationkeeping are required.

Table 11 lists the major perturbations on a 24-hour equatorial circular

orbit which cannot be corrected by initial injection bias (as reported in

Reference 4).

12

TABLE II. GEOSYNCHRONOUS ORBIT PERTURBATIONS

Perturbation Cause Direction Period Displacement ±V/yv

Zonal Harmonics -

t dependent

-

Tesseral Harmonics In-track Secular 7. 15

Solar-Lunar Cross-track Secular 5630 ft/day 150

Solar-Lunar In-track (€.) 2 years + 14.8 nm -

Solar-Lunar Cross-track {€ ) c Z4 hours + 8.9 nm -

A zonal harmonic results from the terms in the gravitational potential

of the earth that are dependent on latitude only and are therefore symmet-

rical about the equator. This is a result of the fact that the earth is not

a perfect sphere. However, drifts caused by this perturbation may be

corrected by injection bias. Tesseral harmonics are those resulting

from the aspherical gravitation field or inhomogeneous mass distributions

of the earth. Hence, one area of the earth will have a greater gravitational

attraction for a satellite than another area. These areas are not symmet-

rical about the equator and thus produce an east-west drift upon a satellite.

The solar-lunar perturbations result from the pull of the sun and the moon

on the earth. A secular perturbation is one which is not periodic but is

a constant perturbation dependent, for example, on the time in orbit. The

periodic solar-lunar perturbations need not be corrected if the displace-

ment shown is acceptable.

If the tolerable drift amplitude for stationkeeping is taken to be greater

than or equal to the periodic perturbations shown in Table II, L. e. ,

€. 214.8 nm and 6 -8.9 nm (referred to as the critical ellipsoid), then i c

an annual AV increment of about 157 ft/sec, dominated by the (.ross-track

correction, is required. However, if a "fine'1 stationkeeping mode,

i. e.,£. <14. 8nmand€ <8.9nm, is desired, the value for AV jumps lu

13

about b35 ft/sec/yr. This stucly will only consider an annual AV increment

of 157 ft/sec t'o r slationkeeping.

To accnunt Tor c nipling of thrust into the attitude control axes during

the attitude maintenance maneuvers, 3 percent of the slationkeeping tota)-

impulse requirement was allotted for this purpose. Thus, the attitude

maintenance total impulse was taken as (Reference Z):

[. 1. 03 am

[. [ lbf-sec (III-6)

d. Repnsitioning

Post-1975 SYNCSATS must be capable of covering any global region.

This implies that the satellite has the capability to perform transfers of

up to ISO degrees in longitude. For any given satellite thrust-to-weight

ratio and change in satellite longitudinal position, a minimum time for

repositioning can be determined. The velocity increment required is a

function only of the repositioning rate. The total AV required per repo-

sition is niven by the following equation:

AP r e p

V AP

(111-7)

V o

p

A P

nominal orbital velocity (ft/sec)

orbital period (deg/day)

repositioning rate (deg/day)

For the mission study of this report, a one-time satellite repositioning

maneuver was assumed to be representative for a post-1975 SYNCSAT,

and repositionint! rate of 15 den/day was used. The total AV required ' ' ' rep

for this maneuver using equation [11-7 is 280 ft/sec.

U

The most efficient method for repositioning is to place the satellite

into an orbit with a period greater or less than 24 hours, causing a

westward or eastward, respectively, drift. For example, the drift is

15 deg/day for an orbit with a 25-hour period and requires a AV expen-

diture of approximately 280 ft/sec for both high- and low-lhrusl devices.

Using this technique, repositioning requires from a few days to approxi-

mately 2 weeks.

A summary of the mission requirements used in this study is given

in Table III.

TABLE III. MISSION REQUIREMENTS

Function

Initial Positioning

Position/ Velocity Error

Tip-off Rate, Each of Three Axes

Stationkeeping

Attitude Maintenance

Repositioning

AV Requirement

50 ft/sec

157 ft/sec/yr

(E-W/N-S of 4 8.9 x 15 nm!

280 ft/sec

ACS Requirement

2 3 lb -sec

+ 0.125- deadband

legr ,2 x To-

deg/sec xv rr ^g1-- rate

B. SYNCSAT PROPULSION SYSTEMS

Sixteen different combinations of high and low thrusters wore incor-

porated into the spacecraft model and evaluated. The high-thrust engines

were 5 pounds and the low thrusters were less than 1 pound. The large

thrusters were used for initial positioning and repositioning, while altitude

15

pw'waspWBWWW»'»'^^

maintenance was performed with the small thruster. The thruster

(i.e. , large ur small) which had the highest performance was used for

stationkeepin^.

A listing of the 16 propulsion systems evaluated is presented below:

5 lbf Thruster

1. NJl , Catalytic

2. N H4 Catalytic

3. N-H, Catalytic

4. N.H r . '2 4

v ■tivu.j.y LLV-

5. N2H4 Cldtalytic

h. N21[-l C.italytic

7. K^ (Catalytic

H. KlH-\ Catalytic

9. X'11, Catalytic

10. m > H adioisotope

1 I. IK) U .uli Jisotope

H/) KK'Clrolyais I'ipr ipellant

11/ ) hllcctolysia iiipropollant

Small Thruster

N2H4 Catalytic

N H Plenum

N H4 Electrolytic

N H Resistojet

N^H. Radioisotope

DART

Cesium Ion

Colloid

Hg Pulsed Plasma

DO Radioisotope

Colloid

HO Electrolysis Bipropellant

Colloid

nioxyamine l(,

-- ...-- --.~ -- ■.■■»..- .- - - — ..-..: - ■.■-^-

■WM*"* i. " i •v^rnmmwmm. iammm*mmmmm mmmmmmwmmmmmmmnmm.wiixi mmmmmm^mm* ■' " .1

5 lbf Thruster

14. N 04/N2H4 Bipropel-

lant

15. C1F5/N H4 Bipr< pel-

lant

16. DART

Small Thruster

N^H, Plenum c 4

]NLH , Plenum 2 4

DART

The following subsections provide a conceptual design schematic for

each of the 16 propulsion systems. System description and performance

data are also included.

1. Hydrazine Catalytic/Hydrazine Catalytic

The development of the Shell 405 catalyst in 1963 permitted Ihr design

of hydrazine thrusters capable of a large number of restarts withuul

requiring the use of catalyst bed heaters or an oxidizer injection systfin

for initiation of hydrazine decomposition. Since then, monopropfl l-mi

hydrazine thrusters have become the "standard" spacecraft propulsi m

system for missions which do not have stringent orientation reqinn-/ < .r

and are not marginal on weight. Hydrazine ha.s excellent siorabil Us , ml

compatibility with most engineering materials and is capable f c .1 •■ pulse operation.

a. ^zlhf Thruster

Steady-state performance for the 5-pound hydrazine thrnster,

such as in Figure 2, was based upon 5b percent Nil, dissociation, MI .m

area ratio of 40:1 and upon 97 percent engine efficiency, giving 2-10 se> on«!

of delivered specific impulse. This performance represents nearls the

maximum achievable with hydrazine. Some hydrazine thrusters have

demonstrated steady-state firings exceeding 2 hours. The only requin ■:

power is that necessary to operate the propellant valves. However, 11,

17

'--—.^^^■^■.- aaBM • 1 1 '•""" ^^'"•""•"■"■'-- - ■ ■ ■ ■■•■-■-

i'liimw '""■ ■''-"'■ ' ■ <.»<•"•■'" ■ •^•"~n*r.J\\ij<..iixmmw*m*mimmm\u.uKimi... i .IHPIUH u'■nwin» ffffm^^g^mmit!^^ iMiJUiiiJin

U .J.-^.^... ^^.^^

iiiiiiwii mtfwmmr^^^^^^mmfmmmf^^ ^^^^^mm^***"! i mmm. Jl i,.lijll I.Jl.i., iPIJIi ■ IIÎ J». 11.! „„-,.^Î.^^^^7™„

certain installations, the low freezing point of hydrazine (33 F)

necessitates the incorporation of a 2- to 10-watt heater oneach thruster.

b. 0.1-lbf. Thruster

Pulse-mode performance of the small thruster was based upon the

same assumptions as those for the large thruster. Assuming that

heaters would be used on the catalyst pack to maintain temperature above

60 F, and using a minimum impulse bit of 4 x 10 lb-sec, a specific

impulse of 200 seconds is achievable. Although no flight-qualified

0. l-lbf thrusters have been built, a present NASA/Goddard development

effort for the Applied Technology Satellites (ATS), Models F & G, will

provide this technology. Hydrazine thrusters have demonstrated

pulsing capability on the order of 1-million hot starts. Several thousand

cold starts should be realizable without significant performance degra-

dation.(Reference 3).

c. System Schematic

Figure 3 shows the system schematic.

2. N^H^ Catalytic/NzHi Plenum

This hybrid propulsion system is a modification of the all-hydrazine

catalytic systern. Mydrazine plenum systems have been developed and

flight qualified by Rocket Research Corporation and TRW Systems. For

this design, the low-level thrusters are supplied gas from a single

catalytic hydrazine gas generator which feeds an accumulator or plenum.

The only system problem encountered with this hybrid system has been

that of maintaining a cool plenum temperature during a long pulse duty

cycle (Reference 3).

19

I ..J....^^^>-^,.-.,.^~^^sa.mak^wÂ!i^,.^,^..„ ■..,..„,..,. ■ ...--■- ■■:-l.-.-^->J..J..^-^ . - .-^^l

• •rt..'nu'm;nrwtKi'i'>'*'.-~ ■ »n i»i »UWMI wm**m "."HIIIüI'M.WJTI^W.II.H«, i.i(,a.n .i.i|i,i,fipiiiiiii«i n^?<g?^''!t'^^W°?l'aBtil^BiaBl^'g»T*l'8|WWW|ITWWy^^ FWWWW

Fill Valve

Transducer

Solenoid

Filter

Regulator

Fill Valve

Transducer (2)

Solenoid (2)

Filter (2)

Lines

Weight (pounds)

0. 25

0. 20

0. 20

0. 20

0. 3

0. 25

0, 40

0 40

0 .40

1 .00

Thrusters

Large

Small

V' Constant Weight

8(0.80)= 6.4

12(0.40)= 4.8

14.80

/\ LARGE THRUSTER SMALL THRUSTER /\

Figure 3. System Schematic for N^H. Catalytic-N H Catalytic

20

ii n i ii i ii i i n i i „t^^M^â^^^.^^^^^,,^^^ ^MMBMMB ■BMMMMMMMaMM ■■■ ^maä

r

^ 5-lbf, Thruster

The large hydrazine catalytic thruster is identical to that described

in Section III.B. 1. All performance numbers remain unchanged.

b. 0. 050-lbf Plenum Thrusters

The catalytic hydrazine gas generator uses Shell 405 catalyst and

feeds a plenum tank having a nominal 35-psia pressure. Specific impulse

for the 50-millipound thrusters'is taken to be a constant 110 seconds

for this study. Actual performance data for an N^H. cold gas plenum

varies between 95 and 110 seconds depending upon the gas temperature.

If individual heaters are used on all low-level thrusters, then the

specific impulse will vary between 114 and 132 seconds, again depending -4 upon the plenum gas temperature. A minimum impulse bit of 5 x 10

Ib.-sec was used.

c. System Schematic

Figure 4 shows the system schematic.

3. N_2Û Catalytic/N2H4 Electrolytic Ignition

The search for an efficient method of initiating and continuing the

decomposition of hydrazine without the use of a scarce catalyst has led

researchers to the concept of electrolytic ignition. The Air Force

Rocket Propulsion Laboratory (AFRPL) first determined the feasibility

of this approach through a contractual program with the Dynamic Science

Corporation in 1969-1970 (Contract F046 11-69-C-0048, Final Report

AFRPL-TR-69-247). Presently, the United Aircraft Corporation Research

Laboratory is under contract (F04611-70-C-0070) to AFRPL for develop-

ment of an electrolytic ignition cell for use in a 0. l-lbf thruster.

21

© —CÄH (p tij—I cäo

A LARGETHRUSTER

©

Weight (pounds )

Fill Valve 0.25

Transducer 0.20

Solenuid O.ZO

Filter 0.20

Regulator 0.30

Fill Valve 0.25

Solenoid (2) 0.40

Transducer (2) 0.40

Filter 0.20

Control Val ve 0.40

Pressure Switch 0.25

Gas Generator(GG) 0.35

Plenum 0.30

Filter 0.20

Lines' 1.0

Thruslers

Large 8(0.8) = 6.4

Small 12(0.4) - 4.8

Total Weight 16.1

SMALL THRUSTER

Figure 4. System Schematic for N?H. Catalytic-N?H . GG Plenum

22

Fabrication and testing of this cell have not yet begun. Since no test data

are available, all performance numbers are to be considered as "best"

estimates and will have to be revised in the future.

a. 5-lb. Thruster

The large hydrazine catalytic thruster is identical to that described

in Section III.B. 1. All performance numbers remain the same.

b. 0. l-lbf. Thruster

This study postulates a 0. l-lbf hydrazine electrolytic ignition

thruster having a pulse mode specific impulse of Isp = 220 seconds and _3

a minimum impulse bit of 5 x 10 Ibj.-sec. This size of low-level

thruster will require approximately 15 watts of electrical power excluding

that required for the valve. No life or reliability data are available.

c. System Schematic

The systam schematic is shown in Figure 5.

I 4. _N2H4 Catalytic/N-jH, Resistojet

The hydrazine resistojet has been under development by both AVCO

(NAS 5-21080) and TRW. AVCO has built and tested a number of proto-

type thrusters for NASA/Goddard Space Flight Center using a porous

ceramic injector configuration. Test data have shown that Isp is

23

A LARGE LS THRUSTER

Fill Valve Transducer Solenoid Filter Regulator

Fill Valve Solenoid (2) Transducer (2) Filter (2) Lines

Thrusters

Large

Small

8(0.8)

12(0.5)

Total Weight

SMALL A

THRUSTER LS

Weight (pounds)

0. 25 0. 20 0. 20 0. 20 0. 3

0 25 0 40 0. 40 0 40 1 0

6.4

6.0

16.0

Figure 5. System Srhcmatic for N H Catalytic-NH. Electrolytic

24

a strong function of thrust level (N0H, flow rate), so thai the designer c. 4

must be careful of his performance numbers. AVCÜ prototype data

are:

/ lb -sec \ ! Isp 1 1

Thrust (lbf) \ lb / m

2.4 x IG-3 120

4.6 x 10"3 155

7.1 x 10'3 177

9.5 x 10"3 190

12.0 x 10"3 200

14.4 x 10"3 206

j 16.8 x 10"3 210 1

1 Specific power for the AVCO prototype thruster is approximately

2 W/mlb^ (10 mlbj. TRW has completed the preliminary development

of a 0. 01-lbf-thrust hydrazine resistojet thruster for both pulsed and

steady-state operation. Reproducible impulse bits based upon pulse

widths as short as 2 0 milliseconds have been demonstrated.

The pulsed mode specific impulse is 180 seconds; steady-stale

operation results in a delivered specific impulse of 200 seconds. The

total power input for a 0. 010-lb^ thrust system is less than 5 walls,

excluding the valve power.

25

i

a. ^-Ib. fhrusto r

The la rue hyclrazine catalytic thruster is identical to that described

in Sccliun [II. 13. 1 .

