by William R. Kerslake, Joseph F. Wassexbauer, ard Paul M. Margssian ... including power to heat the feed system and the neutralizer, was 222 watts.
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A iviEKClJKKY ELECTRUN-BOMBARDMENT ION THRUSTOR SUITABLE FOR SPACECRAFT STATION KEEPING AND ATTITUDE CONTROL
Pl
Microiicne w r r ff
663 July 66
by William R. Kerslake, Joseph F. Wassexbauer, ard Paul M. Margssian Lewis Research Center Cleveland, Ohio *
;
TECHNICAL PAPER proposed for preserL&kn at Fifth Electric Propulsion Meeting sponsored by the American Institute of Aercnacntics apd A S ~ ~ Q L X J ~ Z C S S m Diego, California, M a c h 7-9, 1966
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
-
WASHINGTON, D.C.
-
1965
I
A MERCURY ELECTRON-BOMBARDMENT ION THRUSTOR SUITABLE FOR SPACECRAFT STATION KEEPING AND ATTITUDE CONTROL
by William R. Kerslake, Joseph F. Wasserbauer, and Paul M. Margosian Lewis Research Center Cleveland, Ohio
TECHNICAL PAPER proposed for presentation at Fifth Electric Propulsion Meeting sponsored by the American Institute of Aeronautics and Astronautics San Diego, California, March 7-9, 1966
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
A MERCURY ELECTIION-BOMBARDMENT I O N THRUSTOR SUITABLE FOR SPACECWT STATION KEEPING AND ATTITUDE CONTROL by W i l l i a m R . Kerslake, Joseph F. Wasserbauer, and Paul M. Margosian Lewis Research Center National Aeronautics and Space Administration Cleveland, Ohio ASSTRACT A n i o n t h r u s t o r system, including a new type of mercury feed system and
a shielded n e u t r a l i z e r , has been designed and t e s t e d at a t h r u s t l e v e l of 0.5 millipound.
The use of
8
r a d i a l f l a w propellant d i s t r i b u t o r and a n
oxide-coated brush cathode r e s u l t e d i n discharge power l o s s e s one-fourth of those previously encountered at t h i s t h r u s t l e v e l .
Different lengths and
diameters of t h e discharge chamber wer? t e s t e d t o e s t a b l i s h a compromise between discharge power l o s s e s and propellant u t i l i z a t i o n .
The f i n a l design
discharge chamber w a s 5-centimeters i n d.iameter and 7.5-centimeters long. The complete f l i g h t - t y p e t h r u s t o r system used a permanent magnet f i e l d and weighed (without p r o p e l l a n t ) 1 . 3 6 kilograms.
The power t o t h r u s t r a t i o ,
including power t o heat t h e feed system and t h e n e u t r a l i z e r , was 222 w a t t s per millipound at a t h r u s t of 0.65 millipound and a s p e c i f i c impulse of 3050 seconds.
The oxide-coated brush cathode w a s endurance t e s t e d 1553
hours i n a t h r u s t o r and an i d e n t i c a l cathode w a s heat cycled i n a separate t e s t for 418,000 cycles before f a i l u r e .
INTFODUCTION The primary objective of t h i s work w a s t o demonstrate mercury electronbombardment i o n t h r u s t o r performance s u i t a b l e f o r s t a t i o n keeping and a t t i t u d e c o n t r o l of s a t e l l i t e s .
For simplicity, t h e same b a s i c t h r u s t o r
design and t h r u s t l e v e l were contemplated f o r both t h e station-keeping and
TM X-52163
3 0.5 millipound and a propellant u t i l i z a t i o n e f f i c i e n c y of 80 t o 50 percent.
I n t h e course of' t h e research program, various discharge chamber diameters
.
and lengths were t e s t e d t o determine minimum discharge chamber l o s s e s and
m a x i m u m propellant Q t i l i z a t i o n s .
I n i t i a l l y t h e program w a s conducted w i t h
an e a s i l y varied electromagnetic f i e l d , and l a t e r e f f o r t was concentrated on t h e use of permanent magnets t o reduce weight, power and system complexity. The mercury propellant feed system chosen w a s a positive-pressure l i q u i d feed t o a heated porous tungsten plug which controlled t h e vapor flow r a t e and separated t h e vapor-liquid phase.
I n addition, t h e porous plug served
as an on-off valve because neglible flow passed through t h e plug when it w a s cold.
Performance data a r e presented f o r two f l i g h t - t y p e t h r u s t o r s
as well as f o r an extended t e s t with a preliminary design. APPAFAWS ANG PROCECTJRE F i E i r e c l ( 9 ) a n d 1 (h) are
nhntograohs of t h e f l i g h t - t p e t h r u s t o r with
an electromagnetic f i e l d e o i l , and figure l ( c ) i s a photograph of a permanent magnet version.
