The activities on the development of the high power arc jet HIPARC, the thrust balance, and plasma diagnostic probes are discussed. Modifications of the HIPARC design and a synopsis of the materials used are given. Further experimental results with the TT30 thruster in the 50 kW range are presented. Some first calibration measurements of the thrust balance are also included. Progress concerning the development of plasma diagnostic devices is documented

Auweter-Kurtz, M.

Glocker, B.

Goelz, T. M.

Habiger, H.

Kurtz, H. L.

Schrade, H. O.

Wegmann, T.

English

NASA Technical Reports Server

High Power Arcjet 
M. Auweter-Kurtz, B. Glocker, T. M. Golz, 
H. Habiger, H. L. Kurtz, H. 0. Schrade, T. Wegmann 
Institut fiir Raumfahrtsysteme 
University of Stuttgart 
IRS 90-P7 
Second Progress Report 
NASA Grant No. NAGW-1736 
for NASA Lewis Research Center, Cleveland, OH 
August 1990 
https://ntrs.nasa.gov/search.jsp?R=19900019342 2020-03-19T21:51:36+00:00Z
Contents 
1 High Power Arcjet Thruster 4 
1.1 Modifications . . . . . . . . . . , . . . . . . . . . . . . . . . . . . . 4 
1.2 Materials Used . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 
2 TT30 - Further Experimental Results 8 
3 Test Facilities 12 
3.1 Thrust Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 
3.2 Vacuum Tank Preparation . . . . . . . . . . . . . . . . . . . . . . . 14 
3.3 Gas Supply System . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 
3.4 Data Acquisition System . . . . . . . . . . . . . . . . . . . . . . . . 14 
4 Fabry-Perot-Interferometry 15 
5 Plasmadiagnostics with Electrostatic Probes at the IRS 19 
5.1 Measurements of Electron Temperature and Density with Triple-Probes 19 
5.2 Measurement of Plasma Velocities with Time-of-Flight Probes . . . 22 
6 Optical Diagnostics 23 
2 
Abstract 
This second progress report covers the activities on the development of 
the high power arcjet HIPARC, the thrust balance, and plasma diagnostic 
probes from November 1989 to July 1990. 
Modifications of the HIPARC design and a synopsis of the materials used 
are given. Further experimental results with the TT30 thruster in the 50 kW 
range are presented. 
Some first calibration measurements of the thrust balance are also inclu- 
ded. Progress concerning the development of plasma diagnostic devices is 
documented. 
3 
1 High Power Arcjet Thruster 
1.1 Modifications 
The final High Power Arcjet (HIPARC) design differs in several points from the 
design presented in the first progress report: 
1. Due to your proposal (Fax dated March, 14th 1990) the nozzle contour was 
changed from bell to conical shape, with a 20 degree half-angle. This results 
in a new A,/At ratio of 114, which is greater than it had been before (100). 
2. Propellant was changed from ammonia to hydrogen. 
3. The maximum mass flow rate has been reduced from 1 g/s to 0.5 g/s. 
4. Due to experimental results with the TT30 thruster the throat diameter was 
reduced to 6 mm, which is still supposed to be rather big. 
5 .  In order to perform current distribution measurements, all segments are now 
supplied with power individually, each having in-line a shunt resistor for cur- 
rent measurements. 
6. A pressure tap has been included to determine the chamber pressure. 
7. For easy variation of the propellant injection the injector consists of four 
molybdenum inserts, which are easy to replace in order to check different 
injection angles and bore hole diameters. 
Fig. 1 shows a cross sectional drawing of the HIPARC thruster. In Picture 2 the 
thruster can be seen prior to integration into the vacuum chamber. 
1.2 Materials Used 
All electrical conducting parts as the cathode tube and the nozzle segments are made 
of copper. The cathode itself is made of 2 %-thoriated tungsten. The arc cham- 
ber consists of steatite, containing four molybdenum inserts serving as propellant 
injectors. 
The cathode positioning device is based on an aluminum part which is beared on 
a polyvinyl chloride (PVC) traverse. The cathode tube housing, the segment base 
plate and the gasket elements between the nozzle segments are also made of PVC. 
