Inhouse magnetoplasmadynamic (MPD) thruster technology is discussed. The study focussed on steady state thrusters at powers of less than 1 MW. Performance measurement and diagnostics technologies were developed for high power thrusters. Also developed was a MPD computer code. The stated goals of the program are to establish: performance and life limitation; influence of applied fields; propellant effects; and scaling laws. The presentation is mostly through graphs and charts

Myers, Roger M.

English

NASA Technical Reports Server

Appendix G N 9 2 -1 0 fj_'_-_"
--_r._.a_ j
L,_wrj SPACE PROPULSION TECHNOLOGY DIVISION N/_A
Low,s Rosoarch Cortler
MPD THRUSTER TECHNOLOGY
ROGER M MYERS
SVERDRUP TECHNOLOGY
NASA LEWIS RESEARCH CENTER
MAY 16, 1991
L_ SPACE PROPULSION TECHNOLOGY DIVISION
I III
IN-HOUSE PROGRAM OVERVIEW
N_.qA
Lew_ Rosliuch _nlor
• RE-ESTABLISHED IN 1987
• FOCUSSED ON STEADY-STATE THRUSTERS AT POWERS < 1 MW
• DEVELOPED PERFORMANCE MEASUREMENT AND DIAGNOSTICS
TECHNOLOGIES FOR HIGH POWER THRUSTERS
• DEVELOPING MHD CODE
• GOALS ARE TO ESTABLISH
PERFORMANCE AND LIFE LIMITATIONS
INFLUENCE OF APPLIED FIELDS
PROPELLANT EFFECTS
SCALING LAWS
G-1
https://ntrs.nasa.gov/search.jsp?R=19920000830 2020-03-17T14:50:14+00:00Z
SPACE PROPULSION TECHNOLOGY DIVISION
MPD Thruster Technology
Ikll,mll lqmml_¢il ¢,l_de+
High Power MPD Thruster Test Stand
Power
• 0.39 MW
Thrust stand Vacuum facility
• 0.1 to 4 N • 0.1 gls at 3x10 -4 TORR
Data/control 220 kW thruster
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MPD THRUSTER TEST STAND
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L,_rJ SPACE PROPULSION TECHNOLOGY DIVISION
HIGH POWER ELECTRIC PROPULSION (MPD)
N/ A
Lewis f_¢_ _*r
DEMONSTRATED MPD THRUSTER POWER
300"
POWER 200"
(KW)
100:
°
FACILITY
<
(1 YEAR)
7/87 9/89 8/90
220
10/90 11190 12190
DEMONSTRATED MPD THRUSTER POWER
INCREASING RAPIDLY
.-_ .......
,TECHNOLOG
".i'_.,'_'_.... MPD Thruster Technology .. ": "'"" ,-J_':_.:._!_
...... =:i'.'.::_i _:"........ Effe ::.. i!:,:_T_ruster Scaling and Materials .....,
• 'r._,;.'_._tt ) "_.: •
_'!'- .:,_
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0°
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:_,.,
diameter anodes both 3 and 6 Inches long.! _
.:.. " ".,_: 4"
_meter cathodes ..... -- .;i,:_.:
D0;impregnated tungsten cathodes ;",,-":.