Figure 2 i s a schema€ic view of a t h r u s t o r , i n d i c a t i n g t h e
r e l a t i v e l o c a t i o n s of t h e dischasge chamber, cathode, d i s t r i b u t o r , magnetic c o i l , and a c c e l e r a t o r g r i d s . Flight - m e Thrustor The nominal s i z e (anode diameter) of t h e f l i g h t - t y p e t h r u s t o f was determined t o a l a r g e extend by t h e l i f e t i m e requirements of t h e a c c e l e r a t o r grid.
D e t a i l s of t h e accelerator @;ridediameter s e l e c t i o n may be found i n
appendix A.
B r i e f l y , t h e g r i d erosion i s d i r e c t l y r e l a t e d t o t h e square
of beam current density and inversely t o t h e propellant u t i l i z a t i o n .
t h e d e s i r e d l i f e t i m e of 13,000 hours and t h r u s t of 0.5 millipound, t h e
For
4 minimum accelerator diameter w a s determined t o be 3 t o 5 centimeters at p r o p e l l a n t u t i l i z a t i o n e f f i c i e n c i e s of 80 t o 50 percent. design w a s 5 centimeters i n diameter.
The f i n a l a c c e l e r a t o r
The a c c e l e r a t o r and screen g r i d s were
both f a b r i c a t e d of a 0.16-centimeter-thich molybdenum sheet. d r i l l e d i n both g r i d s on a 0.635-centimeter
Holes were
e q u a l a t e r a l t r i a n g u l a r spacing.
The screen g r i d and a c c e l e r a t o r holes were 0.476 and 0.317 centimeters i n diameter, respectively.
The a c c e l e r a t o r holes were made smaller t o b o t h
i n c r e a s e t h e web m a t e r i a l between holes ( t h u s i n c r e a s i n g t h e l i f e t i m e ) and t o somewhat decrease t,he loss of n e u t r a l propellant, through t h e g r i d system. The screen-accelerator g r i d separation w a s held a t 0.15%.01 centimeter by shielded aluminum oxide b a l l i n s u l a t o r s . By using t h e r e s u l t s of t h e v a r i a h l e geometry t . b u . s t o r t e s t s , t h e discharge chamber w a s designed with an anode diamet,er Da and a length
La
of 5 centimeters
of 7.5 centimet*ers. All sheet metal p a r t s were made of
nonmagnetic s t a i n l e s s st.eel.
A magnetic c o i l produced a tapered f i e l d with
magnitudes of 56 gauss at t h e d i s t z i b u t o r and 24 gauss a t t h e screen.
The
permanent magnets produced a f i e l d w i t h a range of near zero t o a m a x i m u m of 105 gauss.
Four rod magnets, 0.'785 centimet,er i n diameter and 8 . 2 5 cen-
t i m e t e r s long were located between mild s t e e l pole pieces.
This con.fi.guration
appears i n a photograph i n f i g u r e l . ( a ) and i s a l s o sketched i n f i g u r e 3. The chamber cathode was a tant.alum brush (0.5 cm i n diameter and 1 . 2 cm long) coated with Radio Mix No. 3 (57 percent BaC03, 42 percent SrC03, and 1 percent CaC03) and had a surface a r e a of 1.8 square centimeters.
The
cathode w a s supported between two copper rods and w a s c e n t r a l l y l o c a t e d i n f r o n t of and p a r a l l . e l t o t h e plane of t h e d i s t r i b u t o r ( f i g . 3 ) .
The cathode
5 c
w a s approximately 1 centimeter from the d i s t r i b u t o r .
The cathode brush
diameter w a s sized l a r g e enough t o make t h e required l i f e t i m e f e a s i b l e and The l e n g t h w a s determined
y e t s m a l l enough t o reduce t h e thermal l o s s e s . by t h e emitting a r e a required.
T h i s surface a r e a was a c t u a l l y l a r g e r than
nominally required (nominal emission, 1 A/sq cm) because temperature gradiants i n a u n i t of t h i s s m a l l s i z e g r e a t l y reduced t h e emission at t h e ends. The propellant d i s t r i b u t o r w a s of a r a d i a l type (first reported i n reference 4 ) t o increase t h e cathode l i f e t i m e and a t t h e same t i m e produce a high t h r u s t o r t o t a l e f f i c i e n c y .
The inner hole diameter of t h e d i s t r i b u t o r
p l a t e w a s 2.54 centimeters, and t h e distance between t h e cathode mounting block and t h i s p l a t e w a s about 0.32 centimeter.
Other types of propellant
d i s t r i b u t i o n were not attempted i n t h i s program. Propellant Feed System
i n generai, zne Ieea s y s ~ e mcuiibibieG ~ J Ta D & L L C ~ .
i-ESEi-i-G . The difference between n e u t r a l i z e r and i o n beam currents probably represents e l e c t r o n s that were drawn i n t o t h e beam from other sources w i t h i n t h e vacuum tank.
The coupling voltage increased with increasing
beam c u r r e n t as shown i n f i g u r e 9(b) u n t i l it leveled o f f near 35 v o l t s
16 and 0.040 ampere.