11. 0. OSJKJJ). Thruster

Contractor in-house programs have arrived at a new wall injection

prototype thruster design incorporating a spiral-wound heater element

(Figure 6). This thruster has a moderately high chamber pressure of

H. 5 atmospheres and delivers 235 seconds of steady-state specific

inipul.se. The hydra/.ine resistojot can be pulsed as low as 50 milli-

seconds and deliver an average of 190 seconds Isp. These values yield -3

a minimum impulse bit of 2.5 x 10 lb -sec. The new prototype thruster

requires 5 watts for approximately 1 minute prior to ignition. This

electrical input raises the wall temperature to 1000 F.

c. System Schematic

The system schematic is shown in Figure 7.

5' N H Catalytic/N,H . Radioisotope

Both General Electric Company and TRW have developed radioisotope

thrusters using either NM^ or H as the propellant. Very little technology

has been expended on a hydrazine radioisotope thruster. Since N-,H

decomposes exothermally, the isotope power required is considerably

reduced from that of an ammonia radioisotope thruster. Therefore,

thruster inert weight and power required will be less for the hydrazine

system. To achieve long life, capsule temperatures will be 2000 F or

less. The design of such a thruster will be very similar to the DART

system in tiiat there is a radioisotope, re-entry heat shield, propellant

flow tubes, and thermal insulation (Figure 8).

26

n

u o

H

(A o

(X

(VJ ^

u

IM

27

■ ~~ "•"" —■^^WPllBillWIfWWWI^^Pil—^-' ' ■ " ■'■' ii »•■nn "'' ' 1"1"1 II .mi! » 'uu i. J"_ll ■«i^. ■ ^lll, in iifuui.ni.iiM . vmpii

A LARGE

L\ THRUSTER

SMALL THRUSTER

Fill Valve

Transducer

Solenoid

Filter

Regulator

Fill Valve

Solenoid (2)

Transducer (2)

Filter (2)

Lines

Thrusters

Large

Small

Total Weight

8(0.8)

3(0.45)

Weight

(pot inds )

0. 25

0. 20

0. 20

0. 20

0. 3

0. 25

0. 40

0. 40

0, 40

1 0

6.4

1.35

11. 35

A

Figure 7. Systcin Schematic for N?H Catalytic-N-H . Resistojet

28

— ■ -• iai

'"-ii - ■■»■.■iiiiiiu. i. .. ffgmggfjqgfqiiifjgfgjgsqfiîgffiipftffifififfiiiri^^

u 0)

■M en

u X H

O

o

CM ^

co

■r-(

>.<*

kitf^s•--»tiÛ "■•■-fc"'''-1 ■ ■- - jj - .-t.^,..v ^■,-

mrvwiwrr'.'■ • - ^ uviii.n<i»v»'fi'.'y-'''"-l'i-B.iw.iwi>rM)îvyiw?iywi«ilw'ij'.Tiwyi^ffiiii«iii.ip,ii-ii WKm!fn'ril!tl^llSllfSlieS^ftlfmi&^W!^SÎ9l9KWI!llflf^l!IIV Wf WWWBÔTHBIIimUPWWPBB'HWff

5-lb Thruster

The large hydrazine catalytic thruster is identical to that

described in Section III. B. 1.

b. 0. 0Z5-lb Thruster

For a thrust level of 25 niillipounds, a minimum impulse bit -4

of 5 x 1U lb -sec was used for limit cycling. Pulse mode specific

impulse was taken lo be 250 seconds on the basis of a capsule temperature

of 2000 0F. (Steady-state Isp is 220 seconds. ) A vented capsule will be

required to achieve the 10-year life requirement.

c. Conceptual Schematic

Figure 9 shows the conceptual schematic.

6. N H Catalytic/DART

The decomposed ammonia radioisotope thruster (DART) has been

under AFRPL-sponsored development with TRW since 1965 (AF04(6ll)-

11536). A DART prototype thruster was demonstrated at the AEC Mound

Laboratory in January 1967. An advanced DART prototype has been

designed by the Los Alamos Scientific Laboratories and is presently

undergoing evaluation. Since 1969, DART has been a part of the SAMSO

ADP for Advanced Satellite Propulsion, and a current $50,000 study is

concerned with problems associated with spacecraft integration.

a. 5-lb- Th rüste r

The large hydrazine catalytic thruster is identical to that

described in Section III. B. 1.

30

— ...

i-,^!!^.^!»!,!!»!^^1'/'"^/!.'^! »IIIIJ. I.JIilJH IJÎ.Ul,.P.I.JJy..11J,JljlHJll')HI^).f.!ll,1l>'H.>"'.-!'JM'.'-"-^

Fill Valve

Transducer

Solenoid

Filter

Regulator

Fill Valve

Solenoid (2)

Transducer (2)

Filter (2)

Lines

Thrusters

Large 8(0.8)

Small 12(1.5)

Total Weight

Weight (pounds

0. 25

0. 20

0. 2 0

0. 20

0. 30

0. 25

0 40

0 40

0 40

1 00

6.40

18. 00

28. 00

SMALL THRUSTER

Figure 9. System Schematic for N H Catalytic-N,H4 Radioisotope

31

ai-w-iÂ..^-.-.--.. .- .■■-.-J.--—- -.... . ,, . ..... . .. ■ - . . . J ■ - - ■■• '

. „■»■IJ.IMIIU.MPMI.IIIIU-I. <> inMmmmm'mnmnfmmmvwv I i. t I , ,mhnmUm.li»W»w U". .H"i.Hu"HiMH'im-..—■-.■•—- ■.-.-.-- ...-r^r^.^. „.„„„„v. V

b. 0. 0Z5-lbf Thrusler

Performance for DART was based upon a capsule temperature

of <i500OF. For a thrust level of Z5 millipounds, a minimum impulse bit .4

of 5 N: 10 lb.-sec was used for this study. Specific impulse numbers are

310 seconds pulsed, and 280 seconds at steady-state.

c. Design and Conceptual Schematic

Figure 10 shows the thruster design and Figure 11 is the

conceptual schematic.

7. N ". Calalylic/Cesium Bombardment Ion — 2 r

The electron bombardment engine uses an anode-cathode arrange-

ment to ionize a propellant such as niercury or cesium. The ions are

accelerated in an electrostatic field and neutralized as they are emitted to

avoid the limitations of space charge flow (Figure 12). While the ionization

potential for cesium is less than that of niercury, the cross section for

electron-atom interactions for mercury is greater than for cesium. The

result is that both propellants are equally easy to ionize.

NASA/Lewis Research Center mercury bombardment thrusters

(Kaufman thrusters I have flown on SERT-1 and SERT-1I satellites. An

Electrical Optical Systems (EOS) cesium bombardment ion engine has been

tested as an experiment aboard an Air Force satellite. Cesium is easily

handled by passive zero-g feed systems (Figure 13) and has a high mass

utilization efficiency as long as the cesium is kept above its freezing point.

Due to long start and shutdown transients, high-frequency pulsing is not

practical for the ion engine. Also, power requirements are extremely

sensitive to thrust and range from 15 watts at 10 [J.-lbf to 1300 watts at

1 0 mlb . tli rust. t

32

■ ~"*J™"

o V)

'J

s: H H

Ci

Fill Valve

Transducer

Solenoid

Filter

Regulator

Solenoid (Z)

Fill Valve (2)

Transducer (2!

Filter (2)

Solenoid (2)

Lines

Thruste r

—C^J-^ Large 8(0.8:

/T) Small 3(12)

Total Weight

Weight (pounds)

0. 25

0. 20

0. 20

0. 20

0. 30

0. 40

0. 50

0. 40

0. 40

0. 40

1. 0

6. 4

36. 0

46.65

LARGE THRUSTER

SMALL THRUSTER A

Fiwurc 1 1. System Schematic for N^ Catalytic-DART

34

in

ß W

tn o U

D

35

J

?

RESERVOIR

VAPORIZER WICK

Cs VAPOR FLOW

^-VAPORIZER HEATER

figure 13. Cesium Feed System

3 A

a. 5 - lbf Thruste v

The large hydrazine calaJytii I li inistc r i .s idnitiral to iluii

described in Section III. B. I.

b. 0. 001-lbf Thrus'er

The 1-millipound thrust sizr lor llic luw-lrvr] lliru.slcr was

chosen to keep the power requirements around ISO wails. AJlliou^li KOS

life-tested two S-miliipound bombardment ion I h rust e rs, their power

consumption was on the order of I kilowatt. For a l-millipound thi-ustcr

and 150 watts power, a specific impulse of 3000 seconds is projecird. A -3 minimum impulse bit of 1 x 10 Ibf-sec was considered reason,ddi-,

based on a 1-second minimum pulse width. (The above performance goals

were obtained from Dr. Fritz, AFAPL/POIJ-2 of the Air Force Aero

Propulsion Laboratory. )


Figure 14 shows tiie conceptual schematic.

8. N9H. Catalytic/Colloid

The basis for the colloid engine (Figures IS and Kx is an ein I id-

eally conducting propellanl subjected to a high electric field < stablished

between the propellanl and an extractor electrode. The extrai tor ejet (rode

has historically been a small-diameter (4-niiJ bore) capillary needle, but

recent developnient effort has inen expended on a linear slit geometry

electrode version. Once ihv electrode field is established, field emission o

ionization of small-diameter (loo A) droplets oi < urs at the needle lip or

linear slit. The same field wiiich produces ionization also a« i derates tin-

charged droplets to produce thrust. Since the v harLiefi drtjplets may be

positive or negative depending on the polarity of the potential applied to

57

Weight (pounds)

Fill Valve 0.25

T ransdvu-er 0,20

Sulenuid 0.20

Filter 0.20

Regulator 0.3

Solenoid (2) 0.4

Fill Valve (2) 0.50

Transducer (2) 0.40

Filter (2) 0.40

Solenoid (2) 0.40

Lines 1. 0

Thrusters

Large 8(0. 8) = 6.4

\ Small 12(5.

;ht

5) = 66.0 /

Total Weiy 76.6E

LARGE THRUSTER

SMALL THRUSTER

Figure 14. Sy System Schematic for N-,^ Catalytic-Cs Ion

38

1

00

-I oc

o- u O UJ OC -I a. LU

f-TT cc

s > a ca a. CM

00 Q

< ^

LU

o UJ =3 " I- O r»

I 0 2! -- u r: «■

C W

o U

j

I

39

1. .,JM,,_ . ^»^JJ*^.'.' .^Ji

^W^PPO . II. I. .JiiMMpp

•"'""•■"i Man

1

CAPILLARY NEEDLE 14-milO.D. 4-mil BORE

PROPELLANTIN

fk

EXTRACTOR—, ELECTRODE

GROUND SHIELD

Uidijtikittimammjmi '■■••• - -^>^^

5-10kv

NEUTRALIZER

Figure 16. Colloid Thruster Concept

40

'■■'■" ' -n "-■-' i i*iii n i

' '^-^■■-' H.M . . i^.UU.,^-^,^^,.,, ,^ ^^ _

the needles, a neutralize r is IUM r ssa ry lu inui i-.ili/.r i'H- IM an,, Mn i .,1.-,... .-.

of the charged droplets an; generally ^rraler than tin nuissi's ol imiM

produced in ion engines.

Although the colloid engine is d' graded by the raniloinne.is M i: ■

particle formation and the manner ol i idu« in^ the . iiarê, n al^m , <>\ 'i •

electrostatic engines, lias ihe most ell'i« i« n' loi-nuiiion D! . harued i.iii;. le ..

The most successful colloid engine v.urk lias l>e. n ;M r in i n .< : !.\

TRW Systems for the Air Forie Aero i'l-oiadaioii l.alaj r.ii ory (.\i Al'l.i.

TRW built a Colloid Microlhrusler i:.\|>enn.eni ((.Mil for th. Al AIM. in

support of the DODGE-I1 satellite, which was i. am elled. liu ('Ml ili-i.i

hardware was subsequently tested for lonn hours during \')>''),

Since 196V, the TRW colloid thrusti-r has been pari ol Ih. SAMSi;

ADP for Advanced Satellite Propulsion. In l)e. ember l')H), 1 H W was

awarded a 56-month contrail by SAMSO for devilo|)menl ol a !-milli imund-

thrust colloid ihruster. The thruster, i-is<mljliiiu ,i l()-in< h > nbe with

12 individual thrusting modules, will wei^h uliout <i() pounds and tarry t>'>ii,e

25 pounds of propellant, mainly glycerol with sodium io<lide. Ai il .

1-millipound thrust level, the colloid engine will liave a SDC. ili. MI. .i,].-,.

of 1500 seconds and deliver about 55,()()() lbfsei of t^t.il impuUe. 1 i,.

contract calls for ground testini; of ilim lliuhl cpialified thrusi. r^ for

10,000 hours and delivery of three systems for satellite fli^lit tesiinu to

provide satellite slationkei-ping 1\. >• 7 yia'-s.

a. 5 - lb, Thruster

The large hyd razine eatalyt it i h ru&l e r is idi-ntu al to t ,,ii

described in Section 111. B. I.

^Space Business Daily, 2 December IPVI

—— . —' rr- '—• •m ' '■"■J - - ■'"■ '"■ ■

I). P. 01)1-lb Thruster

Tlu" curnnt SAMSO ADF for colloid propulsion was used to

obtain propulsion system characteristics. For this study, thrust level

was I millipound, Up was 1500 seconds and minimum impulse bit was

I x 10 lb.-sec. The propellant is 20 percent Nal and 80 percent glycerol

(pe ri i-nt by weighl).

A lightweight and reliable feed system remains the largest

development effort. No pulse tests have been performed on the linear

slit to date. The needle has been pulsed between 1 and 3 pulses/sec in

duty cycles of 10 and 30 percent. Although the low power required

(70 watts) for the colloid thruster is favorable, the linear slit requires

between 14 and 16 kilovolts for operation and is the weak point in the

system.

i . (Conceptual Schematic


9. ^ i' 1 , Catalytic/l ig Pulsed Plasma

Although several types of pulsed plasma thrusters have been

undergoing exploratory development (Figure 18), only the pulsed vacuum

arc thruster (Figure 19) being developed by Cornell Aeronautical

Laboratories uses a liquid propellant (mercury) which permits this sys-

tem to achieve a total impulse level required for a SYNCSAT (USAF

Contract No. F 336 1 5-67-C - 1 579, R eport AFAPL-TR-68-92). The general

mode of operation is for niercury to be ionized by a high-voltage discharge

and accelerated by the interaction of the discharge current with its own

magnetic field. The PVAT produces only discrete impulses, the effective

"thrust" being governed by the size of these impulse bits and the repetition -6 -5

rale. Typical impulse-per-pulse figures range from 10 to 10 lbf-sec

while repetition rates from zero to 50 pulses/sec are readily attainable.

42

-■-. .-—.■lI-^lA-.-r-.^.l—j--.^ r .;■■■»■:■ ,,"'*d

iim9.m»m-mimmwmim^ iw9iMßmm}».mrMjuiummmm^m^mmmiv^mwn'v.mmK -1 ■ mgppr^glBfgww—»M...,. «mm.im^.m*, vwiinymmw,:-'r-'.~'w'"""■*"■

Weight (pounds)

lill Valve 0. 2 5

T rans'luicr 0. 20

Solenoid 0. 2 0

Miter 0. 20

Ret^Lilalor 0. 30

Solenoid (I) 0.40

L ill Valve (I ) 0. 50

Transducer (2 1 0. 40

Killer U) 0.40

Solenoid (2) 0. 40

Power Clonditione r 1 1 . 00

^\ Lines 1 . 00

L l) iDj/

Thrus te r

-th- 3 Large «((). 8)

8)

= t..40

9. (-0 '\" J Si i ifiil 1 .i ( Ü.