PTFE (polytetraflouroethylene) is the material of which the socket shells are made. 
In Fig. 3 the materials can be distinguished by different colours. 
4 
I 
+ 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
? 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
L 
4 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
?- 
Figure 1: HIPARC Cross Sectional Drawing ORIGINAL PAGE IS 
OF POOR QUALITY 
5 
ORIGINAL PAGE 
COLOR P3nOTOGRAPi-J 
- ~~ - 
Picture 2: HIPARC Thruster 
6 
1 00'0rE 
I I 
0 
(0 
M 
=I= 
I 
I 
I 
I - 
b 
I - 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
L 
m 
0 
.c - 
8 
/ 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I- 
- 
Figure 3: Materials: C 
ple), aluminum (brown 
(blue), PTFE (light blue), other matedds (black) 
7 
. - . - . - . - . -  
I 
. - . - . - . - .  
2 TT30 - Further Experimental Results 
170- 
150- 
n > 130- 
U 
0 
D w 110- 
0 > 
+- - 
90- 
70 - 
50 
As the experiments described in the first progress report were limited to a 400 A 
current by the facility, the facility has been rebuilt, so that the current limit is 
now at 1000 A. With this rebuilt facility tests were performed with hydrogen as 
propellant and mass flow rates of 0.1 g/s, 0.2 g/s, and 0.3 g/s, using the 5 mm 
throat diameter and the 2 mm cathode gap. In the case of the two lower mass flow 
rates the thruster was not operated to its limits, whereas with 0.3 g/s 430 A was 
shown to be the maximum current for the cathode of 4 mm diameter. With a mass 
flow rate of 0.1 g/s, a specific impulse of more than 750 s can be reached, whereas 
with 0.3 g/s it is limited due to the small cathode to about 600 s. The efficiencies 
are rather low, between 10 % and 16 %, increasing with higher mass flow rates and 
decreasing with higher power. With a propellant flow of 0.3 g/s more than 50 kW 
could be reached without any damage of the nozzle segments (Fig. 6). The arc 
chamber pressure was never higher than 900 mbar (Fig. 6)) and the heat flux was 
moderate (Fig. 8). The major part of the heat flux was deposited into the first two 
segments downstream of the nozzle throat which served as anodes (Fig. a), whereas 
the heat load of the nozzle throat was always less than 0.37 kW/cm2. 
propellant: hydrogen m--m w i t h  0.1 g/s 
w w i t h  0.2 g/s 
rn w i t h  0.3 g/s 
I I I I I I I 
0 
Current [A]  
Figure 4: Current - Voltage Characteristic 
These results indicate that specific impulse and efficiency can be improved at  higher 
power levels using a thicker cathode. 
8 
1000 
loo]  propellant: hydrogen with  0.3 g / s  
o w i t h  0.2 g/s 
m with 0.1 g / s  01 I I , I I I I 'V 
50 100 150 200 250 300 350 400 450 
Current [A] 
Figure 5: Current - Arc Chamber Pressure Characteristic 
2.0 
1.6  
1.2 
0.8 
0.4 
propellant: hydrogen 
EB with  0.3 g/: 
o w i t h  0.2 g/s 
* w i t h  0 1 g/: 
5 15 25 35 45 I 
Power [ k W ]  
Figure 6: Power -Thrust Characteristic 
9 
10 
9 
8 
7 
6 
5 
4 
3 
2. 
1 
0- 
3.0- 
2.5- 
2.0- 
1.5- 
5 
El. P o w e r :  2 0 K W  
propellant: hydrogen  
,'p W H 
W 
W 
H 
H 
H w i t h  0.3 g/s 
0 w i t h  0.2 g/s 
m w i t h  0.1 g/s 
I I I I  
15 25 35 45 
Power [ k W ]  
Figure 7: Power - Heat Flux Characteristic 
L - 3 
3.5 I I I I 
fZS3 El. Power: 4 0 K W  
1 0  
0.5 
0.0 
I I 
propel la n t .  hyd -ogen 
masstlow: 0.3 g /s  
5 6 7 8 
Thruster Axis  [cm]  
Figure 8: Heat Flux Distribution 
10 
1 0 0 7  I ! I I I 
90- 
80 - 
70- 
60 - 
50- 
40 - 
30-  
20- 
10- 
Ea 
I I I 
propellant: hydrogen 01 
5 15 2 5  35 45 55 
Power [kW] 
Figure 9: Thermal Efficiency vs. Input Power 
II 15 Ea m 
0 
5 1 
m 
0 0  
mm 
I 
m 
m f m m  
w i t h  0 .3  g/  
0 with 0.2 g /  
x w i t h  0.1 a /  
1 propellant: hydrogen 
I 
01 J I  I I I I 
300 400 500 600 700 8( 
Spec. Impulse [ s ]  
Figure 10: Efficiency vs. Specific Impulse 
11 
- 1 2 -  > - 1 . 1  - E 
0 1.0- 
> 0.9- 
0.8- 
a 
m 
+ -
0 
0.7- 
0.6 
Time [ s ]  
- 
- 
-. 
~ 
- 
- 
- 
- 
H waag2l.sdf 
I I I I 
Figure 11: Voltage vs. Time During Heat-Up and Cool-Down 
3 Test Facilities 
3.1 Thrust Balance 
The thrust balance has been mounted and integrated in the vacuum chamber. First 