The s c a t t e r i n t h e coupling voltage d a t a w a s t y p i c a l
and might have been caused by l o c a l differences i n t h e cathode a c t i v a t i o n . Previously i n a c t i v e cathode areas may become a c t i v a t e d by i o n bombardment, and as space-charge-limited flow propably e x i s t s , a l a r g e r emitting a r e a would require l e s s coupling voltage. Because of t h e s e n s i t i v i t y of t h e n e u t r a l i z e r cathode and s h i e l d p o s i t i o n , most t e s t s were considered successful i f t h e coupling voltage were i n t h e range of 50 t o 200 v o l t s .
The unusually low voltages obtained i n f i g u r e 9
may be i n d i c a t i v e of t o o l i t t l e shielding or t o o much immersion of t h e cathode i n t h e beam t o r e a l i z e a long n e u t r a l i z e r l i f e t i m e . Propellant flow c a l i b r a t i o n .
-
The porous tungsten plug that gave t h e
proper flow r a t e at t h e desired temperature l e v e l had (according t o t h e vendor) c a p i l l a r i e s w i t h a 3.8-micron pore radius and contained 1.6X106 pores per square centimeter.
The porous tungsten plug t i p (shown schematically
i n f i g . 10) was flow c a l i b r a t e d by operating it at a constant temperature and weighing t h e weight loss from t h e r e s e r v o i r .
Mercury flow rates as a
function of temperature f o r two t i p s made from i d e n t i c a l porous tungsten a r e a l s o shown i n f i g u r e 10. t o +3 percent.
Flow r a t e s f o r a given plug were reporducible
Af'ter 100 hours of operation, t h e flow r a t e w a s again checked
and found t o be e s s e n t i a l l y unchanged.
There was n e g l i g i b l e heat conduction
back t o t h e r e s e r v o i r , 15 centimeters away.
I n a seqarate t e s t , t h e l i q u i d
head pressure w a s increased t o 2 atmospheres, and no l i q u i d mercury w a s forced through t h e porous plug. from 400'
Using 8 w a t t s , t h e plug temperature increased
t o 580' K i n 30 seconds, and cooled down t o 420'
a f t e r t h e power w a s turned o f f .
K, 45 seconds
(The temperature cycling range corresponded
17 t o two orders of magnitude change i n flow r a t e ) .
The temperature response
of t h e feed t i p w a s judged t o be somewhat slow f o r an a l t i t u d e c o n t r o l t h r u s t o r and would probably need improving before use i n this a p p l i c a t i o n . Magnetic f i e l d s .
-
The measured magnetic f i e l d s t r e n g t h s f o r t h e
electromagnetic f l i g h t - t y p e t h r u s t o r and permanent magnet f l i g h t -type t h r u s t o r a r e compared i n f i g u r e 11. The electromagnetic f i e l d w a s produced by a s i n g l e l a r g e solenoidal c o i l .
The tapered f i e l d (lower f i e l d i n t h e downstream
d i r e c t i o n ) was found i n previous t e s t s ( r e f . 8 ) t o a i d t h e discharge chamber performance.
A s a r e s u l t of' weight and s i z e r e s t r i c t i o n s , t h e permanent
magnet f i e l d shown i n f i g u r e l l ( b ) contained a hump at an a x i a l p o s i t i o n of about 2.5 centimeters from t h e d i s t r i b u t o r , r a t h e r than t h e continuously diverging f i e l d produced by t h e solenoid.
The permanent magnet f i e l d w a s
a l s o measured 1 . 3 centimeters from t h e c e n t e r l i n e t o give an i n d i c a t i o n of t h e radial gradient.
It, w a s f o m a
trlar;
~ i l criel\; S+i-CI&h
ct t h e 1 3-
centimeter r a d i u s increased about 5 percent over t h e c e n t e r l i n e values f o r
most of t h e a x i a l length.
Greater increases near t h e d i s t r i b u t o r end were
probably due t o shape of t h e d i s t r i b u t o r pole piece (shown sketched i n f i g . 3 ) . Figure 1 2 compares t h e performance with t h e two t h r u s t o r s i n terms of tne dependence of discharge power per beam i o n on t h e propellant u t i l i z a t i o n . The discharge power gradually decreases at lower u t i l i z a t i o n s and r a p i d l y i n c r e a s e s at propellant u t i l i z a t i o n e f f i c i e n c i e s higher than 60 o r 70 percent. The shape of t h e f i e l d of t h e permanent magnet t h r u s t o r apparently was responsible f o r t h e decrease i n t h e maximum propellant u t i l i z a t i o n , but below propellant u t i l i z a t i o n e f f i c i e n c i e s of 50 percent t h e r e w a s no difference i n t h e discharge l o s s e s per beam ion.
Some of t h e d i f f e r e n c e i n t h e maximum
. .*
i
18 propellant u t i l i z a t i o n could a l s o have been t h e r e s u l t of a v a r i a b l e .activ a t i o n s t a t e of t h e cathode from one t e s t t o another.
Other t e s t s s i m i l a r
t o those shown i n f i g u r e 1 2 r e s u l t e d , a f t e r longer running times, i n a d i s chazge power per beam i o n of 400 t o 500 e l e c t r o n v o l t s per i o n at a propellant u t i l i z a t i o n e f f i c i e n c y of 50 percent. Thrustor Performance and Endurance Tests Variable geometry t h r u s t o r .