Total Weight l 31.25

LARGE THRUSTER

SMALL THRUSTER

Figure 17. System Sthnnalic for I^Jl. Catalytie-Colloi<

4 3

-■-■- ■ - ■■■■- -■•- ■■■-■■ -^^..-.- -'■'- ■-■■■ ■'- .:..--...■ ■ - - ■-■■. -..■-■

Weight (pounds)

Kill Valve 0.25

T ransduccr 0.2Ü

Sulenoul 0.20

Filler 0.2 0

Regulator 0. 30

Solenoid (2) 0.40

Fill Valve (2) 0. 50

Transducer (2) 0.40

Filter (2) 0.40

Solenoid (2) 0.40

Power Cond itioner 11.00

Lines 1. 00

Thruster

-C^H) Large 8(0.8) = 6.40

(?) Small 12(0.8) = 9.60

Total Weight 31.25

LARGE THRUSTER

SMALL THRUSTER

I'-Ml Vi' I I System Schematic for Nil Catalytic-Colloid

43

faMM- —■ - - -- ■-^——--—^ ÜÜÜHiiiHiliäi

CERAMIC IGNITER

MAGNETIC FIELD

ANODE

CATHDDEPLATE

INSULATOR

FLANGE

TO MERCURY

FEED SYSTEM

Figure 19. Pulsed Vaiuuiu An I li ru.sl e r

4 5

h MUM ■/..■-.- , ji.jj^ mmm M^^M^^^MMMM

Lifrlimos up to ti x 1Ü iîilses have been dentonstrated on several pulsed

vacuum aw ihruslers (PVATs). To date, the mercury-cathode PVAT, 5

lunvi'ViT, has been run for only about 4 x 10 pulses because of feed system

problems.

a, 5-11) Thrustor

Performance for the N.1I, thruster was Isp = 230 seconds

slrady-state vacuum. All menibcrs are identical to those previously used

in Sect ion 111. B. 1.

b. 0. 00001-lb Thruster

Data used in this study were based on the PVAT being devel-

oped by Cornell Aeronautical Laboratories. This study used a thrust

level of 10 micropounds force, a specific impulse of 1500 seconds and a

minimum impulse bit of 5 x 10 lbf-sec.

The PVAT still needs considerable development in the feed

system area and more life testing of an integrated system. For the

SYNCSAT, power r quirenients are on the order of 100 watts for each

PVAT.

v. Conceptual Schematic

Figure 2.0 shows the conceptual schematic.

10. DO Radioisotope/DO Radioisotope

DO (dioxyaminc) is a new type of monopropellant presently being

characterized by the AFRPL. Because of the high (2600oF) equilibrium

flame temperature of DO, it is difficult to project the early development

of a catalyst and substrate to decompose DO. For this reason, a radioiso-

tope capsule is envisioned as the ignition scheme to initiate thermal

decomposition for the purposes of this study.

46

luH^.H.Jr.v.. ...,.■■■..■■■■-;•:.-<-.II.I. :■•■'.. ..—■■..J^.-i J^^.^-.-..iu...U^ i. ■!■ i.-- ui -» .--■ .i.i ..i ■■-■ •-■■--'■■•■■■ ■■--■ -i..--.-. ■ .I..:..-. !iC*>-,2: ,.-_:_!_.■./;. ■ ■■ ■ ' ..■.■•■ . - .■■■ --,..-.... ^-. - ... ■ , fc | , , „ ■ ......... | 1 | |H Ulrt—^rMJ

A LARGE /ATHRUSFE

Wriuil (pianui

Kill Valve 0. .1')

Transilm o r 0. ^0

Solunuifl 0. JA)

Filler 0. .M)

Ueûlalor 0. M)

Solcnoiij (Z) 0. 10

Fill Valve {.'.) o. so T ransfiueer (i ) 0.-40

Filter U) 0. 10

Solennul (.1) 0. 10

I'D wer (li)nciilii mer 5. 0

Lines 1. 0

I'll ruslc rs

l.artie «(() . «1 '). 1

Small li^ ) .'. 1. 11

'F.tal \\ riuhl '■'>•'>, 1,1

SMALL THRUSTER

'imin iO. System .Schemalir for NJI, ('atalytic - 11^ Fulscd Plasma

47

i ■nia'iMiiii—miaut •■■-■■- ■-■■■■ i i - ..-■ --^.■-^---.^—^-^---—■_....■■-..■.-;..■..■. ....,■..-■.. .... ... ...- , .. ■ mfl .■■J^;:.....J;..ÂIJ^.^.vA^..vl,..-.:.-J.-..:t-..i..;^....J^.,^.J.--...^l:>..»^:-^..--:.^^-.^ ^^^...-l-. ^■.■. ■■ ^JMiM*

The AFRPL pi; r for mane u calculation computer code (ODIE) was

used to calculate a theoretical altitude performance number for DO usiny

a combustion chamber pressure of 100 psia, a nozzle expansion ratio of

i 40 and a 55 percent NIL dissociation. These inputs gave a theoretical

steady-stale altitude specific impulse of 283 seconds. Using a 94 percent

efficiency factor, a realizable specific impulse of 265 seconds might be

obtained.

a. ^-Ih,- Thrust»

Steady-state performance for the 5-pound DO thruster was

taken as 265 seconds, as mentioned above. This number is based on

94 percent engine efficiency and is a "best" estimate. No engine data are

ava' table.


For the 25-millipound thruster, a minimum impulse bit of .4

5 \ 10 lb.-sec was assumed, based on the DART ADP goals. Pulse

mode performance is Isp = 238 seconds based on a capsule temperature

of 2500 'F. No lest data are available.

i . Conceptual Schematic

Figure 2 1 shows the conceptual schematic.

11. DO Radioisotope/Colloid

This is a hybrid propulsion system comprised of two previously

described units. A radioisotopc thermal decomposition ignition system

is postulated for the 5-lb DO thruster. The colloid thruster is of

l-millipound thrust with performance numbers as described in the SAMSO

ADP for the Xal plus glycerol colloid propulsion development effort.

48

Ktasnrv&L:..■.,■..:.. ^. - j,,^.;., r......_./..-■-.-■■.i...--..„...„..,;.-_..•..,.. :_,>.^.., . ....,.-..M^J5:-,t .■^..■.,^.,.,-. . ...:;l^,'..^. .„,. :■.,-.■: i^-.il ** **..'.■.: - ;. -. ■>.„ ■..■■., .■..:^j.î/. i^J ^l.,;.^ :■...-,:.. ;. .■,...1-i-.,.l-.-:■ , „■[, i 1,1 I I 1 I ■" ■lillîi

Fill Valvi-

Transducer

Si)leni)itl

Filler

Hi'milalor

Fill Valve

Solenoid (l)

Transfluccr (Z)

Filter (2)

Lines

I'll ruste i-s

Far^c H(Z)

Small 12(1. SI

Total Weight

\V.MK| pom,

0. .'1

it s

0. .'.(

0. ,'C.

0. ,'(

o. •; (/

0. .":

0. .-(!

0. ■'.()

0. (;

h.. w

FS. ■

LARGE SMALL THRUSTER THRUSTER

Figure 21. System Schemalu- for DO-DO

49

^. ^ .■>.,.-...■ . ^.„^ ..^ ...■;, .^ âaato îMijjm* ^^.*.^-..^^^-^.^^^^*.^.~. -^ ^^fc:. -:^^.. ■ ^ ^.^î >^ .,■ J;..: .^ . ;,w,.. - ^ Lr.. ict'>THrff1rttfr.,i- ;i, . ■c.-i.^.^^^xûij^t...

a. 5- lli. riu'ustor

I'lu' large DO radioisotopf t. ISILT is identical lo thai

«lesi-riljed in Seition lll.li. 10.

1). (i. 00 I-lb Thruster

The small colloid thruster is identical to that described in

Section 111. B. 8.

c. System Scheniatic

Figure LI shows the system scheniatic.

\l. II,O Electrolysis/llÔ Electrolysis u —————— ————^—^ ^ —

The water electrolysis propulsion system employs liquid water

as the propi.Tlant in the storage mode and gaseous hydrogen and oxygen for

the bipropellant rocket thrusters. A zero-g water electrolysis unit provides

the separated gases for the engine. This scheme permits storing liquid

water at low pressure and reduces the total system weight in this manner.

A recent AFRPL contract (F0461 1 - 71-C-0055) to the Marquardt Corporation

provides for the development and testing of a bipropellant 5-lbf thruster

and a bipropellant 0. I-lb thruster using gaseous hydrogen and oxygen.

General Electric is the subcontractor for the zero-g water electrolysis

unit. Twenty watts of continuous power are required.

a. T-lbr Thruster

Steady-state performance for the large 5-lb. GH?/GO^

engine (Figure Z3) was based on Marquardt prototype test data. This

number is Isp 3 50 seconds with a thrust coefficient equal to 1.6. A

pulsing performance of Isp - 3 10 seconds was assumed for this thruster.

Gil./GO, ihrusters provide high performance and highly reproducible

pulses.

50

LARGE THRÜSTER

Fill Valve

1' iMUbilurt.' r

Snlcnoifl

ilter

Regulator

Solenoid (<i)

Fill Valve (Z)

Transducer i2)

Filler (Z)

Solenoid (Z)

Lines

Power Conditioner

<D Thruste r

Larjê H(Z)

Small 1Z(.8)

Total Weight

\\ .■i-lii

Üii""1'''''

0. .'.'>

0 ..'. ()

0. .'. 0

0. zo

0. 30

0. 40

0. bO

0. 40

0. 40

0. 40

1 . 00

1 1. 00

U,. 0

'). (,()

40.85

SMALL

THRUSTER

Figure 22. System Schematic for DO-Colloid

51

i ^—SPARKPLUG IGNITER

PROPELLANT VALVE

Figure 23. Water Electrolysis Bipropellant Thruster

52

it

b. .O-iilb. Thruster

The small thruster pulse mode Isp was assumed to IJC

280 seconds. Using a 25-millisecond pulse width, a minimum impulse

bit of 2. 5 x 10 lb--sec was taken. Steady-state performance for the

0. l-lbf thruster was assumed to be Isp =310 seconds.

In general, this system lacks feed system and integration

testing. System weight reduction is at the t-xpense of system complexity.

A high system reliability must be demonstrated early in its development.



13. H^Q Electrolysis/Colloid

This hybrid propulsion system is designed to provide a high-

performance chemical rocket engine with electrical colloid thruster. Both

units have been described previously.

a. 5-lb. Thruster

The large GH-/GOp bipropellanl thruster is identical to the

water electrolysis system described in Section III.B. 1Z.


The small colloid thruster is identical to that described in

Section III. B. 8.


The conceptual schematic is shown in Figure 2 5.

53

—lÄ—3

h

—cJj-3

X"

-(5>

THRUSTERS

A LARGE

A SMALL

X-

Fill Valve (3]

Transducer

Solenoid

Filter

Electrolysis Cell

Transducer (2)

Pressure Switch (2)

—(?) Solenoid (2)

Filter (2)

Lines

Thrusters

Total Weight

Weight (pounds'

0. 75

0. 20

0. 20

0. 20

15. 0

0. 40

0. 50

0. 40

0. 20

1. 00

30. 00

49. 05

Figure 24. System Schematic for H?0-H?0

54

__, M^^MHM ^^HMMMiM mmmm

® X"

<D E

CELL

X-

A LARGE THRUSTER

Jr'ill (4)

Transducer (Z)

Solenoid (2)

Filter (2)

L^J J Regula lor

0 Eleclrolysis Cell

T ransrlvu:e r (2 )

Pressure Switch (2,

Solenoid (2)

Fill Valve

Filter (2)

Transducer

Filter

""£^H3 Solenoid

Power Conditioner

Lines f 8

SMALL THRUSTER

Til rusters

Large

Small

Total Weight

Weight (pounds

1 . 00

0. 10

0. -10

0. 10

0. 30

15. 00

0. 40

0. 60

0. 40

0. 2 5

0. 40

0. 2 0

0, 2 0

0. 20

11. 00

2. 00

15. 00

9.6

57. 05

i

Figure 25. System Schematic for ILO-Colloid

55

'"^'-- ■■■■^ft-i ,,— - :...■- . i ,..,.,■.._,.-!-■■: -o.. .-■--.: -- - -- ■ ..■>-■■ . ■..--.-—„w..-..v.. ■■.■•■>!■.. ■■>-.,.■.. iij.^.-.■■?-

14. N2,04/N2H4 Bipropellant/NzH1 GG Plenum

Although a 5-lb- bipropellant thruster using N?0. and hydrazine

has not yet been developed, a large number of 5-lb. engines using N_0 ./

MMH and N2O4/50 percent N H.-50 percent UDMH have been built and

tested. This system is designed to use the common hydrazine tank for

both the high and low thrusters, thereby providing for simplicity and weight

reduction. The earth storable bipropellants are well defined in terms of

properties and characteristics. The only required power is that needed

to operate the propellant valves.

a, 5-lb Thruster

Since no 5-lb,, thrust N00 , /N-H . test data were available, i Z 4 c 4

steady-state 5-lbf thruster performance numbers for 1NLO ./MMH were

used. This yielded steady-state operation at Isp = Z80 seconds for an area

ratio of 40:1 it a mixture ratio of 1.40.

b. 0. 050-lb. Thruster

The low-thrust hydrazine gas generator plenum system was

capable of giving Isp - 1 10 seconds and a minimum impulse bit of _4

5x10 lbf-sec. These numbers are identical to the hydrazine plenum

system described in Section I1I.B.2.

Conceptual Schematic


15. C1FC/N,II. Bipropellant/N.M. GG Plenum b ^ —t c,—4 ——^^——————

This bipropellant system is similar to the No0 . system described c, 4

in Section 111. B. 14, except that care must be exercised in the selection of

a CIF_ storage tank. A lank material compatibility problem exists which

56

Fill Valve

Transducer

Solenoid

Filter

Regulator

Solenoid (2)

Fill Valve (2)

Transducer [Z]

Filter (2)

Solenoid (3)

Control Valve

Press. Switch

G G

mi

SIWALL A THRUSTER

I Thruster

Iâi-ge

Small

8(1.30)

2(0.4)

We iglit (poiinds )

25

2 0

2 0

20

3

0.4

0. , 50

0. , -10

0. 40

0. 60

0. 40

0. 25

0. 3 5

Total Weight

0. 3 0

0. 2 0

1 . 00

10. -I

4. H

2 1.15

j

Figure 26. System Schematic for Ôj/N.,1 IÂy 14 GG ITenum

57

k

was nol fxpe ricncfd with NjO.. No performance data for a 5-lb. c -i [

GIF /N ,11, thruster are available. The only required power is that needed

to OIK'rate the propellant valves.

-i- lb . Til rust er

For this study, a 10-second specific impulse gain above the

N ü./iXjH system was assumed. Therefore, steady-state performance

was Isp - Z'.'O seconds at a mixture ratio of 2.0.

I). 0.05Ü-lb Thruster

Once again, the hydrazine gas generator plenum performance

numbers were identical to those used previously in Section III.B.2.

c Conceptual Schematic


16. ADP DART/ADP DART

The development status of the decomposed ammonia radioisotope

thruster (DART) has been described under Section I1I.B.6. The all-DART

concept has been favored by the AFR PL since 1968, and an in-house study

was accomplished by Mr. E. Barth to detail its potentials and mission

applications (Reference 5).