calibration measurements have been carried out without the thruster being moun- 
ted. The influence of temperature was simulated by heating critical parts by a 
hot air blower. During the heat-up and cool-down process balance data was re- 
corded with constant weight on the balance. Fig. 11 shows the voltage vs. time 
characteristics for this case. 
Several datasets have been recorded to determine performance characteristics, each 
set beginning with zero weight, increasing weight successively up to M 600 g, and 
subsequently decreasing weight successively down to zero. Different initial stress 
values resulting in different zero-weight voltages have been established. As Fig. 12 
shows, linear characteristics can be achieved by reasonable initial stress only. 
12 
1.8 I I I 1 ! 1 I 
1 . 6 4  
1 .4  
1 . 2  
1 . o  
0.8 
0.6 
04 
0.2 
0.0 
+--e woog07.sdf 
I++ waag06.sdf 
waag04.sdf 
0 100 200 300 400 500 60C 700 
Weight [g] 
Figure 12: Voltage vs. weight characteristic at different initial stress values, showing 
one load/release-cycle each. 
13 
ORIGINAL PAGE 
COLOR PHOTOGRAPH 
Picture 13: Gas Cooler 
3.2 Vacuum Tank Preparation 
The thrust balance has been integrated into the vacuum tank. New feed-through 
power supply leads have been installed, because the old ones didn’t fit the new 
thrust balance. 
In order to avoid thermal overload of the vacuum system, a gas cooler has been 
built in just in front of the outlet (Fig. 13). 
3.3 Gas Supply System 
Since the gas supply system was limited to a maximum mass flow rate of 0.3 g/s 
hydrogen, a new mass flow controller with a range of 0-0.5 g/s hydrogen has been 
integrated. 
3.4 Data Acquisition System 
In order to perform current distribution measurements, three additional isolation 
amplifiers have been installed, each of them leading to our DEC PDP-11/73 data 
14 
acquisition computer via newly installed ADC channels on our Intercole Spectra 
data acquisition system. 
4 Fabry-Perot-Int erferornet ry 
With the FPI, it is planned to measure Doppler-shifts of spectral emission lines of 
an accelerated plasma and therefore the velocities of the emitting particles. The 
temperatures of the heavy particles can also be measured from the Doppler broade- 
ning of the emission lines. The technical concept of the FPI was already described 
in detail in the former progress report [l]. 
The experimental setup of the Fabry-Perot-Interferometry has now been build up. 
It is shown in pictures 15 and 16 and can be compared to figure 14. 
Picture 15 shows the optical rack on the left hand side and the data acquisition rack 
on the right. The data acquisition rack includes the computer and the the transient 
recorder on the two lower levels. The optical rack which is shown in more detail 
in picture 16, is divided into four levels. On the upper level, the reference light 
sources and collimating optics are mounted. The reference light passes through 
a hole down to the second levels, where the light collecting optics are mounted. 
The silver colored assembly on the left includes the collecting optics and the beam 
splitter. By several plane mirrors, the light is passed through the FPI which is on 
the right hand side. After the FPI an additional mirror reflects the light to the 
third level with the focussing lens, the narrow band filter and the photomultiplier. 
The fourth level includes the power supply for the photomultiplier and the ramp 
generator for the FPI. The whole rack is mounted on a marble block in order to 
suppress vibrations which will disturb the FPI calibration. At the moment only a 
slight effort is done to suppress external vibrations. 
At the moment the calibration of the optics without the mirror plates of the FPI 
is being performed. In a second step the mirror plates will be mounted and the 
effect of external vibrations on the setup will be examined. Depending on the 
results the most effective vibration isolation has to be installed in order to perform 
measurements. It is planned to start reliable operation of the FPI in spring 1991. 
15 
4 
\ 
I 
3 
b - 
\ ' E  
B 
I 