-
A long duration run w a s conducted t o
endurance t e s t t h e oxide-coated brush cathode i n a t h r u s t o r .
A 5-centimeter-
diameter t h r u s t o r of t h e type used i n reference 6 was u t i l i z e d f o r t h e t e s t . The t h r u s t o r w a s modified by s u b s t i t u t i n g a radial flow d i s t r i b u t o r f o r t h e uniform-flow type.
The brush cathode and coating were i d e n t i c a l t o those
l a t e r used i n t h e f l i g h t - t y p e t h r u s t o r .
The t e s t proceeded f o r 1553 hours
during which time t h e t h r u s t o r was d e l i b e r a t e l y shut down and r e s t a r t e d 54 times without removal from t h e vacuum chamber.
The average t e s t values
a r e l i s t e d i n t h e l a s t column of t a b l e I. The t h r u s t o r i o n beam was maintained between 0.015 and 0.020 ampere by a d j u s t i n g the cathode heating power t o give an emission of 0.2 t o 0.4 ampere. This mean value of t h e discharge power per beam i o n s t e a d i l y dropped from
an i n i t i a l value of 600 t o a f i n a l value of 400 e l e c t r o n v o l t s per i o n . The discharge voltage w a s held constant a t 35 v o l t s .
This p o t e n t i a l w a s
a compromise between lower cathode s p u t t e r i n g rates at lower voltages and
a lower discharge chamber loss per beam i o n at somewhat higher voltages. The cathode heating power slowly rose from about 1 5 t o 22 w a t t s at 1500 hours.
(Details of t h i s heating power curve may be found i n r e f . 3.)
The
magnetic c o i l power loss ( 2 7 W) w a s considered unduly high because of an
19 i n e f f i c i e n t design.
An e f f e c t i v e value of 8 watts w a s t h e r e f o r e used when
computing t h e power-to-thrust. r a t i o value of 256 watts per millipound l i s t e d i n table I.
Also, a value of 8.7 w a t t s w a s added t o t h e t o t a l power t o
represent t h e feed system power.
(The a c t u a l feed system f o r t h i s endurance
t e s t w a s a steam-heated mercury vaporizer w i t h an o r i f i c e plug.) A t 1551 hours t h e r e vas a l a r g e pressure excursion i n t h e vacuum f a c i l i t y
t o the
Al5hough emission was r e e s t a b l i s h e d a f t e r t h i s
t o r r range.
excursion, one of t h e two cathode heatter wires w a s broken, and t h e other f a i l e d a f t e r 2 hours of operahion.
There w a s l i t t l e erosion of t h e heater
wires, and t h e f a i l u r e w a s i n a portion w e l l protected from d i r e c t i o n bombardment
.
Microphotographs of wire cross s e c t i o n indicated a l a r g e amount
of a second phase, which may have been tantalum n i t r i d e or oxide.
i n d i c a t e probable f a i l u r e by gas embrit5lement. 8
.
LUO,L*Lll&
u l 2 l
tk
i&,L&2Xd
Flight thrustors.
-
This would
A t l e a s t h a l f of t h e oxide
CGZ,kik.
An i n i t i a l t e s t w a s performed on t h e complete f l i g h t -
type t h r u s t o r , which included a porous plug feed system, electromagnet c o i l , and n e u t r a l i z e r .
The average values of t h i s t e s t a r e a l s o l i s t e d i n t a b l e I.
S p e c i f i c values of beam current, discharge voltage, cathode heating power, a n d magnetic f i e l d were chosen t o optimize t h e propellant u t i l i z a t i o n , discharge power l o s s e s , and cathode l i f e t i m e .
The r e s u l t i n g power-to-thrust
r a t i o was 247 watts per millipound at a t h r u s t of 0.65 millipound.
This
power t o t h r u s t r a t i o was somewhat greater than normal because t h e discharge power was higher due t o operation at a higher propellant u t i l i z a t i o n .
Also,
t h e a c c e l e r a t o r impingement current (for unknown reasons) w a s s i x times
i t s u s u a l value.
The n e u t r a l i z e r coupling voltage w a s a l s o somewhat high
20
at 130 t o 200 v o l t s and t h e net a c c e l e r a t i n g voltage of t h e i o n beam w a s reduced from 4000 t o 3800 v o l t s when t h e beam t h r u s t w a s calculated. The f l i g h t t h r u s t o r was next modified t o incorporate a permanent magnetic field.
The permanent magnetic f i e l d t h r u s t o r w a s t e s t e d f o r 13 hours a t
e s s e n t i a l l y constant conditions t o determine i t s steady s t a t e operating performance.
The r e s u l t s of this t e s t a r e l i s t e d i n t a b l e I.