^- lb . Thruste r i

There is no 5-lb engine in this concept. All engines are of

the ().()2S-lb. class. Repositioning may pose a problem.

b, 0. 02^-11) Thruster

The ADP DART performance- goals were used for the all-DART

system. Thrust level was 25 millipounds and a 20-millisecond pulse

58

Wright (pounds

BIPROPELLANT THRUSTER

SMALL THRUSTER

Fill Valve Ü. 2 5

T ranscluce r 0. 20

Solenoitl 0. 20

Filler 0. 20

Regulator 0. 3

Solenoitl (2) 0. -1

Fill Valve (Z ) 0. 50

Transducer (2) 0. 40

Filter (2) 0. 40

Solenoid (3) 0. (,0

Control Val ve 0. 40

Pressure Switch 0. 25

G G 0. 3 5

Plenum 0. 30

Filler 0. 2 0

Lines 1. 00

Thruste rs

Large 8(1. 30) 1 0 -1

Small 12(.4)

t

4 8

Total Weigh 21 15

Figure 27. System Schematic for C 1 F-/ N-, 11 - N 11 GG Plenum

59

produced a mininmm Lni|)ulse bit of 5 x 1U Ibp-sec. Pulse niode

Isp 3i() seconds and steady-state !sp - Z80 seconds were used, based on

a capsule yas tempe ralure of 2500 'F.


Figure ZcS shows the conceptual schematic.

C. HROPFLSION SYSTEM WFLGHT

A computer program was developed to facilitate the task of calculating

spacecraft geometry, weighing the propellant required and sizing propellant

tankage and pressurization systems. This program will weigh the entire

propulsion system for a geosynchronous orbit with up to 10 different

propulsion functions being performed. The program will accept variations

in the following parameters: (1) satellite parameters such as initial gross

weight, life, initial angular momentum and repositioning rate, (2) center-

body parameters such as bulk density, geometry (i.e., spherical,

cylindrical or rectangular) and L/D ratios, (3) solar panel variables such

as ideal specific weight, percent life degradation, specific surface area

and height-to-length ratio, and (4) parameters for the small thruste rs such

as minimum thrust level, minimum pulse width, deadband half-angle for

accuraicy and the minimum average angular rate for limit cycling. In

addition to these, data on the propellant storage tank materials and their

properlies and storage pressures are required. The Isp which would be

realized during each propulsion requirement for a given propellant or

propellant combination is also required. Information obtained from the

program is; ( I ) solar panel and centerbody dimensions and weights, (2)

onboard power, (5) satellite moments of inertia, (4) impulse, AV and

propellant amounts for each propulsion requirement, (5) total impulse,

AV and propellant recpured for the mission, and (6) storage tank sizes and

weights and the amount of pressurant required along with a tank to store

it.

(>0

... ...>^L,î..::-..li--. i i i i.i jiwli.ini'iii LJIJIUH TWO^fit^'lii'itii'IJ*' T't'iinifn'iilfriitflît'ftTCîittrfWLlVVIfi

r HSff^Wfy; «^■w;-*? w.«.s3^R<pw?,.»«r J»."«*»'*»;« H-ri.i»j»*!*a*i n^»jrrT^TT»^^Tvn»^r™^T^r^T*rT7Ti^^^

Fill Valve

Transduce r

Solenuid

Filter

Regulator

Fill Valve

Solenoid (Z)

Transducer (Z'

Filter (Z)

Lines

Thruste rs

5(12

Total Weight

Wei^lit (pounds 1

0. Z^

0.2 0

0. 20

0. 20

0. 30

0.25

0. 40

0. 40

0.40

1 . 00

60

(,3. 60

f

THRUSTER THRUSTER

Figure 28. System Schematic for DART-DART

61

.t.l'ff<.^^^l'r?VMTH-W|.-îT^-Tru ..,.., >f-J.I*7^t)wi.y..; j ■■T^^..^....,..-,,,.^^^^^,^^^^^^,^^;^^^,^^^.^^^,^,,^^^^

For this study, four different missions were analyzed. These were:

1. 10-year life, 3000-pound satellite performing north-

south stationkeeping. (This mission will be designated as 10-3NS. )

1, 10-year life, 30ü0-pound satellite not performing north-

south stationkeeping, (This mission will be designated as 10-3. )

3. 5-year life, ZOOO-pound satellite performing north-south

stationkeeping. (This mission will be designated as 5-2NS. )

4. 5-year life, 2000-pound satellite not performing north-

south stationkeeping. (This mission will be designated as 5-2. )

Values assumed for other spacecraft parameters which remain con-

slant from system to system in this study have been discussed in

Sections lll-A-2 (Satellite Geometry) and I1I-A-3 (Mission Requirements).

Tables IV, through Vll present the results from the computer for the

above four missions. These tables give the propellant weight, inert

weight, total weight of the propulsion system without a power penalty,

power penally and the total system weight with the power penalty.

All of the systems investigated in this study require some power;

some require more than others. All systems have solenoid valves in the

feed systems and these require power to operate them. Since the power

Ii'V.ls of these arr so small, these power requirements were ignored in

assessing a power penalty. It was also assumed that the hydrazine sys-

lems would not require heaters to keep the hydrazine as a liquid. There-

to n-, power penalties were assessed upon the water electrolysis,

hyu ra/iiu' nsislojet, colloid, cesium and mercury pulsed plasma systems,

''ov.-rt- MiMialties used are presented in Table V11I.

p mammmmmmmmmmmm** f mmmmmmmmmmmmmmmmmmmmmmym mm m\m\imm* iP-_piMf.iii'i"wyf.'ûijnn.i ma

TABLE IV. PROPULSION SYSTEM WEIGHT FOR 3000-POUND SATELLITE WITH A 10-YEAR LIFE PERFORMING

NORTH-SOUTH STATIONKEEPING

_ High Thru sie r

Low rhruste r

Propellanl Weight

Inert Weight

Total Weight

W/C) Pov/i r Powe r

Penalty

Total Weight

With Powe •

H20 Colloid 200.638 65.604 266.242 94. f 360.842

N2H4 Cat ^ Ca Ion 188.820 89.097 277.917 198. 0 475.917 1

DO Colloid 227.357 54.655 282.012 92.4 374.412 j

N2H4 Cat. Colloid 243.969 47.083 291. 052 92.4 383.452

N2H4 Cat H« P.P. 2-13.969 50. 181 294.150 132. 0 426.150

H20 ■v 622.212 66. 969 689. 181 2.2 691.381

HART HART 621.055 1 12.420 733.475 - 733.475 !

N H4 Cat DART 640.693 97.554 738.247 73H.247

N21I4 Cat. N2Il Radioisotope 751.936 65.551 817.487 817.487 1

C1F5/N2H4 N2II GO Plenum 776.006 57.489 833.495 - 833.495

N204/N2H4 i;2H4 GO Plenum 794.055 60.957 855. 012 - 8 5 5.012

N H /Cat. N2H4/Resistojet 826.525 51.933 878.458 6.( 88 5.058

no DO «50.282 73.368 923.650 - 923.650 j

N2H4/Cat. N2H4 Cat. 1013.323 02.811 1076.134 - 1076.134

N2H4 Cat. NpH Electrolytic 1031.291 64.715 1096.006 19.8 1115.806

N2H4 Cat. N2H4 GO Plenum 1070.732 66.7 32 1137.080

i

- 1 n" )86

t

N ,11 Catalytic

I Ig Pulsed Plasma

63

' "W W^ ' • ■ I'-■^^WIWWIiPPiiiPWWIpWBi MMMi

TABLE V. PROPULSION SYSTEM WEIGHT FOR 3000-POUND SATELLITE WITH A 10-YEAR LIFE NOT PERFORMING

NORTH-SOUTH STATION KEEPING

Total Total

Hiyli l.ou 1 'ropi-llai t llKTl Weight Power Weight

1 11 r u s l e r I h rust f r Weight Weight W/O Power | Penalty With Power |

"a" follol.l 107.713 i.o. 89 3 168.606 94. 6 263.206

DO follcia li^.-i'il 49.997 185.288 92.4 1 277.688

.\,114 Cat. Collni.l ISi.-Ut. 42. 46 194.897 92.4 287.297 1

N ,11. Cat. iiii. ij. i'. 152.4 56 49.425 201.861 132.0 333.861

N ,11 , Cat. Cd Ion 1-12.634 86.94 > 229. 580 198.0 427.580

11/' "a" 20i. 51 S 57. 55 i 261.071 2.2 263.271

DART DART 204.042 85. 397 288.039 - 288.039

DO DC) 240.824 50.62 1 291.448 - 291.448

M ll( Cat. N ,11 . Had ioisolope 251).217 4 3.493 293.710 - 293.710 |

N^llj C..I. DART 227.546 66.94 1 294.487 - 294.487

N ,11 Cat. N.ll, Ncvs istujct 298. 569 29.162 327.531 6.6 3 54.131

CIF./N.ll, "V1-! <■i(■, Plfiuim 337.498 41.189 378.687 - 378.687

■-'•L'^'U N.llîC 1 'lumiin 54 1.879 42.098 383.977 - 383.977

N.ll, Cat. N>ll| (;(; 1 TcMiun i 5 74.084 5 7.4 32 411.516 - 411.516

v^ll. Cat. N , 11 ('a 1 110. 1)20 5 7. 70 0 447. 780 - 447.780

N , 11 , Cat. ..,11 . I'.lu i rolyti( 480.295 42.078 522.373 19.8 542.173 " 1 "

mm* mm

TABLE VI. PROPULSION SYSTEM WEIGHT FOR 2000-POUND SATELLITE WITH A 5-YEAR LIFE PERFORMING


High Thruster

Low Thruster

Propellanl Weight

I n o 11 Weig it

Total Weight

W/O Power Power

Penalty^

Total Weighl

With Power

H20 Colloid 95.842 60. 788 156.630 94.6 251.230

DO Colloid 1 13.987 49. 07(1 163.057 92.4 255.4 57

N2H4 Cat. Colloid 125.266 40. 9S5 166.221 92.4 258.621

N2H4 Cat. Hg P. P. 125.266 47. 071 172.337 132. ) 304.3 37

N2H4 Cat. Cs Ion 107.001 84.902 191.90 3 198. 0 389.90 5 !

C1F5/N2H4 N,H4 GG Plenum 293.170 37.867 3 31.03 7 - 331.057 |

N0H, Cat. 4

DART 260.949 70.897 3 31.84 6 - 3 31.816

DART DART 246.6 54 86. 588 333.242 - 333.242

H20 H2U 275.462 59.459 3 34.921 2.2 337.121

N204/N2H4 N2H4 GG Plenum 300.932 39. 578 340.510 - 340.510

N2H4 Cat. N^ll , Radioisotope 300.454 45.910 346.364 - 346.364

DO DO 331.742 54.365 386.107 - 386.107

N2H4 Cat. N2H4 Resistojet 377.696 32. 846 410.542 6.6 417. 142

N2H4 Cat. N2M4 GG Plenum 4 12.496 59. 171 451.667 - 451.667

N2H4 Cat. N2H4 Cat. 512.553 42.285 554.838 | 5 54. 8 38

N2H4 Cat. N2H4 Electrolytic; 56 5. 962 45. 784 611.746 1 9. 8 631. 546

I

65

--■■-•- iîWifiutÛx.... i

TABLE VII. PROPULSION SYSTEM WEIGHT FOR 2000-POUND SATELLITE WITH A 5-YEAR LIFE NOT PERFORMING


1 High i Thni.Hlc r

Low J'h nislcr

Propellant Weight

i Total ! Total Inert Weight Power Weight

Weight I W/OPower| Penally With Power

ll/i ' Colloid

DO ' Colloid

N.ll Cat. Colloid

N^ll Cat. llg P. 1J.

DO DO

N'all4Cat.

C1FC/N,11, N.ll, GO Plenum

N Ü /N,llt N',11 C.C, Plenum

Nall4 Cat. ■ DART

N ,11 Cat. Cs Ion

Njll, Radioisotope

DART

Ul0

XJlj Cat.

N£ll( Cat.

DART

"z0

N,ll C.C Plenum

N , 11. Re s i s t o j c t

\,ll Cat. ; N31I j Cat.

NJI, Cat. X.ll Electrolytic

64.581

«3.015

94.475

94.475

114.763

123.835

1 38. 901

141.611

117. 022

91.535

101.587

130.146

160.191

192.057

299.488

573.085

58.712

47. 01 3

38.910

46. 71 1

44.932

36.944

31.038

31.698

57.894

83.931

74.993

55. 364

27.020

23.933

52.664

3 7.286

123,293

130.028

133. 385

141.186

159.695

160.779

169.939

173. 309

174.916

175.466

176.580

185. 510

187.211

215.990

332. 152

410.371

94.6

92.4

92.4

132,0

198.0

2.2

6,6

19. 8

217.893

222.428

225.785

273,186

159,695

160.779

169.939

173.309

174.916

373.466

176.580

187.780

187,211

222,590

332.152

430.171

!:

66

|^ -

TABLE VIII. POWER PENALTY

System

Electrolytic

Resistojet

Cs Ion

Colloid

Hg P. P.

H20

Penalty (pounds)

19, .8

6. ,6

198. Ü

92. 4

132. 0

2. 2

67

mtlmmmmlmmmmm ■

The ideal method of assessing a power penalty would be based on the

continuous power required by a thruster. However, thruster requirements

are given as, say, 100 watts. It is not stated if this is a continuous

requirement or one that is required just during the pulse itself. Also, no

information is given as to power required between pulses nor the amount

of ''warm-up" time required by a particular thruster. Therefore, power

penalties were assessed on the basis of the amount of power required by

a thruster limes the number of thrusters onboard. The weight of a solar

panel which would then provide this power was added to the propulsion

system weight. In this manner, the maximum power penalty has been

assessed upon the systems.

Figures Z9 through 32 present the total propulsion system weights

graphically. Figures 29 and 30 have no power penalties whereas Figures

31 and 32 include them. It should be noted that in each figure, the

systenis are arranged in order of increasing total weight for the mission

with north-south stationkeeping.

Several observations are to be made from the two figures without

power penally. The most important is the large weight saving obtained

when the small thruster is of the electric type and the mission requires

north-south stationkeeping. In the case of a 10-year, 3000-pound

satellite, this saving stands to be as much as 871 pounds between the

N-,H. catalytic-N ,11 gas generator (GG) plenum system and the H?0

electrolysis-colloid system, and as little as 395 pounds between the

all-HpO electrolysis system and the N fl . catalytic-Hg pulsed plasma

system. This weight jump between electric and chemical small thrusters

is not very large for a mission which does not require north-south station-

keeping, but there still are savings with electric propulsion. It is

interesting to note that in all four missions, the H_0 electrolysis-colloid

is the lightest propulsion system combination. For missions 10-3NS and

10-5 (i.e., 10 years, 3000 pounds, with and without north-south), the

all-water elei t roly sis system is the lightest of the chemical systems.

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For Mission 5-2NS, this system lacks only 4 pounds of being Lite lightest

of all chemical systems. There are some systems which possess properties

that translate themselves into total propulsion system weight handicaps for

some of the missions investigated.