• ':.:i;
G-3
L_rj SPACE PROPULSION TECHNOLOGY DIVISION
MPD Thruster Technology
/JmofBdor Ncmac oor 41_¢1NM_ rE
ii
Retouch _let
Performance Measured With Hydrogen and Argon
.2
Efficiency .1
O Hydrogen (_
I-1 Argon 0
0
@
0
..--Argon unstable
[3"" above this Isp
2"D, 3"L Anode 25mg/s flow rate
750 A discharge current
o I I I
1ooo 2000 aooo 4000
Speclfic Impulse, sec
Performance dramatically improved with hydrogen
• Efficiency increased by 2X
• Isp Increased by 50%
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L_,T_ SPACE PROPULSION TECHNOLOGY DIVISION NI_A
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MPD Thruster Technology
.3--
.2
Efficiency
.1
Thruster Performance
Geometry and Applied Field Effects
Jd = 1000 A, i_1= 0.1 g/s argon
O 2 inch diameter anode
I-I 3 inch diameter anode
ZS4 Inch diameter anode 1800
I
0 .15
O
-- _ O
°°
I I
.05 .10
O
Specific 1400
impulse,
s 1000
I 600
.20 0 .05 .I0 .15 .20
Applied Magnetic Field, T
• Efficiency increases with applied field strength
• Specific impulse increases with both anode radius and
applied field strength
CD-|1-$40_3
G-4
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SPACE PROPULSION TECHNOLOG Y DIVISION
I
MPD Thruster Technology
Lewle Heee_ch Genle_r
Anode Power Deposition
Applied Field and Geometry Effects
1.0 m
.9
Anode .8 --
power
fraction .7 --
.6
.5
.00
O 2 in. diameter anode
D 3 in. diameter anode
A 4 in. diameter anode
8
O
.05
O
O
I_D 0
I I
.10 .15
Applied magnetic field, T
.20
I Increasing applied field strength and anode Idiameter decrease anode power fraction
CD.II1 *SU2S
SPACE PROPULSION TECHNOLOGY DIVISION
HIGH POWER ELECTRIC PROPULSION (MPD)
iaw_ _ ¢emee
MPD THRUSTER HIGH CURRENT
HOLLOW CATHODE TECHNOLOGY
Higharea emitter Lowarea emitter
Three hollow cathode assemblies fabricated
and prepared for evaluation
¢1_1_-|1 I:ll
G-5
Z_TJ SPACE PROPULSION TECHNOLOGY DIVISION
MPD Thruster Technology
J¢11 _lcm_¢c_f •llna_lWf
I
Level• Rea4mrch ¢4mler
I
IScaling Issues I
• Megawatt class operation required for missions of interest
• Cannot operate megawatt class steady-state in current facilities
• Must be able to correlate MW class pulsed thruster operation and
steady state data •
• Data must enable rational extrapolation to high power levels
How do we realistically study MPD thruster performance and life iusing curr ntly available facilities?
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L_TJ SPACE PROPULSION TECHNOLOGY DIVISION
JWmN IICI_ID_C_V emlcR_,irl Llnwlu Re•earth Center
III I
MPD Thruster Technology
" (Diagnostics I
• X-Y probe positioning stand
- Electrostatic probes
- enclosed current contours
- Axial applied B field distribution
• Plume imaging
- Correlate ion density distribution with applied field
• Spectoscopy
- Non-invasive temperature and density measurements
CO-gl-SU211
G-6
L TJ SPACE PROPULSION TECHNOLOGY DIVISION
II |
MPD Thruster Modeling
N/ A
L_II RelNmt¢h C4mleT
IProgram Outline I
_onservation of mass -->density (p)
Conservation of momentum _ velocity (Vr ,V0 ,Vz)
Conservation of energy _ temperature (3")
t=quatlon of state _ pressure (P)
Evaluate transport coeffs, hall parameters, etc...
Fluid loop
f0hm's law and maxwell equations -.., Induced fields (B0)_
Field loop Maxwell (Ampere's) equation _ current density (j) JOhm's law _ electric field (E _ plasma potential)evaluate energy source, sink terms, etc...