A beam current
of 0.0225 ampere w a s produced w i t h 1 4 4 w a t t s t o give a power-to-thrust of 222 w a t t s per millipound.
ratio
The discharge l o s s e s and a c c e l e r a t o r impingement
were normal and much lower than t h e electromagnet f l i g h t t h r u s t o r t e s t . Values were estimated f o r both t h e n e u t r a l i z e r cathode heating power and propellant feed power, although no c o r r e c t i o n for n e u t r a l i z e r coupling voltage
was m a d e t o the calculated t h r u s t .
No n e u t r d i z e r was used, and a steam-
heated vaporizer replaced t h e e l e c t r i c a l l y heated porous plug. Cathode heat cyclin@;.
- For a t h r u s t o r
t o perform an a t t i t u d e c o n t r o l
mission it must be a b l e t o be cycled on and o f f a g r e a t many times.
This
imposes t h e stress of thermal cycling and t h e g r e a t e s t stress w i l l be on t h a t component t h a t has t h e l a r g e s t temperature v a r i a t i o n , namely, t h e cathode. Therefore, a s e r i e s of t e s t s was undertaken t o thermally cycle t h e cathode only.
An oxide-coated tantalum brush i d e n t i c a l t o t h a t used i n t h e f l i g h t
t h r u s t o r w a s mounted from a p a i r of copper supports and heated i n a b e l l
jar. After an i n i t i a l period of 1/2 hour t o decompose t h e a l k a l i n e e a r t h carbonates, the cathode heating voltage w a s snapped on and off i n a c y c l i c manner allowing 15 seconds f o r heating and 15 seconds f o r cooling.
The
cathode reached an equilibrium temperature of 1250° K and 17 w a t t s heating power i n 10 seconds.
During cooling, t h e temperature dropped below 900° K
21
( l i m i t of t h e pyrometer) -in 5 seconds. i n t h e mid-10"
The b e l l jar pressure w a s normally
t o r r range.
The first t h r e e t e s t s Bailed a f t e r 5000 t o 15,000 cycles.
A fourth
t e s t , c o n s i s t i n g of a bare brush with no oxide coating, w a s made t o separate t h e e f f e c t s of any chemical r e a c t i o n o r physical gas absorbtion between t h e oxide and t h e tantalum brush from t h e e f f e c t s of thermal f a t i g u e . The b m e brush w a s cycled 84,000 t i m e s at which point t h e t e s t w a s stopped because it exceeded t h e 50,000 cycle requirement of t h e estimated mission. The core wires were b r i t t l e , hawever, and t h e brush borke a p a r t when it
w a s removed from t h e holder. A f i f t h t e s t , i n which a d i f f e r e n t solvent w a s used t o clean t h e brush
before coating, w a s attempted with t h e i d e n t i c a l type of brush cathode. T h i s t e s t , s i m i l a r l y run t o t h e f i r s t three t e s t s , l a s t e d f o r 418,000 cycles
before t h e core wires or t n e brusil L L U ~ ~L; Z ;;:.
zf cx5de
f x m hniqh,
either
by evaporation or by s p a l l i n g away, was n e g l i g i b l e . The key t o this improved cathode l i f e t i m e w a s probably not i n t h e solvent used t o clean t h e brush but i n a thermal gradient t h a t e x i s t e d i n t h e cathode. To r e a c h a surface temperature of 1250° K, t h e i n t e r i o r core temperature
m u s t b e about 1600° t o 1700° K.
A t 1703O K, t h e r e a c t i o n r a t e between barium
oxide and tantalum ( r e f . 9 ) becomes ( t h e o r e t i c a l l y ) s i g n i f i c a n t l y high. Chemical r e a c t i o n rates a r e very s e n s i t i v e t o temperature, and perhaps local v a r i a t i o n s i n t h e brush oroxide coating could cause s p e c i f i c brushes t o o p e r a t e e i t h e r above o r below a c r i t i c a l r e a c t i o n temperature.
Nevertheless,
t h e demonstration of 418,000 cycles proves that under t h e c o r r e c t conditions
a b r u s h cathode should conservatively be able t o meet a mission requirement
22
of 50,000 thermal cycles.
CONCLUDING REMARKS A lower l i m i t f o r t h e t h r u s t of an electron-bombardment i o n t h r u s t o r
appears t o be about 0.1 t o 0.3 millipound f o r s a t i s f a c t o r y performance i n
a s a t e l l i t e c o n t r o l system.
If t h e chamber diameter i s reduced t o 2.5
centimeters or lower, w a l l recombination l o s s e s r e s u l t i n a power t h a t becomes p r o h i b i t i v e l y high.
discharge
To avoid w a l l recombinations, t h e
t h r u s t o r may be made l a r g e r i n diameter, b u t keeping t h e t h r u s t l e v e l constant r e q u i r e s a reduction of t h e beam current and, hence, plasma density.
At
t h e s e lower d e n s i t i e s , t h e propellant u t i l i z a t i o n w a s sharply reduced.
A
low propellant flow of 0.020 ampere of n e u t r a l s i n a 5-centimeter-diameter t h r u s t o r (corresponding t o a t h r u s t of 0.25 mlb) caused severe d i f f i c u l t y i n maintaining a discharge and e x t r a c t i n g a bean of more than 50-percent propellant u t i l i z a t i o n e f f i c i e n c y .