For Mission 10-3NS, the three systems, allN^H. catalylic, N?ll

catalytic-N_H4 electrolytic and N?H. catalytic-N^H. GG plenum, have

approximately the same total propulsion system weight. However, this

trend disappears in Mission 5-2NS and 5-2, For these missions, the

N-H, catalytic-IvLH. GG plenum system is considerably lighter than the

other two combinations. This is a result of the large minimum impulse

bits required for N?H. catalytic or N?H. eletrolytic small thrusters.

Hence, for either of these two systems to be competitive for the less

strenuous missions, i. e. , without N-S stationkeeping, the minimum

impulse bit must be reduced.

A similar circumstance occurs with the electric thrusters and in

particular the N?H, catalytic-Cs ion system. However, in this rase, the

inert weight of the Cs ion thrusters and power-conditioning equipment is

the reason for the increased weight. For the less strenuous missions,

this inert weight begins to overshadow the propellanl savings obtained

with the large Isp of the Cs thruster when compared with the other

electric systems.

There are several different observations to be made if a power penalty

is included in the total system weight. These results are shown in

Figures 31 and 32. The jump in system weight is not as large when going

from electric to chemical. For Mission 10-3, there are several chemical

systems which weigh less than the Hg pulsed plasma or cesium ion systems.

For these two missions with the power penalty, the water electrolysis

system is the lightest chemical system.

73

„,„„ ^^,,wu,^MJx.,„.;WW,.WWW^ imarm

For Missions 5-2NS and 5-2, the effect of the power penalty on the Hg

pulsed plasma and cesium ion engines is quite evident from Figure 32.

For Mission 5-2NS, the weight of these two systems is comparable with

the chemical systems whereas in Mission 5-2, most of the all-chemical

systems weigh less than the electric systems. The water electrolysis-

colloid system is the lightest for Mission 5-2NS, but for Mission 5-2, the

all-DO system weighs less than any other one.

It is interesting to note that for all four missions investigated, the

INLH . catalytic-N^H. electrolytic system requires the most total impulse

to do the same mission. It also requires the most AV for mission

accomplishment. The N-^H, catalytic-N^H^ GG plenum will accomplish all

four missions while expending the least amount of total impulse.

D. SYSTEM COSTS

The cost of a propulsion system is an important and integral part of

a total system analysis. The development cost data for an advanced

propulsion system are very difficult to obtain. In addition, a significant

portion of the development cost is expended for flight qualification. The

cost of a system not previously developed for a similar application will

necessarily be higher than that of a system already flown. Therefore,

the costs shown here reflect the total system cost even though it may

have already been spent. Therefore, all values shown have a common

basis. It should be stressed that these are rough cost estimates and

should be considered as such. These estimates do have value in that

they provide a relative ranking of a system's cost. Table IX provides

the costs for the individual thruster concepts, while Table X gives the

cost for an ei. ire propulsion system combination.

74

^„-r-^rrrr^TW i..;^.rj.V^..T>n^"V-^rT-rT>-^->-n-T. | -,.. u....... vi ■ J û^.-,;.!.^! J.H^.J.,., .i...,,.!..^^ n^Ly-uîja.^ >N^*!H1M?f ^,-A'V^7y'!^.Jt',lIV,M'rrt.»Jl-V.>r<'S* ■'■ 7T ■.■J'TVI^f / W.f,'-?!.■'^T T-^rv-j *^r""- '-.--F.T'" '^

E. SYSTEM RELIABILITIES

Reliability data and combining techniques were used as suggested by

a recent JPL report (Reference 2). Data presented in that report were

derived from a review of previous reliability studies, reported component

reliabilities and failure rate values. However, no failure rate data were

included. Reliabilities for the noncyclic components were based on a

I-year mission duration, and the cyclic component reliabilities were

used based on 10,000 cycles. Reliabilities were not improved through

use of redundancy. A listing of component reliabilities used is provided

in Table XI. Reliabilities used for the valve-thruster combinations

are shown in Table XII.

TABLE IX. THRUSTER COSTS

Thruster

Large Thruster

N2H4 Cat.

DO

H20

N204/N2H4

C1F5/N2H4

DART

Cost

Million

2. 5

3. 0

5. 0

3. 6

4. 5

4. 7

Where Obtained

AFRPL best estimate

AFRPL best estimate

Marquardt Corp. contract and AFRPL best estimate

AFRPL best estimate

AFRPL best estimate

AFRPL best estimate and Reference 3

Small Thruster

N,II, Cat. 3.0 2 4

N,1I . GG Plenum 3. 5 c, 4

N.ll Electrolytic 1. 0

AFRPL best estimate

AFRPL best estimate

AFRPL best estimate

75

■■< ■ ■■■ ■■■:,-..JU^:. .,.■'■:■-

. . n.-.i.iûM'.vn-. *...—.,„,—„,v„..~ ■ vnu-i'T.--<iM.iip<i-i«.»f.><->iir~->i ' i J.II ijiii»ia\iiiii>i.iiiipi IIPHIIH»|JI.»I«!II.I iui"» .miiM)iiw"wJiipini«.»Mi .u .1 .ii^Kiinn«^nfqinnn««fimiimi|ui>nîn><q|a

TABLE IX. THRUSTER COSTS (Cont)

Thruster Cost Where Obtained

b. Small Thruslez*

NpH4 Resislojet

Cs

Colloid

Hg P. P.

DO

H20

2.6

N?I-I, Radioisotope 3.5

DART 4. 7

4.0

5. 5

4.85

3.5

5. b


Reference 3


AFAPL best estimate

SAMSO ADP

AFAPL best estimate

AFRPL best estimate

Marquardt Corp. contract and AFRPL best estimate

76

'■-^ nfifKn^m^mmmamn^m^^^^m^^mmmm^mt^mfm^m^^'i wrnmjmw n.n PH ,.!..■..—..-..■ . .——«

TABLE X. TOTAL PROPULSION SYSTEM COST

Large Thruster

I N2H4 Cat.

II N2H4 Cat.

III N2H4 Cat.

IV N2H4 Cat.

V N2H4 Cat.

VI N2H4 Cat.

VII N2H4 Cat.

VIII N2H4 Cat.

IX N2H4 Cat.

X DO

XI DO

XII H20

XIII H20

XIV N204/N,H4

XV C1F,/N H 5 2 4 XVI DART

Small Thruster

N2H4 Cat.

N2H4 GG Plenum

N2H4 Electrolytic

N H4 Resistojet

N-H. Radioisotope

DART

Cs

Colloid

Hg P.P.

DO

Colloid

V Colloid

N2H4 GG Plenum

N2H4 GG Plenum

DART

Total System Cost (millions)

3 . 0

3 . 5

3, . 5

5. , 1

6. 0

7. 2

6. 5

8. 0

7. 35

3. 5

8. 5

5. Ü

10. 5

7. 1

8. 0

4. 7

77

TABLE XI. PROPULSION SYSTEM COMPONENT RELIABILITIES

Noncyclic Components

Filter 0.9999

Propellant Tank 0. 9999

Plenum Tank 0.99988

Pressurization Gas Tank 0.99988

Line Heater 0. 99985

Pressure Transducer 0.99980

Bladder 0.99968

Fill Valve 0.99910

Lines and Manifolds 0.99850

Cyclic Components

Gas Generator 0, 994Z

Electrolysis Cell 0.9925

Pressure Switch 0.9925

Relief Valve 0. 9925

Regulator 0. 9900

Solenoid Valve 0.9871

Bipropellant Solenoid 0.9830

78

MMMHMMHMBMHMiMliMHI

TABLE XII. VALVE-THRUSTER RELIABILITY

Catalytic Monopropellant 0. 9958

11. Electrolysis 0. 9960

Bipropellant 0.9958

DART 0. 9976

Monopropellant Plenum 0. 9958

Electric Types 0. 9970

Radioisotope Types 0.9976

The method used to obtain the reliability of each feed system employed

the normal equation which is the product of all the component reliabilities

raised to a power equal to the number of times that particular component 1^

appears in the system. The JPL report then suggests the following

equations for the total system reliability.

1. The large thruster doing stationkeeping

4 Rc; ~ RF Rv T

^LT LT v "^LT

whe re:

RQ = total large thruster system reliability bLT

R = large thruster feed system reliability 11 LT

R -T „ = large thruster - valve combination reliability

2.. The large thruster not doing stationkeeping

R.S ^ RF RV2-T LT ^LT V LT

79

^^..1.. .,^■,^.U,J„^^„ ^,.1,^,1.., p,,^,..,,..,!,, |i,W,,,„^

3. The small thruster doing staticmkeeping

Rs = RF RV4_T

ST ST v ST

He = total small thruster system reliability ST

R,-, = small thruster feed system reliability ST

Rv T = small thruster valve combination reliability. ST

4, The small thruster not doing stationkeeping

Re = Rp RV2

T öST rST v -iST

Therefore, using these equations, a reliability for both the large

thruster system and the small thruster system was obtained for each

propulsion scheme. The propulsion systems can thus be ranked either

according to the reliability of the large thruster or the reliability of the

small thruster.

Table XIII shows the reliabilities for the large and small thruster

feed systems and the total large and small thruster system. It should

be pointed out that these numbers are based on the conceptual system

schematics in the previous section, and therefore, are "conceptual

reliabilities" only. However, they are useful from the standpoint of

obtaining a relative ranking of the various propulsion schemes.

From Table XIII, the reliability ranking (high to low) for the large

thruster feed systems may be summarized as in Table XIV.

80

l,^^m^tmHm^m^m*m*i^*mmmm*mmmmimmmtmmmmKmm

p!,.iw«.,i»iui.i| ^WIU'WIÎWWWJIV'WWKPI^^

TABLE XIII. SYSTEM RELIABII. riY

Small Thruster

l.a rui i hr usl cr Small Hi", iff

Large Thruster

1 cod System

1 olal System S\

1 rrd

stem 1 ..t.,|

System

N2n_1 Cat. N2II4 Cal. 0. 'ITS'I] (1. '^DK? 1). "-K'M i). 'i..: MI

N2H_1 Cat. Njll, Plenum 0. "-^S'M I). 'MZ'M) 1). • i i ,S ' l' 1 (i. ••.\\:.

N2n4 Cat. N2M Flectrolytic 0. ''r>S M o/isn.s? 0. ■ i - S u | n. ' ! V i ■ N.n. Cat. N.llj Resistoiel 0. 'l^S'M i). 'i-.O.^V II. ■'r y', 1 i). ••! ,-;i

N2II4 Cat. is', H . R adioisotope 0. 'V-xu| n, •■ 0,S7 1). ■•■-.VM (i. '-i ■.' ■

N2H4 Cat. DART 0. o.|i,s.| 0. 'r',,s, i) 1). | 1' ■..; ii. '"'iVi;-

N2H4 Cat. C H I o n 0. "-|i,--l O.'i '.h- ') I). I.|I,-,S 1). ■i-.',-

N2I14 Cat. Colloid 0. "■liir-1 0. u',s,,i) 11. 1.1,,-S.) i). ' • .!'.

N2H4 Cat. Ilg' Pulserl Plasma 0. ')AiÂ 0. 'HHi.O 0. i.l,,'-,.! n. ■'. .'.'.

DO DO 0. flSH'M 0. 'ii''7 5 0. l:",S')| 0. ■' I', I

DO Colloid o. 'i-ir^.i n. u-iaoo 1). '■1'-- ■; o.'' ■■ .:'

H-,0 "a0 0. 'H.'-O.'. 0. ''lT.d() II.'

1 '\ i). •17 1-

N204/N2H4 K ,11 . 1 'loimm L, 4 0. 'H'l-IO 0. 'UI-IOT t).' ^•'.78 (1. M'iill

c i r5 N2II Plenum 0. ')] q-io 0. OO-IOS 0. ' .'.l:7H 0. ''P'lll

DART DART 0. •i.-lS.ii o. .--.p, 1 0. ' RK'-I i). * i ■ * r'

HI.

■-—- --- .,-..-./ . .., fgn^gk

wmmmm.

TABLE XIV. LARGE THRUSTER FEED SYSTEM

RELIABILITY RANKING

Type

Monopropellant

Monopropellant

Water Electrolysis

Bipropellant

Propellant

Same as in small thruster

Different than in small thruster

Same as in small thruster

Fuel is same as small thruster propellant

A similar ranking for the small thruster feed system is in Table XV.

TABLE XV. SMALL THRUSTER FEED SYSTEM

RELIABILITY RANKING

Type

Monopropellant

Monopropellant

Plenum

Water Electrolysis

Plenum

Propellant

Same as in large thruster

Different than in large thruster

Same as in monopropellant large thruster

Same as in large thruster

Same as fuel in bipropellant large thruster

82

râ.j^^to^^j^ft^^ ^ ^^ tamaiMtMMtaaia . ■aî..,^;,;...-.-.....-.........,...,.■.:.. ;.-. ....,.■. ,.-. ^L...^..- ....^

ÜIHBIWBWtWBWW

The.se rankings serve to reiterate the obvious, that I he more

complicated or the more components in Ihe feed system, (he lower the

reliability ofthat feed system. In like manner, large and small Ihruster

total system reliability rankings may be obtained from Table Xlfl. For

the large thrustcr, this ranking is shown in Table XVf.

TABLE XVI. TOTAL LARGE THKUSTEK SYSTEM

RELIABILITY RANKING

Type Perform North-South Propellant

Nuclear Thermal No Same as in small Ihruster

Gatalytic No Same as in small thruster

Nuclear Thermal Yes Same as in small tli r uster

Catalytic Yes Same as in small thruster

Nuclear Thermal No Different than in small thruster

Catalytic / No Different than in small thruster

I-LO Electrolysis No Different than in small thruster

ILO Electrolysis Yes Same as in small thruster

Bipropellant Yes Fuel same as propellant in small thruster

The ranking for the small thruster system is in Table XVTI,

S3

mtrntmnamammm ■■' -■ ■■■ ■ ■

ii -■■ 'n,■■,■,■„■■ ^^ . ...-; aAteatttiiattttflaagffli

TABLE XVII. TOTAL SMALL THRUSTER SYSTEM

RELIABILITY RANKING

Type Perform North-South Propellant

Nuclear Thermal No Same as in large thruster

Nuclear Thermal Yes Same as in large thruster

Electric Yes Same as in large thruster

Catalytic Yes Same as in large thruster

Nuclear Thermal Yes Different than in large thruster

Electric Yes Different than in large thruster

Plenum No Same as in monopropellant large thruster

H20 No Same as in large thruster

Plenum No Same as fuel in bipropellant large thruster

I

These total system rankings once again point out the fact that the

simpler the conceptual diagram, the more reliable the system should be.

In addition, the nuclear thermal thruster is more reliable than the catalytic

systems. It must be pointed out, however, that there is probably a

difference in the reliability of the thrusters for the DO, DART and N-jH^

radioisotopes even though the value used for each is the same. The same

applies to the electric-type thrusters. Common values were used because

of the lack of data on these new thrusters. Therefore, care must be used

in extracting just a number from these tables.

84

F. PLUME EFFECTS, INTEGRATION AND HANDLING

There are several other areas where the different combinations may

be compared on a qualitative basis. One of these is plume effects. Two

areas within plume effects warrant discussion. These are the signature

of the plume and contamination of solar panels, and sensing devices by

the plume. Table XVIII shows the propellants, their signatures and

possible contaminations.