Figure 14: Schematic Drawing of the Fabry-Perot Interferometry 
16 
C!!RIGINAL PAGE 
COLOk' PHOTOGRAPH 
Picture 15: FPI: Apparatus 
17 
Picture 16: FPI: Optical Rack 
18 
5 Plasmadiagnostics with Electrostatic Probes 
at the IRS 
With means of electrostatic probes, we want to measure the distributions of the 
electron temperatures and electron densities in our plasma-jets, as well as the velo- 
cities of the plasma particles by time-of-flight measurements. The directions of the 
plasma-streamlines can also be detected with electrostatic probes by measurements 
of the ion-current at  different angles of attack. 
5.1 Measurements of Electron Temperature and Density 
with Trip le- Probes 
Electrostatic triple-probes are used to determine the values of the electron tempe- 
rature and the electron density instantaneously, within a short time on the order 
of the probe response time. Several triple probes of different dimensions are being 
built up. The technical concept of the probes was described in the former progress 
report [I]. 
First measurements of the electron temperature were performed with the DT2 and 
the ZT1 MPD-thrusters of the IRS. While the data acquisition works convenient, the 
xy-table which moves the triple probe through the plasma jet causes some problems. 
The maximum speed of the xy-table in radial position is about 4 cm/s. By typical 
jet diameters of 20 to 40 cm depending on the ambient pressure, the movement 
through the jet plume takes 5 to 10 seconds. Figure 18 shows a radial profile of the 
electron temperature of the DT2 plasma plume. The triple probe was moved from 
the left to the right. Despite the reproducable shape of T, at reasonable values in the 
center of the plasma plume, one can clearly see a small maximum of T, at a position 
of about -85 mm which does not occur symmetrically. Also there is an remarkable 
rise of T, in the outer regions of the j e t  after the probe was moved through. Such 
a behavior occured also when we moved the probe from the right to the left as one 
can see in Figure 19. The appearence of such a maximum also has been detected 
before the probe reaches the jet axis in the ZT1 thruster (see Figure 20). Besides 
the fact that the absolute values of T, are affected by different angles of attack in 
the outer regions of the plasma jet, it is supposed that the behavior of the probes 
is caused by heating effects. To avoid those heating effects, it is planned to build 
up a faster xy-table. The radial movement of the new table should be driven by a 
pendulum to reach speed of about 20 to 30 cm/s. The new table will be operational 
until November. 
Picture 17 shows the triple probe during operation in argon. 
19 
,ORIGINAL PAGE 
COLCF: PHOTOGRAPH 
2.80- 
2.60- 
2 2.40- 
- 
Picture 17: Triple-Probe Measurement 
I=2000A - 
m=0.8g/s A- 
p =0.02lmbor - 
- 
Y 
a, 2.20- 
t 2.00- 2 
+ 3 1.80- 
1.60- 
a 1.40- E 
Q, 
2 1.20-  
i. 1.00- 
0 
0 
W 
5 0.80- 
U 0.60- 
0.40- 
0.20- 
0.00 
X =  107mm - 
- 
- 
- 
- 
- 
- 
- 
! I O  I J J I I I J J I , , , , , , , , , , , , , l , , , l , , , , , , , ( , l , ( , , , , , ( ,  I , , , , , , I ,  
Figure 18: Radial Distribution of the Electron Temperature (I) 
20 
2.80- 
2.60- 
2 2.40- 
a, 2.20- 
n 
u 
I=2000A - 
m=0.8g/s. Ar 
p =O 02l rnbar  
X=  107rnn-I 
- 
- 
- 
21 
3.00 I r v  I I T s I  , 1 8 8  I I I 1 r t  n r I v n 1  c t r s  > I v s  ( I t I  1 8  I r t  
- 
- 
- 
- 
- 
; 1.00- 
0 
o 
w 
- 
0.80- 
2 0.60- 
- 
- 
0.40- - 
0.20- - 
0.00 ! I 1  l l l l l ~ , l l l l l l ~ l l l l l J J J I I I I I I I I I I J 1 , , , , , , , , , , ,  I , , , , , , , ,  
0 100 200 300 -300 -200 -100 
2.0 
1.8- - >, 1.6- 
Y 
i- 
a, 
a 1.4- 
5 1.2- 
b 1.0- 
u 
0 
a 
E 
0.8- 
0.6- 
2 0.4- 
0.2 -
0.0 
u 
c 
-I- o 
w 
-I 
-300 
ZT1 RUN1 31 PTP23/.5/5AP HTEl26.SPF(SDF) 
I I I J J J , , , , , , , , , , , , , l , , , , , , , , , , , , , , , , , , l  , , ( , , , ,  
I = 4000A 
m=Z.Og/s AR 
p =O.lrnbor 
X=41Omm 
- 
- 
- 
- 
- 
- 
- 
I ,  1 , , , , , 1 , , , , , , , , , , 1 , , , , , , , ( ,  , , , , , ( , , ,  I ( , ,  I " " " " '  -200 -100 
ORIGINAL PAGE 
COLOR P H O T Q G R A ~ ~  