A good compromise f o r a mercury e l e c t r o n -
bombardment t h r u s t o r of 0.5-millipound t h r u s t w a s a chamber 5 centimeters i n diameter and a propellant u t i l i z a t i o n of 50 t o 60 percent.
The low
propellant u t i l i z a t i o n i s j u s t i f i b l e f o r t h e intended mission of s a t e l l i t e control. The permanent magnet t h r u s t o r system, c o n s i s t i n g of a propellant tank and vaporizer, t h r u s t o r , and n e u t r a l i z e r , i s considered t o be adaqtable t o launch environment and mission requirements. e f f i c i e n t f o r t h e low t h r u s t l e v e l .
It has l i g h t weight and i s
Furthermore, t h e cathode has demonstrated
( i n a separate t h r u s t o r ) s u b s t a n t i a l l i f e t i m e i n vacuum tank t e s t s .
The
b e s t power-to-thrust r a t i o achieved with t h e complete t h r u s t o r w a s 222 w a t t s per millipound a t 0.65-millipound t h r u s t .
If other t h r u s t o r missions can
i 23
t o l e r a t e or require higher t h r u s t , the same 5-centimeter-diameter f l i g h t type t h r u s t o r could be used a t a higher e f f i c i e n c y or a lower r a t i o of power t o t h r u s t than obtained herein.
Data of reference 2 i n d i c a t e , i n f a c t ,
t h a t a 5-centimeter-diameter t h r u s t o r operates e f f i c i e n t l y up t o 4 millipounds of t h r u s t .
Based on t h e s e data, a 5-centimeter-diameter t h r u s t o r , operating
at a t h r u s t of 3 millipounds and a s p e c i f i c impulse of 4000 t o 5000 seconds, would produce a power-to-thrust r a t i o of about 200 w a t t s per millipound. (This includes a power l o s s of 13 watts f o r t h e feed system and 15 w a t t s
for n e u t r a l i z a t i o n . )
The accelerator l i f e t i m e , however, would be more
c r i t i c a l at t h i s higher t h r u s t l e v e l .
24 APPENDIX A
*
An endurance run w a s made with a 10-centimeter-diameter permanent magnet t h r u s t o r s i m i l a r t o t h a t described i n reference 10.
This t h r u s t o r
w a s operated a t a beam current of 0.13 ampere f o r t h e f i r s t 768 hours and 0.25 ampere f o r t h e l a s t 615 hours.
A t o t a l of 13821 hours w a s accumulated
on a s i n g l e molybdenum a c c e l e r a t o r g r i d 0.156 centimeter t h i c k .
The hole
s i z e and p a t t e r n were i d e n t i c a l t o those reported herein f o r t h e f l i g h t type t h r u s t o r (0.32-cm-diameter spacing).
holes on a 0.64-cm e q u i l a t e r a l t r i a n g u l a r
After 1383 hours t h e a c c e l e r a t o r w a s reweighed, remeasured f o r
thickness, and rephotographed.
The average hole diameter enlargement w a s
0.02 centimeter while t h e m a x i m u m (near t h e g r i d c e n t e r ) w a s 0.05 centimeter.
The thickness of t h e g r i d w a s reduced 0.007 centimeter over t h e center 5-centimeter diameter and a diminishing amount towards t h e outer edges of the grid.
The calculated weight l o s s from l i n e a r measurements was 5.7k.7
grams,
while t h e measured weight l o s s w a s 5.36 grams. The t o t a l impingement on t h e a c c e l e r a t o r g r i d was 2.35 ampere-hours
at an accelerator g r i d voltage of -1000 v o l t s , a net a c c e l e r a t i n g voltage of 4000 v o l t s , and an average propellant u t i l i z a t i o n e f f i c i e n c y of 80 percent. The erosion from t h i s impingement, i f continued l i n e a r l y , would r e d m e t h e web material between holes t o zer3 with a t o t a l erosion given by t h e r a t i o of web thickness t o t h e m a x i m u m wear, 0.05 centimeter, times 2.35 amperehours or 1 4 . 9 ampere-hours.
The l i f e t i m e of t h e a c c e l e r a t o r w a s assumed
t o be t h e point at which t h e web thickness w a s zero. Using t h e technique of reference 11, an erosion of 1 4 . 9 ampere-hmrs,
and a l i f e t i m e of 13,000 hours, f i g u r e 13 was prepared.
The t h r u s t o r
25
diameter becomes a function of t h r u s t and propellant u t i l i z a t i o n .
The
a c c e l e r a t o r impingement due t o charge exchange w a s calculated by equation B(16) of reference 11. A d i r e c t impingement value (estimated from t h e measured impingement values of t h e 1383-hr t e s t minus a calculated charge exchange v a h e ) was a l s o added t o t h e charge exchange value.