TABLE XVIII. PLUME EFFECTS

Propellant Signature Contamination

N2H4 Low Low for short term; has not had long term studies done

NH3 Low Low

Hz0 High High because of frozen water

N204/N2H4 High High

C1F5/N2H4 High High

Cs Low Thought: to be high but little data are available

Hg Low Same as Cs

Colloid Low Same as Cs

DO High High? No studies made

It must, however, be kept in mind that the contamination from any

one of these propellants may be lowered by appropriate positioning of

the thrusters or by careful tailoring of operating conditions. If the thrustc> rs ;i ro

pointed away from the solar panels and the sensors on the satellite, then

there may be essentially no contamination from any of them. However,

85

■ l m\\\ ^^m^mmmmmmmmmmm

from a plume effects standpoint, the all-hydrazine or ammonia systems

will produce the least plume effects whereas the bipropellants and the

water systems probably produce the worst effects. The cesium, mercury

and colloid may produce the worst contaminations of all but this isn't

really known because no real in-depth studies have been made using the

propellants. In addition, since two of the missions investigated here

are 10 years in length, it should be pointed out that no work has been

done on long-term plume effects of this duration.

Integration and handling are two other areas where comparisons are

required. The systems employing a radioisotope for a heat source have

unique problems in that they require special handling both on the ground

and within the spacecraft. Adequate shielding presents a problem because

of the excessive weight buildup of the containers. GIF- also has a ground-

handling problem as a result of its corrosive and toxic nature.

Certain systems have inherent problems or lack flexibility. The

cesium system is one of these because the entire feed system must be

kept warm (above 83. 3 F) to avoid the problem of frozen cesium in the

feed lines. If the cesium freezes, then the wicking process of feeding

the thruster will not work. Hence, there are inherent problems in the

feed system. Because of the electrolysis cell in the HÔ electrolysis

system, there is very little flexibility. The cell must be sized to

accomplish the task and cannot be split up in order to redistribute the

weight throughout the satellite. Also, the storage of gaseous hydrogen

and oxygen in the mixed condition presents a potentially explosive problem.

Care must be exercised in this area.

The electrical systems (Cs, Hg and colloid) are very complex

systems. They require large voltages and power. In addition, the

action of the accelerated particles can degrade and limit the life of some

of the engine parts.

86

i ;

■»'.^■llflMMT.I.MlM«!»»!!^ ■7i-f"ii"i'- ^jap^B^BpwpwiwffipiPffg wwaw^ "ii"n."^nw-'-' WPWJI

SEC7'ION IV

CONCLUSIONS AND RECOMMENDATIONS

Based upon the calculated data, .several conclnsinn.s may 1)0 made.

Conclusions as to system weight will bo based on data including power

penalty.

1. Eor the more strenuous missions, i.e., with north-south

slationkeeping, tlie systems using an electric small thrusler liavo a c-on-

siderable weight saving. This is particularly true if the electric concept

is a colloid system. The cesium and mercury pulsed plasma do not offer

as large a weight saving, and in one case, none at all (cesium on a

5-year satellite).

2. Eor missions requiring no north-south stationkeeping, the

electric systems offer no advantage from a weight standpoint. If the

mission life is for 5 years or less, a considerable weight disadvantage is

incurred as a result, of the power penalty required.

3. The water electrolytis or nuclearthermal systems appear to offer

some weight savings over the other "all-chemical" systems for the more

strenuous missions. Eor the less strenuous mission, these systems are

comparable in weight.

4. Although the all-hydrazine catalytic systems appear to offer

no weight savings, the effect of the hydrazine plume is less than all

other systems. Furthermore, because of the number of currently

operational catalytic systems, further development and flight qualification

are minimized.

87

mß . mmmmmmm******miimm*> ''■'■M*!^.»ul^JM.îi>.ii.i.uwiiiuj;i4iuipiui.ifpNi»pit]tij|ilM).MUjiuti.iiii

5. If the actual power penalty is anywhere near that estimated, the

electric systems offer no advantage for missions not requiring north- south stationkeeping.

6. The weight savings obtained for missions with no north-south

stationkeeping indicate that development of an all-DO system is warranted.

7. The nuclearthermal systems appear to have very good reliability

whereas the bipropellants and the water electrolysis system have worse reliabilities.

8. The reliabilities for solenoid valves are very poor and additional

development in this area is needed. Also, considerable reliability work and life testing must be done in the electric thruster

area.

9. The hydrazine electrolytic system suffers from a poor minimum

impulse bit of 0. 005 lbf-sec. If this could be reduced to 0. 004 lb,-sec,

then this concept may compare more favorably with the hydrazine catalytic systems.

10. The nuclearthermal systems have an integration and handling problem which must be solved.

The following recommendations are thus put forward.

1. Advanced development of the electric thrusters and, in particular, olloid thruster th

2. Life and reliability work on electric thrusters

3. Reduction of the power requirement of the electric thrust ers

88

piHWU miMiWIIilJl^^

4. Development of the water electrolysis thrusters

5. Reduction of minimum impulse bit for the hydrazine electrolytic

and the cesium ion thrusters

6. Development of DART and DO nuclearthermal thrusters

7. Improvement of valve reliability

89/90

APPENDIX A

SATELLITE PROPULSION SYSTEM

WEIGHT PROGRAM DESCRIPTION

The computer program can be divided into IS distinct sections. This

is shown on the overall logic diagram on the next page. The program

calculates certain data in each of these sections. The flow through these

sections is as shown in the diagram. Following the diagram is a short-

description of the calculations in each section.

91

^wate ■ -- MI^M

SI - READ INPUT DATA

I S2 - SIZE AND WEIGH SOLAR PANELS

Z S3 - SIZE AND WEIGH CENTER BODY

± S4 - CALCULATE MOMENTS OF INERTIA

S5-CALCULATE MAXIMUM PROJECTED AREA

I S6 - CALCULATE IMPULSE, AV AND PROPELLANT AMOUNT FOR 10 PROPULSION FUNCTIONS

I S7 - GET TOTAL IMPULSE AND AM

T S8-GET TOTAL AMOUNT OF EACH PROPELLANT REQUIRED

S9-SIZE AND WEIGH PROPELLANT STORAGE TANKS

1 S10 - IF REQUIRED, CALCULATE AMOUNT OF GN2 FOR PRESSURIZATION AND SIZE AND WEIGH A TANK FOR IT

I S11 - SUM UP ALL WEIGHTS TO GET TOTAL PROPULSION SYSTEM WEIGHT

S12 - PRINT OUT ALL RESULTS FOR SYSTEM

i 813 - IS THIS LAST SYSTEM TO BE CALCULATED?

NO

YES

S14 - ARRANGE SYSTEMS WITH "LIKE" PARAMETERS BY INCREASING TOTAL SYSTEM WEIGHT

Z S15 - PRINT SYSTEMS IN ORDER OF INCREASING WEIGHT

<g)

Figure 33. Logic Diagram for Computer Program

92

ti I I—ii——*-

SI - READ INPUT DATA

The data canis for a particular .system are read into the computer.

SZ - SIZE AND WEIGHT SOf-AK I'ANi-ILS

The initial gro.ss woi^lil of the satellite is used to determine the

onboard power from (he solar panels usini; e'.|ualion III-l in the main body

of this report. Tlic; panels arc then sized and weighed using the ideal

specific weight, percent life degradation, specific surface area and

height-to-length ratio.

53 - SIZE AND WEIGH CiENTERBODY

The centerbody weight is equal to I lie initial gross weight minus the

weight of the solar panels. The centerbody can then be sized using the

bulk density, centerbody shape code and the dimension ratios supplied as

input.

54 - CALGULATE MOMENTS OF INERTIA

Tiie moments of inertia foi* the spacecraft are calculated here.

Referencing the axis system as set up around the spacecraft in Figure 1

of the main body of this report, the equations for the moments of inertia 2

in slug-ft" are:

For spherical centerbody:

l M C.W I p L MS1 SI

MSF / 1) \ (A-l)

113

— .

/D\2 + i^SP xlp - 2_^SP psp\ Iyy = 2MCB /D\2 + 2MSP xL - ZMSP IXSP\Z (A-2)

Izz = ^CB ^f + ^SP (v2p + 4p} + ,A-3,

(f + ^ 2

2 Msp

where: ^-r^n = rnass of centerbody

M„p = mass of solar panel

D = centerbody diameter

Subscript CB = centerbody

SP = solar panel

For cylindrical centerbody;

I = MCB /D\2 + 2MSP YL + 2MSP /D.V (A-4) xx . __

1 yy

^8 (f) + ^-P Y- + ^ (f

M

~4g

2M

CB T/Df + L? + 2 MSP X2p - (A-5)

3g

94

zz M

CB 4g

2 M,

3

2 M

T2g" SlJ

^SP p , r^

4. I x •SI

(A-6)

For rectangular centerbody:

xx M

CB 12g

2 M

, CB

SP(!CB)

Y CBJ

Z M SP

H SP

(A-7)

yy

M CB (X

12g

2 M

CB

SP 8 w

z2 ^ ^CBi

2 M

"""3g" SP 4. (A-H)

zz M

CB 12g CB

2 jVl SP YCB

CBJ

Y

2lV1SF (YZ

SP

XSF (A-9)

S5 C A L C U LA T E M AX 1 M Li M J > R O J E C T ED AR F.- A

This area is required in the calculation of solar pressure lorrtM -

tions and is the area seen when looking along ihe '/. axis in Figure I,

including solar panel and centerbody projected area.

9S

S6 - CALCULATE IMPULSE, AV AND PROPELLANT AMOUNT FOR TEN MISSION FUNCTIONS

The ten mission functions are calculated in the following order

where it has been assumed that the satellite is repositioned halfway-

through its lifetime. It is possible, however, to eliminate any number of

these functions. In such a case, the computer automatically sets the

impulse, AV and propellant required to zero.

1. Despin

2. Tipoff

3. Injection

4. One-half of the total E-W stationkeeping

5. One-half of the total N-S stationkeeping

6. One-half of the total attitude maintenance, i. e. , solar pressure, limit cycle and contingency.

7. Repositioning

8. One-half of the total E-W stationkeeping

9. One-half of the total N-S stationkeeping

10. One-half of the total attitude maintenance

11. Stationkeeping contingency

The method for calculating the impulse, AV and amount of pro-

pellant for each propulsion function is as follows:

1. DESPIN

The impulse is calculated by dividing the satellite initial

angular momentum by half the maximum distance between thrusters in

the x-y plane (Figure 1). If the centerbody is a sphere or a cylinder,

then this maximum, distance is the diameter. If a rectangle, then use the

largest dimension (y or z) of the centerbody.

The amount of propellant equals this impulse divided by the

Isp for this function, and the AV required is obtained from the normal

equation.

96

AV u Jsp In (M /M ) 0 o e (A-10)

where: M - mays of satellite at beuinninti of maneuver o

M •- mass of satellite al end of maneuver e

£■ XiifQFF RATE

The impulse for tipoff is read into the computer as input data.

The amount of propellant and ^V are then t:alc:ulaled as for Despin.

5. INJECTIQN ERROR

The AV for this is read in as input data. The amount of pro-

peJlanl is determined through the use of equation A-10 and the impulse

by multiplying the amount of propellant by the Isp.

-1. E AST -W EST ST AT 10 N K EE iJ1NG I From Table I, tbe AV is equal to

AV = 7. 1 5 t m (A- 1 1

where t is the satellite life in years. The impulse and propellant are m

then calculated as an Injection Error.

5. NüRTIJ-SOUTU STATIONKEEldNG

From lable II, the AV is equal to

AV I SO t 111

Impulse and propellant ctre as in cast-west st al ionkeeping.

97

•mT.i^wrr -■.■c„,,,■r,.^,^lm.^^1^^r,^..M.-MF^W'i.^^^,yi■^m'yl''M^ ^v^^KU.i^w W!WWlW«W?BpiW!W?WW»'

1

6. STATIONKEEPING CONTINGENCY

The amount of impulse is equal to 3 percent of the sum of the

east-west impulse and the north-south impulse. AV and propellant are

then calculated as in Tipoff Rate.

7. SOLAR PRESSURE

The impulse is calculated from equation III-2 in the form

I = 5.91 t (0. 35) A. t ml s

2 where A = maximum projected area, ft . The AV and propellant are

then calculated as in Tipoff Rate.

8. LIMIT CYCLE

A value is calculated from both equation III-3 and equation

III-4. The impulse required for limit cycling is then the larger of the

two numbers. AV and propellant are then calculated as in Tipoff Rate.

9. ATTITUDE MAINTENANCE CONTINGENCY

The impulse is calculated as per equation III-5 with AV and

propellant as per Tipoff Rate.

1 0. REPOSITIONING

The AV is determined from equation III-7. Impulse and pro-

pellant are then calculated as per Injection Error.

S-7 - GET TOTAL IMPULSE AND AV

The total impulse and AV are obtained by summing the impulse and

AV from each of the ten propulsion functions.

98

^.■WlWLIMJ!l>'JI.Û'!ll.'M|li|.l'l'IP..'«llJlffWK-i|»W.'l'JI*llMII!.l||-!M lüRIPilüPPP i. mmg^mmmmmm m * mmmmmmmm *m ^^^^^^T.

S-8 GET TOTAL AMOUNT OF EACH PRO PE LEANT REQUIRED

In this section, the amount of each propellant used for l lie ten pro-

pulsion functions is summed. This set-lion has the capability of handling

a bipropellant large thruster. By using the mixture ratio on this, the two

propellants can be split out to give correct propellant sums. An ullage of

1. 5 percent is then added to each propellant sum.

S-9 - SIZE AND WEIGHT, PROPELLANT STORAGE TANKS

Spherical tanks are designed for proprilant storage. They arc

sized by getting the total volume required from knowing the total pro-

pellant and the propellant density. The tanks are then weighed using the

equation

. , f 2 77- CBRJ n weight = Q-g p cr„

A- 12)

where: <5

R

P

operaling pressure, psi

tanl^ radius, inches

tank material density, lb/in"

yield stress of lank material, psi

. 3

A safety factor of 1.25 is used in this calculation. From this weight, the

thickness of the tank is determined. This thickness is then compared with

that coming from the normal stress equation

q]R (A-13)

The largest thickness is then selected and compared with the miniimini

average workable thickness (input data) for that particular metal. The

largest of the three thicknesses is selected and the lank rewcighed usim

this thickness. Then 15 percent of this weight is added to account for

99

.<i,„"""""..ip....,., " '' '•"""•■ '•'•"""* ■ ■ '■> immm 1.1 iPPlHnpiiPRpimiiiiiiRp^

fittings, flanges and attachment points. The program has the ability to

not design a lank for any particular propellant if so desired.

S-10 - PRESSURIZATION SYSTEM

The program can design a pressurization tank, weigh it and weigh

the gaseous nitrogen placed in it if so desired. It is also set up to

pressurize only some of the propellant tanks. To simplify the calculations,

an isothermal expansion of the GISL was assumed.

S- 11 - SUM UP ALL WEIGHTS TO GET A TOTAL PROPULSION SYSTEM WEIGHT

All weights are summed.

S-I2 - PRINT OUT ALL RESULTS FOR SYSTEM

The results for the particular system just calculated are printed out.

S-I3 - IS THIS LAST SYSTEM TO BE CALCULATED?

A check is made to see if there are more systems to be calculated.

Up to 99 systems may be calculated with one computer run.

S-I4 - ARRANGE SYSTEMS WITH "LIKE" PARAMETERS BY INCREASING TOTAL SYSTEM WEIGHT

Systems with common parameters (i.e., doing the same propulsion

functions, having same initial gross weight and satellite life, etc.) are

ranked according to increasing total propulsion system weight. If the

computer received no inert weights (pipes, valves, thrusters, etc., as

input data) for a particular system, then this system's total weight will

appear as zero in the listing.