Convergence on exhaust velocity V new _<0.01 V o_d
OX eX w
plasma potential: (bnow < 0.01d_ ola)
No Yes-_-_Evaluate thrust, specific impulse, efficiency,..."_
I Write to data files J
_Done
C0-91-S4|37
L_.TJ SPACE PROPULSION TECHNOLOGY DIVISION _JJf_A
L/NO|AJC'E f/¢N_tOOf IIl_ f_lff _IW|I _tllirch Conter
MPD Thruster Modeling
Comparison Wit h U. Stuttgart Model/Experiment
(6kA, 6 g/s)
Stuttgart-experiment Stuttgart-model NASA LeRC-model
I Current fractions into anode segments I
Segment 1 : 46% 44% 51%
Segment 2: 27% 27% 22%
Segment 3: 27% 29% 27%
NASA LeRC code in agreement with Stuttgart MPDT experiment/model I
CD-gI.S4P3_
G-7
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SPA CE PROPULSION TECHNOLOG Y DIVISION
MPD Thruster Modeling
Lewll floselrch Cenier
Comparison with Princeton University
Half-Scale Benchmark Thruster
6.0
3.0
0
ENCLOSEOCURRENTCONTOURS(HE^SURE,)
tz.__, 1.5g/.. qU^SI-STE^DYOeER^TIO,
nfl
?n
r, FI
_ S_
IO
0
Thrust Characteristics
/
& o 3 ¢ls J
C] EXPT (, /
• con_.: ,-:Y
T - 0.2 j2 (u..° R.g.._..,o.) _p_
-- 0 12 If, 2
CURn[NT (k^)
ENCLOSED CURRENT CONTOURS (PREDTCTE_)
12.4 k.A, 1.5 |/s. STEADY-STATE OPEI_TXON
dllOIBACl ilaemo.ooe lealcro4JyF
SPACE PROPULSION TECHNOLOGY DIVISION
MPD Thruster Modoling
Lowls Ro|oorch Cinto_'
Comparison with Princeton University
Half-Scale Flared Anode Thruster
0 5 I0 cm
ENCLOSED CURRENT CONTOURS (MEASURED)
7.9 k.A, 3 |11, QUASI-STEADY OPERATION
OZ
0 5 lO cm
ENCLOSED CURRENT CONTOURS (PREDICTED)
7.9 kA. _ G/s. STEA_Y*STATE OPeRATiON
z
Inn
Ùn
cn
4R
2n
Thrust Characteristics
1"36 = 3 BI= (_XPT) /
• 4 - 3 el= (co.r.) /
A& - 1.5 =/8 (EXPT)
x_- l.sc/, (coRr.) /__
---T - o.,s_2 (,-,.o,. R....,) ../.,.._--
_T - 0.120 J2(Datn Recressina) /_n _}_
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CunnrNl (k,',)
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SPACE PROPULSION TECHNOLOGY DIVISION
MPD Thruster Modeling
lU/_A
I.ewle ReNmrch C..ence_,
IStatus I
• Self-field version of MPDT code operational
- Modest execution times 3-5 hours VAX'CPU)
- General agreement with experimental results
- Thruster performance evaluations underway
• Applied-field version of code under development
- Routines for applied-B distributions incorporated
- Preliminary testing/modification in progress
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L_'13 sPACEPROPULSIONTECHNOLOGY_IVlSgONN/ISA
Ai'WdO_,K:_ I(_.= _,l O¢;• _,_¢ r0n,= _ Lewis Ret, ewch Center
KEY TECHNICAL ISSUES
G-9
°L_ SPACE PROPULSION TECHNOLOGY DIVISION N_A
KEY SCALING ISSUES
TWO PRIMARY CONCERNS
POWER LEVEL SCALING
- QUASI-STEADY VS. STEADY STATE
ISSUES MUST BE ADDRESSED USING
THEORETICAL MODELS TO ESTABLISH TRENDS AND
DEPENDENCIES
HGIH FIDELITY PERFORMANCE MEASUREMENTS
DETAILED DIAGNOSTICS OF PLASMA AND ELECTRODE PROCESSES
USED TO:
A. ESTABLISH FUNDAMENTAL RELATIONSHIPS
B. VERIFY MODELS
Lh_,_ SPACE PROPULSION TECHNOLOGY DIVISION NI_A
LOW0S ROso&fch Conlor
L I I| I IIII
PERFORMANCE EXPECTATIONS:
MUST EVALUATE EFFECTS OF :
PROPELLANT AND APPLIED FIELD
ELECTRODE SIZE AND SHAPE
PROPELLANT INJECTION
RELATION BETWEEN QUASI-STEADY AND STEADY-STATE:
MUST ESTABLISH DATA BASE WITH CORRECT PROPELLANT IN THE
APPROPRIATE OPERATING RANGE (j2/rh?)