It i s i n t e r e s t i n g t o
note t h a t f o r a given t h r u s t , propellant u t i l i z a t i o n , and erosion rate, t h e a c c e l e r a t o r g r i d l i f e t i m e i s proportional t o t h e f o u r t h power of g r i d diameter (eq. ( 6 ) , r e f . 11).
26 REFERENCES
1. Molitor, H. H.
, "Ion Propulsion
System f o r S t a t i o n a r y - S a t e l l i t e Control,"
J . Spacecraft and Rockets, vol. 1, no. 2 , pp. 170-175, 1964. 2. Pawlik, E.
V. and Nakanishi, S., "Experimental Evaluation of S i z e E f f e c t s
on Steady-State Control P r o p e r t i e s of Electron-Bombardment Ion Thrustor ? " NASA 'I" D-2470 (1964). 3. Kerslake, W. R . ,
"Preliminary Operation of Oxide-Coated Brush Cathodes
i n Electron-Bombardment Ion Thrustors,," NASA TM X - 1 1 0 5 (1965) 4. Kerslake, W. R . ,
"Cathode D u r a b i l i t y i n t h e Mercury Electron-BomSardment
Ion Thrustor."
Paper No. 64-683, AIAA (1964).
5. Nakanishi, S. and Pawlik, Eugene V . ,
"Preliminary Experimental Operation
of High-Voltage I s o l a t i o n Device f o r Propellant System of an Ion Rocket ,'I NASA
TM X-1026 (1964)
6 . Reader, P. D.,
"Scale E f f e c t s on Ion Rocket Performance," ARS J., vol. 3 2 ,
no. 5, May 1962, pp. 711-714. 7 . Kauhan, H. B.,
"Performance S o r r e l a t i o n f o r Electron-Bombardment Ion
Sources,11 NASA TN D-3041 ( 1 9 6 5 ) . 8. Reader, P. D
.,
" I n v e s t i g a t i o n of a lO-Cent,imeter-Dimeter Electron-
Bombardment Ion Rocket)" NASA TN D-1163 ( 1 9 6 2 ) . 9. Rittner, E. S.,
"A T'heoret,ical Study of t h e Chemistry of t h e Oxide Cathode,"
P h i l i p s Res. Rep., vol. 8, 1953, pp. 184-238. 10. Reader, P. D.,
"An Eiectron-Bonibardnient Ion Rocket with a Permanent Magnet , ' I
Paper No. 63031-63, AIAA (1963).
(See also Astronaut. and Aerospace
Eng., vol. 1, no. 9, Oct. 1963, p. 8 3 . ) 11. Kerslake, W . R . "Charge-Exchange E f f e c t s on t h e Accelerator Impingement
of an Ekectron-Bombardment Ion Rocket," NASA T2N D-1657 (1963).
27
TABLE I.
- PERF0F"CE
OF MERCURY ELECTRON-BOMBAFiDMENT THRUSTORS
'light-type t h r u s t o r s Phrustor cathode endurance Electro- Permanent test magnet magnetic Beam c u r r e n t , A Propellant flow r a t e , equivalent amperes of n e u t r a l s Net a c c e l e r a t i n g voltage, V Accelerator voltage, V Discharge voltage, V Thrust, mlb Length of t e s t , hr Neutralizer cathode l i f e , hr Beam power, W Discharge power, W Discharge cathode power, W Neutralizer cathode power, W Magnetic c o i l power, W Propellant feed power, W Accelerator d r a i n power, Total power, W -- I Power t o t h r u s t r a t i o , w / C u S p e c i f i c impulse, sec Weight of complete t-hrustor system ( l e s s p r o p e l l a n t ) , Kg
0.023
0.0225
0.018
0.035 4000
0.047
0.034 4000 -1000 35 0.52 1553
-1000 25
"0.65 95
4000 -1000
30 0.65 13
58
------
92 .O 21.6 1 9 .o
90 .o 12.9 1 6 .O
15.5 8 .O 8.7 6 .Q
b15 .5 0 b8.7 0.9
170.8
144 .O
72 .O
8.0 15 e 0 b15 .5 '27 .O b8 .7 1.o
"4050
3020
'128.2 2 47 3340
5 .O
3 .O
------
0r.0
LUL
773 YLlY
a
Calculated f o r a net a c c e l e r a t i n g voltage of 3800 v o l t s ( i n s t e a d of 4000 V . ) t o c o r r e c t for n e u t r a l i z e r coupling.
b
Neither n e u t r a l i z e r nor e l e c t r i c a l feed system w a s used; estimated values.
C
Design of electromagnet t o o large; t o t a l power reduced by 1 9 w a t t s .
___
r G r o u n d e d shield
,
&Neutralizer cathode
C-65-1396
(a) Electromagnetic thrustor, exhaust end. Accelerator insulator shield7 Cathode terminal Thermal isolation tube7 Thermocouple -,
i
~
LPorous plug heater
\
i LMoijntin? pI;ltp
LFrom propellant reservoir
C-i4733 (b) Electromagnetic thrustor, upstream end with outer screen removed.
C-65-35" (c) Permanent magnet thrustor with outer screen removed.
Figure 1. - 5-Centimeter-diameter flight-type thrustors.