Picture 21: Time-of-Flight Probe 
5.2 Measurement of Plasma Velocities with Time-of-Flight 
Probes 
The technical concept for time-of-flight measurements of natural plasma fluctuations 
was also described in the last progress report. 
A first student thesis on the time-of-flight measurement of natural fluctuations has 
been performed with the DT2 and the ZT1 MPD-thrusters. The major result of the 
thesis is, that the density of charged particles at low massflows and high currents 
seems to be very large at  the center axis of the plasma jet. So far the natural 
fluctuations are not clearly marked and difficult to detect. For the investigated 
thruster parameters, the method seems to allow only a rough estimation of the 
plasma speed. The next step of the time-of-flight measurement will be to generate 
fluctuations and to detect plasma streamlines for probe axis alignment. Picture 21 
shows the time-of-flight probe during operation. 
22 
6 0 p t ical Diagnostics 
During the designing and construction phase of the high power arcjet and its test 
stand, the qualification of the optical plasma diagnostics have been continued. For 
that purpose especially emission spectroscopic measurements were carried out both 
at  the arcjets of lower power and at  MPD-thrusters. First of all general spectra for 
different propellants and various operating conditions were taken. These are im- 
portant for the line identification and to determine the different plasma parameters 
from the line broadening and the Doppler shift of selected emission lines subse- 
quently. An Abel inversion technique is being installed for the determination of the 
plasma parameters at  different radial positions from the measured radial emission 
profiles. 
The measured general spectra at  the nozzle type magnetoplasmadynamic (MPD) 
thruster of the DT-series yielded to new results concerning the onset-phenomena, 
where various instabilities of the arc and the plasma flow appear which limit the 
high power levels of these devices. Figure 22 shows the voltage versus current curve 
of the MPD-thrusters running with argon as propellant. These curves rise steeply 
at  high power levels ( x  4200 A), which results in a decrease of the thrust efficiency. 
At a current level somewhat above this break in the voltage curve, oscillations of 
the plasma can be detected in low and high frequency ranges, Fig. 23, which can be 
measured with an electrical probe. Figures 24-27 show spectroscopic measurements, 
done at  the MPD-thruster DT2-IRS running with 0.8 g/s argon at different current 
levels, between 2700 A and 3500 A. These measurements show that already at 
current levels of about 3000 A,  spectral lines of the second ionized argon appear 
and so at  least in the core of the plasma plume the second ionization of the argon 
atoms starts. 
23 
- 
> 
u 
20- 
0 
1 0 0 ,  , . , . , , . , . , 
4 oscillation-onset 
mass flow: 0.8 g/s Ar (DT2: 10% anode gas) 
, ' , ' I '  
7 0 0 1 ,  , .  . , . . . , , . . . . , . . . . I  
600 j  DT2 
4 : Arlll-Lines 1 
other: Aril-Lines 
C;r:en? [ kA ] Wavelength [ n m  ] 
Figure 22: Voltage vs. current of the 
MPD-thrusters DT2 and DT5, running Figure 25: Intensity vs. Wavelength at 
with 0.8 g/s argon, arrows indicating the 2978 A 
onset of oszillations 
0.028 
0.019 
0.009 
0.000 
o,olo 1 0.8 g/s Ar ( 10% anode gas ) current 4962 A / . , . , . , , I .  
0 1 2 3 4 5 
other Arll-Lne 
current 3122 A 
320 325 330 335 340 
Frequency [ Megahertz ] 
Figure 23: Frequency Spectrum of the Figure 26: Intensity vs. Wavelength at  
Plasma Oscillations at the DT2-Thruster 3122 A 
Wovelength [ n m  ] 
current 2784 A 
320 325 330 335 340 
Wavelength [ n m  ] Wavelength [ n m  ] 
Figure 24: Intensity vs. Wavelength at Figure 27: Intensity vs. Wavelength at 
2784 A 3507 A 
24 
References 
[l] HIPARC, High Power Arcjet, M. Auweter-Kurtz et al., Institut fur Raumfahrt- 
systeme, University of Stuttgart, IRS 89-P19, First progress report, NASA Grant 
No. NAGW-1736, October 1990 
25 


High power arcjet

High power arcjet

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server