MUST MEASURE PERFORMANCE, CURRENT DISTRIBUTIONS, PLASMA
AND ELECTRODE PARAMETERS
G-IO
_1 +l_'o_oGv_C_tq
SPACE PROPULSION TECHNOLOGY DIVISION
I
PERFORMANCE EXPECTATIONS
NASA
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• Sovey, J. and Mantenieks, M. "Pertormance and Lilelime Assessment ol Magnetoplasmadynamic Arc
Thruster Technology", J. Propulsion and Power, Vol.7, No. 1, Jan-Feb 1991
L'_'I_ SPACE PROPULSION TECHNOLOGY DIVISION
II _IOP,+CI i i _.,I_i _ r Olll ¢ I¢lll II
FACILITY REQUIREMENTS
NASA
Lewis Rlselach _OI'
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L_,'rJ SPACE PROPULSION TECHNOLOGY DIVISION
I I
FACILITY REQUIREMENTS
N/ A
Lmm_ch C4mer
EFFICIENCY CHANGE IN Van AND Vd
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by Saber with self-field thrusters
4" D, 3"L ANODE, 0.1 G/S ARGON, 1500 A DISCHARGE, Bz = .1 T
SPACE PROPULSION TECHNOLOGY DIVISION NASA
Le_wi Rese_ch COntO,
POTENTIAL MPDT FACILITIES
THRUSTER POWER, MW OPERATION
FY FACILITY H2 AR TIME, HR
PRESENT LERC T5,T6 0.1 (DEM) 0.2_t (DEM) CONT
1992 LERC T5 0.7-1 1 1 - 2
1993 LERC T5 1 - 1.5 2 4 - 6
1995 LERC T6 1 - 1.5 2 'CONF.'
1995 LLNL MFTF 1-5 'CONT.'
1998 LERC T6 1 - 5 1 - 5 'CONT.'
ESTIMATED
COST,$K
250 K
400 K
3500 - 5000
5000 - 7000
TBD
G-12
L_11_ SPACE PROPULSION TECHNOLOGY DIVISION
i i
MATERIAL LIMITATIONS
N/t /t
Lev, ns Res4.uch ¢erder
ANODE:
MEASURED HEAT FLUX AT HIGH POWER > 5 KW/CM 2 *
- LITHIUM HEAT PIPES LIMITED TO < 0.5 KW/CM 2
- OPTIMIZED BEAM DUMP (Cu) LIMITED TO ~ 5 KW/CM 2
SSME THROAT HEAT FLUX ~ 16 KW/CM 2 (relevance?)
CATHODE:
• CURRENT DENSITIES AT HIGH POWER > 100 A/CM 2.
LONG LIFE CATHODES LIMITED TO CURRENT
DENSITIES _<20 NCM 2 (LOW W F TWT CATHODES)
INSULATORS:
• KNOWN TO FAIL AFTER PROLONGED EXPOSURE TO UV AND
HIGH TEMPERATURE
• WE MUST SELECT GEOMETRIES WHERE PERFORMANCE AND
ENGINEERING LIMITS CAN BE EVALUATED
• PRINCETON UNIVERSITY
L '113 SPACEPROPULSIONTECHNOLOGYDmS,ONN/ A
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FACILITY LIMITATIONS:
MUST MEASURE PERFORMANCE AT PRESSURES < 5 X 10 -4 T
FACILITY PRESSURE HAS LARGE EFFECT ON ANODE HEAT XFER,
NOT CLEAR ON CATHODE
THRUSTER VIABILITY:
SHOULD FOCUS ON DEVICES WHICH MATCH ENGINEERING LIMITS
FOR:
ANODE HEAT TRANSFER
CATHODE CURRENT DENSITY
INSULATOR LIMITS
G-13


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