.
Magnetic field coi Is7
Radial flow distributor r) Propellant
k1
-
25.4
C-- La
j,1 -f
Discharge chamber Screen +DaI
11
'LCathode grid-, I 1------------I '!
lonization chamber p e r
Accelerator power supply
I Flgure 2. - Schematic view of variable geometry electron-bombardment thrustor. (Outer screen to reduce Stray currents and suppress arcing i s not shown. 1
Distributor
wle oiece r- - r - - -
(mild steel) Radial-flow distributor 7
7 Permanent magnet $1 Screen pole piece7 '\\ (mild steel) '-
~
',
Fine mesh screen
Propellant
4-
-
Neutralizer
za
Boron nitride shield
_/.on IJU
5-Cd
diameter
Cathode power
1
Permanent magnet
.'
lonizationchamber power supply Figure 3.
'
I
Ion-beam power supply
- Permanent magnet flight type thrustor.
Accelerator power supply
I II I
LID Diameter,
0 2.0 0 1 . 5
cm 2.5 5.0
Magnetic field, G
I
Screen Distributor 50 210 38 120
0.
dc
N
m
-~ .1
I
w
~~
.3 .5 .7 .9 Propellant utilization efficiency, vu, percent
Figure 4. - Performance of two different diameter discharge chambers in variable geometry thrustors. Net accelerating voltage, 4ooo volts; accelerator voltage, -lWvolts; discharge voltage 14 to 30 volts; neutral propellant f l w , 0.047 ampere.
Discha rge voltage, V
current,
22. 8 32.5 28.6
19 19 19
5.0 20.8 25 21.2 1.5 31.5 1.0 __ 25.5 19.0 2.5 1.5 31.8 -
29
2.5 1.5 1.0
22.5
Magnetic field, G
GqTGxr
19 19
I
63 48 60
62 63 60
W
0
1 2 3 4 5 Length -to-diameter ratio, La/Da
Figure 5. - Discharge chamber length varied in 5-centimeterdiameter variable geometry thrustor. Net accelerating voltage, 4ooo volts; accelerator voltage, -1000 volts; neutral propellant flow, 0.047 ampere.
5
I
lon-chamber discharge potential, V 0 30
grids,
0.152 .145
.32 .64
t
' wT
N
$ "
_ _ _ _ _ _ _ _ _ _- Neutral
--
U c
E L
U 3
2
4
6
20
0
8xld
Net accelerating voltage, V
40
60
Magnetic field strength at screen, G
Figure 6. -Accelerator impingement currents versus net accelerating voltage for 5-centimeterdiameter variable geometry thrustor. Screen thickness, 0.16 centimeters; magnetic field at distributor, 100 gauss; at screen, 35 gauss; ratio of net-to-total accelerating voltage, 0. & ion-beam current, 0.025 ampere.
Figure 7. - Discharge chamber performance for the f I ight -type electromagnetic th r ustor. Discharge chamber diameter, 5 centimeters; net accelerating voltage, 4ooo volts; accelerator voltage, -loo0 volts; cathode heating F e r , constant at 35 watts.
L
0
c
e
al 2 a l
U
0l
k2
a4
p6
8x103 i
Net accelerating voltage, V Figure 8. -Net accelerating voltage versus accelerator impingement c u r r e n t for flight-type thrustor. Beam current, 0.030 ampere; discharge voltage, 35 volts; magnetic field, 22 gauss at screen, 60 gauss at distributor; ratio of net-to-total accelerating voltage, 0.4 neutral propellant flow, (Lobo ampere.
0
1"
E 3
3 v)
E
-I
OP
-
,
t
Field
r-Distributor
Thrustor
centerline --- On At anode
I
o
Permanent-magnet 0 Electromagnet
I
4
Screen
0
"- - 0 E .0)
i
(a) Electromagnet th rustor.
c
P
!E .-
5
z
S
.0
E m c
-I
Y
0
2 4 6 8 10 Axial position along anode centerline, cm
Propellant utilization efficiency,
vu,
.9 percent
Figure 12 - Performance of flight-type thrustor discharge chambers. Net accelerating voltage, 4ooo volts; accelerator voltage. -lo00 volts; discharge voltage, 30 volts; neutral propellant flw, 0.050 ampere; magnetic field as s h w n in figure L
(b) Permanent-magnet thrustor. Figure 11. -Magnetic field strengths of electromagnet and permanent-magnet th rustors.
Propellant utilization efficiency, percent
10 -
E, L 0)
c 0)
.-E
8.
'c)
L
0 c
m,
0)
6
0)
U
m L
0
4
2
0
.5
I
1.0 Thrust, mlb
I
1.5
I
2.0
Figure 13. - Calculated thrustor accelerator grid diameter for 13, MW)-hour lifetime. Impingement included from both direct and charge-xchange ions. Molybdenum accelerator grid had 0.32centimeter holes on 0.64-centimeter equalatera1 triangular spacing and was 0.16-centimeter thick.
NASA-CLEVELAND. OHIO E-3249