S- I 5 - PRINT SYSTEMS IN ORDER OF INCREASING WEIGHT

This increasing total weight listing is printed out.

100

^t^1-■4■Tl*?l^'j^■^■^■.-•■■J■^• ^■,-Lfa.i-;<Aixs-iuiii\-.'.■*;<•..T»~M>.-..'»atfaJm^JMamWatt tBntMtaaatf■■*.■ .1 —^..-.^.• -aau&iä>- .t>.-f..-■.■■ ^.«^.■'.J*,.I■ -.,- ...-^ -:.i--.-i-..-^/„^^ ihvvv'miJWI■ ~"jfltotfim —. .■..■-.■>'..■.-■ -^ .^. f,..^--^.^.^.. ....:..■.._.^.:;^_-,.^..^*a

m™*^mmmm*mmmmmmmmmmmmmmmmmmi^^^*^*******~**>*'^^*m*m mm^

APPENDIX B

SATELLITE PROPULSION SYSTEM WEIGHT

PROGRAM USER MANUAL

This section describes the procedure i'or writing the input required to

use the computer to weigh the total propulsion system to accomplish the

proposed post-lQ?1! mission model. Twelve different types of input data

cards are required to manipulate the program. Cards 2 through 12

describe any particular high-low thruster combination desired. Card 1

tells the computer how many such cases or combinations will be investi-

gated. Up to 99 different combinations may be calculated during one

computer run. However, a complete set of data cards (cards 2 through

12) must be furnished for each case. In the event of a special study

involving relatively few changes on repeated cases, the original data

deck for that case must be reproduced and only those cards containing

changes must be repunched and inserted.

The input variables required by the program are defined on the

following pages. This is followed by a listing of card formats and input

instructions. Also included is a list of suggested values for some of the

variables required by the program.

1U1

■•.,...i_y'..iiv>'";^',: :.j".--.^. -:-.-. ., —i,--^,'.w.-i ^.(.V-'-^L- .•■->•■'-J.i ••■■*■ ^ ■,■-■• ■■ --..-..-; i—v-i-■'-;■-

TABLE XIX. COMPUTER INPUT VARIABLES

C(i)

DBHAID

DEN(i)

IDWHB

IDWPET(i)

IDWSEP(i)

IDWWP

INOP

INOSTR

IS(i)

ISCBC

LL(i)

MM

OPPRE

The Chemical used in a thruster

The Dead Band Half Angle In Degrees 3

The DENsity of chemical C(i) in lbs/ft

Means Do We Have a Bipropellant large thruster. Depending upon the value of IDWHB, the computer will decide whether there is a biprop or not

Is a code that identifies which chemicals are expelled under pressure and which ones are not (Do We Pressurize Each Tank)

Is a code which identifies which chemicals are stored in a tank and which ones are not (Do We Store Each Propellant in a tank)

This is a code which tells the computer Do We Want a Pressurization system

The Number Of different Propellant, C(i), in the system

The Number Of jvystems To Be Run or calculated. A system refers to a particular large-small thruster combination

A code which tells the computer which thruster (large or small) is used for mission function i

A code which tells the computer the geometry of the centerbody. (Satellite Center Body Code)

The number of systems which perform common mission functions

A code which tells the computer whether to add inert weights or not

The initial storage Pressurant PREssure in psi

102

TABLE XIX. COMPUTER INPUT VARIABLES (Continued)

PTDEN

PTSIG

PWMIN

SCBDEN

SCBLOD

SCBZTX

SDV(3)

SIAM

SIG(i)

SISP(i)

SIT(2)

SLIFE

SPHTL

SPISW

SPPCLD

SPSSA

SREPRA

The density of the tank material used for storing a pressurant in lbs/in? (Pressurant Tank DENsity)

The yield stress of the tank material used for storing a pressurant in psi (Pressurant Tank SIGma)

The minimum pulse width of the small thrust or in seroiuls (Pulse Width MINimum)

The S_atellite Renter ßody bulk DENsity in lbs/ft

The_Satellite Center Body Length Over Diameter if il is a cylinder and the Z/Y ratio if it is a rectangle

The S_atellite Center Body Z To X ratio if it is a rectangle

The Satellite Delta Velocity required for injection error in ft/sec

The Satellite Initial Angular Momentum in ft-lb-sec

The yield stress (SIGma) for the storage tank material for chemical C(i) in psi

Are the System ISP's for a given propellant system for 10 steps or functions required in the mission in sec

The total impulse required for tipoff rate in lb .-sec (Sâtellite Impulse X0tal)

The Satellite LIFE in years

The Solar Panel Height To Length ratio

The Solar Panel kleal Specific Weight in lbs/k\v

The Solar Panel PerCent Life Degradation

SJola1" Panel _Specific Surface Area in ft /kw

The Satellite REPositioning RAle in degrees/day

he

f

103

*•*—-■—■ umaMttttiStfaiiaaiu m ^ i tuatouatä* muäaâätBg -.-vi—..- ■..., ^■L...:..^.^,.^,,:..,..,^.^...,.....-..,, ,

■^> ■»•i.-.t ..!.i,;n-..|.m|.|iM1j.u..L.wii.nMi.n ..I.I .I|.I..JW> ,..■■'! ■■■in i. .i.ii i-, ^.>—.n"J,in;nan"»;u»"<".»»"V<M «»'■■■«"» ',■'■">»' >-T^r~-^TJr^-^B~yTiT^y!r^»-ntTT^lTT"

TABLE XIX. COMPUTER INPUT VARIABLES (Continued)

SUB(i) The propellant (SUBstance) or propellant combination used in a thruster

SWOT Is the Satellite \VeiGhT (initial gross weight) in pounds

TAMA(i) The name of the storage TAnk MATerial for chemical C(i)

TAPR(i) Is the storage TAnk PR es sure for chemical C(i) in psi

TIIEDMI The minimum achievable angular rate of the satellite for limit cycling in degrees/sec (THEta Dot Minimum)

THRMIN The small THRuster MINimum thrust in lbsf

TMDEN(i) The storage Tank Material DENsity for chemical C(i) in lbs/in.3

TMMT(i) The minimum workable thickness for storage tank material TAMA(i) in inches (Tank MinimuM Thickness)

WGTI(i) The inert weight of the propulsion system - includes, pipes, valves and thrusters in pounds (WOT Inert

104

Card No.

10

1

INPUT DATA CARD FORMATS

4 pr

(

Conto] and Setup (3112)

Large and Small Thruster Names 3 (3A6);

Propellant Name and Propellant Density 3 (2A6), 3F10. 3j

Propulsion Function and System Specification 101?., 9X, II, 9X, II, 9X, II, 10X, F6.3)

5 Propulsion Function Isp's (10F8. 3)

6 Ceneral Satellite Specifications (5F10.4, 9X, II, 2F10.4)

7 Solar Panel Specifications (4F10.4)

8 Thruster Specifications (FIG. 4, 3 F10.6, 2F10.4)

9 Propellant Storage Tank Materials 3 (2A6);

Storage Tank Material Properties (3F10.3, JF10.2, 3F6.4)

Pressurant Storage Tank Material Properties and Operating Pressures (2F10.3, F10.2, 3F10.3, IX, II, IX, 11, IX, II, 5X, II, IX, II, IX, 11) .

12 Thruster, Piping and Plumbing Weights (4F10. 3, IX, II)

105

. A-^-m.i:.^. ......... .,„., .... .-:......._....,.:.,..,,_..._. ^jMjauuJ^MlltUaiaaaiL^läläi^llÜBIaUäi '<tUa**miiM*lLULl ■ ■• • ....,■■.-.. -.■:■. J

'J^wwâ^rmfva B*W^...»WWVWVMWJWMIU^^^^^^ WNsm

CARD NO. 1 (3112)

CONTROL AND SETUP

Variable Columns Remarks

INOSTR 1-2 Place the total number of distinct system cases to be investigated in these two columns. The number must be right oriented. (Limited to 99)

LL (1) 3-4 Place the number of system cases that perform common mission functions here. The number must be right oriented.v

LL (2) 5-6 Place the number of systems cases in the second group which perform like mission functions here. The number must be right oriented.''"

LL (30) 61-62

At the end of the computer output, the computer ranks all system cases with common parameters (i.e., initial gross weight, life, etc.) or mission functions (i.e., north-south stationkeeping, repositioning, etc) in order of ascending total propulsion system weight. These variables (LL (I) ) tell the computer how many system cases are in each common grouping. If LL(1) -■ 5 and LI.(2) - 8, then the first 5 cases in the input cards will be ranked together and the next 8 input cases will be ranked together. There ran he up to 30 such groupings.

106

jIM^nmmifuuvNpip!^

CARD NO. 2 [3 (3A6)J

LARGE AND SMALL THRUST ER NAMES

This card can be filled out four different ways, Mali h Ihe .system lo he

investigated with either CASE A, R, C. or I) and inpnl ;u < «rdin^ly.

Variable

CASE A

SUB(l)

SUB(2)

CASE B

SUB(I)

SUB(2)

CASE C

SUB(l)

SUB(2)

CASE D

SUB(l)

Columns Remarks

The large thruster is a rnonopropellan.. ...v. .J...aii iniiisicr is monopropellant. The chemjcaj used in the lai-ê Ihruster is no the same as tliat used in tlie small I lirunter.

I, 'Mu; sniall l lirusl oi- is a l

1-18

19-36

Place a desc riplor for llio laruo Ihruslor hero. Example: CAT AI, Vl'IC' X^ll-l

l-'lace a desiriptor for llic small (liruster hero. Example: DART (KWi)

The large thruster is a monopropellant. The small Ihruslor is a monopropellant. The chernical used in each Ihruster is I he same.

1-1!

19-36

Place a descriptor for the large Ihrusier here. Example: CA'I AI.YTIC Nil M

SUB(3)

Place a descriptor for Ihe small Ihruster here. Example: EI.K( T ROl, V I fr Nil M

The large thruster is a bipropellant. The small ihruslor is a mmn.- propellant. The chemical in Ihe monopropellanl IS the same as ihe fuel in the bipropellant.

19-36

•18 Place a descriptor lor- Ihe bipropellanl large ihruster here. Kxample; \'\(i/ HYDRA/INE

I'lace a desiri])tor fur llu; miaioprojjell.nil small Ihruster here. Example: f A I A I, V I H Nil 14

The large thruster [s a bipropellanl. The small ihrusier is a monopropellant. The chemical in Ihe monopropellanl is NOT Ihe same as the fuel fn the Tiipropellanl.

1-18

19-36

37-54

I'lace a descriptor for Ihe bipropellanl large Ihrusier here. Example: N'l O/ IIYDRA/INE

NOT I'SED

Place a desi rijilor lor Ihe moiio])i'opi-llar.l small Ihrusier here. Kxample: DAR 1 (Nil 5)

107

pp. ..,,,! i...,.^^^,.^.. ,.., »TÎWWTB» "" ... ..n. ..MI»?, ll.liiillllllUIIWI lll|>|,|.|,III.IJI|l|W,m. II I l.l<!ll|IJ|ll „ I ll,W![|f^|^pB^îî|^^W^WîWipiW^W^^^PWW^^^^^

CARD NO. 3 [3 (2A6), 3F10. 3]

PROPELLANT NAME AND PROPELLANT DENSITY

li i.- ii- ! 1. . ..riuiL in ih,' in^lnu linns rn.iU hinu I In' . .1.-1' . hosi'n lor c ' A H 11 .'..

\ i!i.,!.],■

i .11

HI \. II

:ii:\.^l

\-.i-' II

< 11 I -I,'

1 (- '■■

i V - 1.

Hr. un.ll LIM .lli.in

i Ol. I t

I HI. -!

Uonia rk.s

I'l.ui' tin* » licinu .il synilml herti for llu* rhemical use«! in llu* larL;o thrustor.

I'lait." Itu; thfiiiital synibul hero for ihc chemical USLMI in the small lliruslor.

NO I VSF.D

The density of chemical Cll) in lli/fl

Ihc (iensily of Chemical ( (-1 ill Ib/fl

I'Uui' the themical symbol here for Ihe chemical iiscfl in l»o[ti the larue and small thrnster.

NOT rsi-:n

Ihc density of i homical CM) is lb/ft .• '

i)i-:N( I i

! IKNI.:I

\- !•■ : i

coi. i i

coi, it

col.

I Ol.

I'lace the chemical symbol here for Ihe oxidi/er in tue lariê lliruslor bipropellant.

IT.oc the chemical symbol here for Ihe fuel in the lariie Ihruster bipropellant

NOI CSKIl

Ihe density of the biprop oxidizer in lb/ft

Ihe density of the biprop fuel in lli/fl

I'lace the chemical symbed here for the biprop oxidi/er.

Place the chemical symbol here for Ihe biprop fiel.

Pia. c the . hemic al symbol here for Ihe . hen.;. . in the monoprop small thrnster.

Ihe density of the biprop oxidizer in lb/ft

Ihe density of the biprop fuel in Ib/fl

i he density of the monoprop . bemical in lb'ft''

inial point i- c-pli. illy itiêrted in input.

108

.>.l-J..kiK».W.».. '. -. ' -.-— i-^JJ -i- —'-i- ■-.■i '^■' ^...-^. u.*^^! .^^^.^.-l. ■•.-.^■iV^-J.J^.Ja-L..^J-^^ltm.'. —Jl ttüjMiiMiafltiiflcagtatotadAteaMMaiaQMaatwt ^...îu,.:.:, aanu ^■.-■:.- | .-.-..-^-. - .■;-..'.\J..^...^..WI.J.I......V^.V..

T-MI'"?-.!-!^/ .^ti-w'-..- y^py^r-rrwvvwrr-'**. ?n!n?7Tyr^^"AwywHi'B'.*w^:*f«ff'g*??^M.'»i'^' gjjiiii,MwiiM»1Hiff^-ju,^/rw^^^wy,''yy^^^'^'F^^*^^

CARD NO. 4 (1012, 9X, II, 9X, II, 9X, II, iOX. F6. 3)

PROPULSION FUNCTION AND SYSTEM SPECIFICA'ITON

Decimal Vari able Columns Location

None

Remarks

IS(1) 1-2 Despin cude;:

IS(2) 3-4 None Tipcjff rate code-

tS(3) 5-6 None In Joel; ion error code

IS (4) 7-8 None E-W stationkeopinti

IS (5)

IS(6)

IS (7)

tS(8)

IS (9)

IS(10)

IDWHB

DVVWI"

INC^P

VVGTR

9-10

11-12

13-14

15-16

17-18

19-20

21-29

30

31-39

40

41-49

SO

SI-CO

6 1 -()(>

None

None

None

None

None

None

None

N o n e

N o n e

COL ()3

code:':

N-S stationkoeping code:;:

Stationkeeping contingency code-'

Solar pressure code-:

Inmit cycle code:'':

Attitude maintenance contingency code:':

Repositioning code:;:

NOT USED

Dipropellant largt; thruster code:':';

NOT l'SED

Pr csbii rization curie- -■:'

NOT USED

Number oJ" chemicals code- - -:--

NOT USED

Ihn mixture ration for the i)iprüpellant large thruster. must be written as the ratio C(ll/Cf2)

109

■^■^■.,.,...■,.■.. ■,......~-.^w,....l| ■ ■■: ■ '■<■ ^.^™'***.:***.™.*mm™M*..„ .***,.•,.. îrt.i.j-^-^-

■^T---^^TT^.-.,B-rT-^... . —ny^p..^^ TV-J,. -,-. n ..v -i», !■-■ ^rvff-.-•' vv^—>". i -1- ,i. ■#/ v * .-'■■* ^, fn;.'---^-Tr-"v'T^^^^''r"''T--^""T^- ifflBTB! »ffWBWffgg ^*-^y tt',-'t"f'v^^-1-' .^1

CARD 4 (cont)

::: These ten codes tell the computer which thruster (Large or small) is used for each of the ten mission functions. If a function requires the large thruster, then [S(i) = 1, If a function requires the small thruster, then IS(i) = Z or 3 depending on whether SUB(2) or SUB(3) was used on Card 2, If it is desired to eliminate one of the ten functions from the system, then IS(i) = 0 for that particular function. More than one propulsion function can be eliminated at once. All numbers must be right oriented in the fields.

0 if the large thruster is a monopropellant

1 if the large thruster is a bipropellant

= 0 if all of the system is a blowdown

= I if any part or all of the system is to be pressurized from a gas bottle

The total number of chemicals placed on Card 3

110

^^mnmm^

CARD NO. 5 (10F8. 3)

PROPULSION FUNCTION Isp's

Variable Columns Decimal Location

SISP(l) 1-8 COL 5

SISP(2) 9-16 COL 13

SISP(3) 17-24 COL 21

SISP(4) 25-32 COL 29

SISP(5) 33-40 COL 37

SISP(6) 41-48 COL 45

SISP(7) 49-56 COL 53

SISP(8) 57-64 COL 61

SISP(9) 65-72 COL 69

SISP(IO) 73-80 COL 77

R em a r k s

Despin Esp

Tipoff rate Isp

Injection eiTor I.sp

E-W stationkeeping Isp

N-S stationkeeping Isp

Stationkeeping contingenc y Isp

Solar pressure Isp

Limit cycle Isp

Attitude maintenance contingency Isp

Repositioning I.sp

f

These are the 10 Isp values in seconds which are obtainable from l he

particular propellant(s) and the duty cycle for the 10 mission funclions.

1 11

*-"—-: ■—^—---■nir ...,.■,. ■-1._^,... ..-. ■- ama ■■-■- :■■■„..^._....., .■■.,.■ ^1?-^.,^..^-^^^^.!:.........■■■■..-...,.^L....'....^^-.i.^-îi^,,;.' -'-■■ - ■■- ■■ -^ -.■..:-. ., ....■■■ ...-■ i . . ..- . i-irihiiil

•■"■ " wwmmmm w rnrnrnm***1 ■'■ "m pnmnipillliilliPiiPiiqniiff^^

CARD NO. 6 (5F10.4, 9X, II, 2F10.4)

General Satellite Specifications

Variable Columns

SWGT

SL1FE

ID

SCBDEN 11-^0

SCBL0D 2 1-30

SCM/.TX 3 1-40

41-50

SREFRA 71-80

Decimal Location

Col 6

Col 16

Col 26

Col 36

Col 46

51-59

1SCBC 60 None

SI AM 61-70 Col 66

Col 76

Remarks

The initial gross weight of the satellite in pounds.

The bulk density of the satellite centerbody is lb/ft3

If the centerbody of the satellite is a cylinder, then this is the L/D ratio. If the centerbody is a rectangle, then this ratio is the ratio of the dimensions of the end of the rectangle which faces the earth.

This is used only if the centerbody is a rectangle. It is the ratio of one side of the end of the rectangle to the length or height of the rectangle.

The life of the satellite in years.

Not used.

Centerbody shape code

The satellite initial angular momentum in FT - LB - SEC

The satellite repositioning rate in degrees per day.

1 if centerbody is a rectangle Z if rente f body is a sphere ; if e enter body is a cylinder

112

■ ■■■ iÂtf.rLaj- £i-J*tUi -■ ■■ --' ■-■■■- ■ ■-■ ■■-■-.--■ ... e....„...,...,..- ■ „. . - -.... . .. .......-.-, ■■ij.î.i,■■'..■■.■..■.J.. - ■ . .- - ' ■ . .l../.,w.,.l-....^!....>-1..v......: ...^.■.^.;-.;..-.:^;-Jt....^^.'-.-..^..-.. .:i——.*■■■ ■.■.:.;.^.^JJlj4..J.,lüriiWll.'MifltT<AV7a

»■Ill I UII-W'MWMWIWJ^'inililJÛJgyHljMil^ HWWUBPômBHiwww '■'" '"■'■'-' rfnuttmim

CARD NO. 7 (4F10.4)

Solaz* Panel Specifications

Vaz-iable Columns Decimal Location Remarks

SP1SW Col 6 The solar panel ideal specific weight in LB/KW

SPPCLD 1 i-ZO Col 16 The solar panel percent life degradation at the end of the satellite life.

SPSSA 21-30 Col 26 The solar panel specific surface area in FT^/KW,

SPIITL 3 1-40 Col 36 The ratio of the height of the solar panels (the side adjacent to the centerbody) to the length of the solar panel (the side perpendicular to the centerbody)

113

mtät*^,^.±t..*Â*Livivä±*UU... .,.-..■!,..*.., ^..^„.^^.l.^^.L.^^L.^.^^*^^^,.^^^^^^.... ....... ■...,■„ ..:■.. ■ .1. ■..-. ...■.:,■..,- .. .... , ... . , . J..^..... .. ^_^ h n i läiäii

CARD NO. 8^10.4, 3F10.6. 2F10.4)

Thruster Specifications

Variable Columns Decimal Location Remarks

THRMIN 1-10 Col 6 The minimum thrust of the thruster used for limit cycling in LBp.

PWMIN 11-20 Col 14 The minimum pulse width obtainable with the valving for the thruster used in limit cycling in seconds.

DBIIAID 21-30 Col 24 The dead band half-angle for limit cycling in degrees.

THEDM1 3 1-40 Col 34 The minimum achievable average angular rate in limit cycling in degrees per second.

SIT (2) 11-50 Col 46 The total impulse in LB-SEC required for tipoff rate.

SDV (3) 51-60 Col 56 The delta V in FT/SEC required for injection error.

114

CARD NO. 9J 3 jiAG)

Propellant Storage Tank Materials

Variable Columns

TAMA (1) 1-12

Remarks

A descriptor for the lank material for storing C (I). Example: TITANIUM

TAMA (2) 13-24 A descriptor for I ho tank material for storing C (Z;.

TAMA (3) 25-36 A descriptor for the lank material for storing C (3).

115

CARD NO. 10(3F10.3> 3F10.2, 3F6. 4)

Storage Tank Matex'ial Properties

Variable Columns Decimal Location Remarks

The density of tank niaterial TAMA (I) in LB/JN3

The density of tank material TAMA (2) in LB/IN3

The density of tank material TAMA (3) in LB/IN3

The yield stress for tank material TAMA (I) in PSI

The yield stress for tank material TAMA (2) in PSI

The yield stress for tank material TAMA (3) in PSI

The minimum workable thickness for tank niaterial TAMA (I) in INCHES

TMiVIT (2) 67-72 Col 68 The minimum workable thickness for tank material TAMA (2) in INCHES

TiVlMT (3) 73-78 Col 74 The minimum workable thickness for tank material TAMA (3) in INCHES

TMDEN (1) I-10 Col 7

TMDEN (2) 1 1-20 Col 17

TMDEN (3) 2 1-30 Col 27

SIC. (1) 3 1-40 Col 38

SIC (2) 41-50 Col 48

SIC (3) 51-60 Col 58

TMMT (1) 61-66 Col 62

116

....rJ,tV;i.... -.-ii i^d». ..-i- - i. J ■■-'-' utait* . .-.. 11 OiVfr.,, ■,-■.<'■-■• ■'-i.jiiBi

CARD NO. 11 I2F10.3, F10.2, 3F10.3, 3(lx, 11), 4x, 3(lx) II

Pressurant Storage Tank Material Properties and Operating Presaurcs

Variable

01'PRE

Colurnna Decimal Location

1-10 Col 7

RxMnarks

rJ'he initial storage jaressure of the pi-c.S' surant (N^) in HSIA.

PTDEN 1 1-20 Col 17 The density of 6AL-4V Titanium used as t he tank material for the pressurant - This has a value of 0. 1 6 1 LB/1N3

JTS1G 21-30 Cnl 28 The yield stress of the 6AL-4V Titanium used as the lank material for the pressurant - This has a value of 176, 000 PS1A

TAPR (1 31-40 Col 3 7 The storage pressure of chemical C (1 ) in PS1A

TAPR (2; 41-50 Col 47 The storage pressure of chemical C (2) in PSIA

TAPR (3, 51-60 Col 57 The storage pressure of chemical C (3) in PSIA

Not usei

IDWPET (I) 62 None 1 'res su rizat ion code fo r chemical C: f I )"

63 Not used

IDWPET (2) 6-1 None Pressunzation code for Chemical il (2 r

6 5

117

Nol used

...■.,-^.,4l,.....^ .:-^..„ ^.u^l,^....,.^.^..^^^..*.~i..~.^ —^

Variable

1DWPET (3;

1DWSEP

IDWSEP (2!

IDWSEP (3)

Columns Decimal Location

66 None

67-71

72

73

7-1

75

76

None

None

None

Remarks

Pi-essurization code for chemical C (3):::

Not used

Storage tank code for Chemical C (I)**

Not used

Storage tank code for chemical C (2):'::':

Not used

Storage tank code for chemical C (3):'::':

0 if chemical C(i) is pressurized from gas bottle 1 if chemical C(i) is blown down

0 if chcinical C(i) is stored in a lank 1 if chemical C(i) is not stored in a tank

118

ra*v-.y*WA?<ja^-*,i**s«a*^ '--^k^^nÛiitmur.^luuu.^r.-âWl îr.tnlii-titiTiYT^- lii.-iii ^J..^^..^,.l>^._- .^

CARD NO. 12 (4F10. 3, IX, II)

Thruster, Piping and Plumbing Weights

Variable Columns Decimal Location

WGTI (1) 1-10

WGTI (2)

WGTI (3)

WGTI (4)

MM

11-20

21-30

31-40

41

42

Col 7

Col 17

Col 27

Col 37

None

Remarks

The inert weight (piping, valves, thrusters, etc. ) of the system for one propellanl lank per propellant

Same as above fur two I auks per propellant

Same as above for three tanks per propellant

Same as above for four lanks per propellant

Nut used

Inert weight code I

* = 0 if computer is to add these inert weights tu propellant and tank weights

= 1 if computer is not to add these inerl weights

119

I» ir ■ ' m^mmm -—-'-■- - ^ ■

Some input variables have recommended values to be used. These variables and (he values are listed below.

Card Nu.

1 1

Variable

SCBDEN

SIAM

SREPRA

SPISW

SPPCLD

S PSSA

DBHAID

THEDMI

SIT (2)

SDV (3)

PTDEN

PTS1G

Value

20.0 lb/fl3

300. 0 lb-ft-sec

15.Ü deg/day

88.0 Ib/kW

80.0 percent

100. 0 £t2/kW

0. 125 deg (Coarse Mode)

0. 100 deg (Fine Mode)

0.0002 deg/sec

23.0 lb/sec

50. 0 ft/sec

0. 161 lb/in3

176, 000. 0 lbf/in2

ll should be noted however, that it is not mandatory that any of the above values be used. These are only recommended as being repre- sentalive values for a post-1975 SYNCSAT satellite.

120

i .--■- - ^ - _____

This completes the input cards required lo investigate one thruster

combination. If a second system is desired, repeat cards I through 1^ for

system 2, and stack them immediately behind card 12 for system 1. Card 1

is not repeated, but the value of 1NOSTR just updated. The diagram on the

next page demonstrates the stacking procedure required to calculate more

than one propellant system and the control cards required by 1 lie IBM 7040

computer.

121

mr ^ - ' -- ■ ■ ■ ' ■• -

/

INPUT DATA CARDS 2 THROUGH 12 FOR PROPELLANTSYSTEMNO.3

INPUT DATA CARDS 2 THROUGH 12 FOR PROPELLANT SYSTEM NO. 2

INPUT DATA CARDS 2 THROUGH 12 FOR PROPELLANT SYSTEM NO. 1

INPUT DATA CARD NO. 1

ACS COMPUTER PROGRAM (BINARY DECK)

Figun- 34. ACS Compute-r Pvugvzm and Input Data Card Arrangement

122

Mii.-itT^'|TBi».lVlijJi!»*

The next page shows a sample propellant case. The large thrustei-

ls a bipropellant using ClFr and N2H. with a mixture ratio of 2. 0. The

small thruster used for solar pressure corrections, limit cycling and

attitude maintenance contingency is a N^H. gas generator plenum. The

satellite initial gross weight is 3000 pounds and has a cylindrical center-

body and square solar panels. The GIF,, is stored in a tank made of 301

cryostretched stainless steel, while the N^H, is stored in 6A1-4V

titanium. Both are stored at 150 psi and the prcssurant (ISL) is stored

at 3000 psi. The different values for Isp for the same propellant and

thruster for different propulsion functions is a result of the duty cycles for the functions being different.

123

,_^_^ i ■ ■ • ...... -, «» „ , : .-■.- ■ ■■ [ll - ■—

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REFERENCES

1. Barlh, E, C. and Hawk, W., Attitude Control Rocket Requirements, Air Air Force Rocket Propulsion Laboratory.

Z. Nunz, G. J. and Oberstone, J. , Propulsion Systems for Advanced Geosynchronous Satellites, SAMSO TR-70-171, 15 May 1970.

3, Holcomb, L. B. , Satellite Auxiliary - Propulsion Selection Techniques, J PL Report 32-1505, 1 November 1970.

4. Gultman, P. T. , Effects of Gravitational Perturbations on the Behavior of a Satellite in a Nominal 24-Hour Equatorial Circular Orbit, Aerospace Report TDR-469 (5501-50) - 3, July 1965.

•5. Barth, E. C. , "Application of DART In Space Relay and Data Management Satellite, " AFRPL Internal Memorandum, 20 May 1968.

149

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|pWii.»M^..im.îJ1i,MW,l|.|i)^.„.i,Ai^ IF^TWipFTT^nf—RTffll

RAYMOND D. KLOPOTEK, Capt, USAF Air Force Rocket Propulsion Laboratory

Project Engineer, Subsystems Branch, Liquid Rocket Division of the

Air Force Rocket Propulsion Laboratory.

He was a Project Engineer in the Combustion and Heat Transfer

Section where he planned, managed and conducted the Titan 111 Transtage

Combustion Program. He was also a Project Engineer on the Pulse Motor

Combustion Instability Program.

He was assigned as an Advanced Plans Officer and was directly

responsible for the Space and Ballistic Missile and portions of the labora-

tory's overall long-range planning. Currently, he is responsible for the

conception, definition and analysis of new liquid rocket propulsion

technology for satisfying advanced Air Force missions.

WALDEN L. S. LAUKHUF, Capt, USAF Air Force Rocket Propulsion Laboratory

Project Engineer, Subsystenis Branch, Liquid Rocket Division of the

Air Force Rocket Propulsion Laboratory.

Upon entering active duty, he was assigned to this branch where he is

responsible for the conception, definition and analysis of new liquid rocket

propulsion technology for satisfying advanced Air Force missions.

Currently, he is working on an in-depth design and analysis study for

advanced satellite propulsion concepts.

150

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