GETAWAY TETHER EXPERIMENT 

(GATE) 

A FREE FLYING TETHER EXPERIMENT 

{NASA-CR-1799 12) GETAWAY TETEER EXPEBIHENT ~3 7- i 34 55 [GATE): B FREE FLYING TETHEE EXEEBIMZNT 

Findl Report {Alafama Univ., EuntsviLle.) 

68 P CSCL 22B Unclas 63/18 43451 

A Final Report 

For 

Marshall Space Flight Center 

National Aeronautics and Space Administration 

Grant #NAG8-586 , 

BY 

Michael Graene , PH. D . Assistant Pmfessor 

University of Alabama in Huntsville 

Huntsville, Alabama 35899 

December 1986 
TABLE . 

OF CONTENTS 

TOPIC 

Objectives ....................... 

Mission Sequence .................... 

Electrodynamics .................... 

Induced potential ................. 

Equivalent Circuit ................. 

Charging Currents ................. 

Electro charging ................. 

Electro boost ................... 

Conclusions . ' ................... Deployment and Altitude Control ............. 

Coordinate System ................. 

Deployment ..................... 

wFlipwl 

Conclusions .................... 

Tension/Impact Dynamics ................ 

Command and Control System ............... 

Thermal Profiles 

Power ......................... 

Physical Layout 

Project Schedule 

Conclusions 

References ....................... 

...................... 

.................... 

..................... 

.................... 

...................... 

PAGE 

2 

6 

10 

11 

15 

21 

27 

29 

31 

34 

35 

37 

43 

45 

48 

50 

52 

54 

58 

62 

64 

65 
I/-- 

- 

\ 

1 
OBJECTIVES 

Orbital reboost and power generation using electrodynamic tethers has 

been suggested as a means of increasing the operational flexibility and 

orbital lifetime of satellites . Excess energy generated by solar arrays 

can be stored as orbital energy and later extracted from the orbit during 

peak power demands. The GATE experiment will demonstrate this practical 

tether application and will measure the dynamic circuit impedance. The 

small forces generated will scale to the small size of GATE so that observable orbital changes will occur. The dynamic impedance measurements complement those of the Tethered Satellite System (TSS) first flight and may help 

plan the second TSS electrodynamic mission. 

1 

Micrometeoroid hazards to space tethers have been the subject of uncertainity because prior measurements on flux was relative to flat plates 

and hypervelocity impact tests were performed on flat plates. Application 

of these data to tethers involves extrapolation in dimensions and in damage 

mechanics. A short series of tests performed by Martin Marietta AerospaceL 

in support of the TSS was performed on stainless steel and woven Kevlar 

under no tension. Extrapolation of the data to tensioned Kelvar is 

questionable. Additional tests are scheduled by MMA for the fall of 1986 to 

assess hypervelocity impacts on tensioned tethers. These tests, while 

addressing the damage mechanisms, do not answer the flux issues. A long 

duration exposure in space can provide this data. Shock waves generated by 

micrometeroid impacts on a tether are expected to have unique signatures. 

These waves and ensuing vibrations are detected, analyzed, and the resulting 

data is transmitted to ground. These data will also be of value in assessing 

damage to tension cables in large space structures and to electrical or data 

transmission lines. 

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The GATE capability to excite ultra low frequency radio waves is not 

equal to that of the TSS, but GATE can provide additional data. The efficiency of the short antenna compared to that of the long antenna (TSS) is of 

interest for communications purpose. A second radio objective, telemetry 

reception by radio amateurs, promotes worldwide interest in space. The 

flight was 

extremely successful in this regard. The mode of transmission is a subset 

of the packet radio protocol now in a period of explosive growth in 

interest. Low cost terminals are available and are in wide use in the 

United States and will be world wide. Using the GAS 007 experience, 

coverage can be expected throughout the Americas, Japan, Europe, and 

Australia. The data mode and power level will allow unattended reception on 

the ground. 

of Get Away Special 007 by the Alabama Space and Rocket Center3 

Near real time data is possible using amateur radio. 

Although much theory exists for calculating the radar cross section of 

long wires, the theory has ignored the effects of resistance. This introduces much error in grazing incidence calculations. Also, the effects of 

reflections in 

the open literature. The objective to measure these with the rendezvous 

radar is of interest to NASA in future Shuttle operations near space 

tethers. This is of interest in detecting and avoiding tethers. Also of 

interest is the ability to discriminate the satellites at the ends of the 

tether. Tests of the rendezvous radar at White Sands in 1985 for the TSS 

program4 showed negligible interference by the electrodynamic tether. However, tether straightness was an issue in the tests and the measurements 

were made only at angles expected in the TSS flight. Orbital flights appear 

to be the only way very long wires can be supported in radio frequency "free 

space" for radar measurements 

from long wires on radar processors has not been addressed 

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5 
MISSION SEQGENCE 

The accompanying figure show the mission sequence for GATE in general. 

After a 

spring later 

phases. The next phase is stabization using the control laws developed. 

Experimentation consists of each of the objectives. 

release from the GAS cannister the satellites will separate using 

induced initial velocity and then the gravity gradient during 

Different orientations are required for GATE for the charging experiments and for the electro-boost experiments. Using the larger satellite 

(mother) as an electron collector and the smaller (daughter) as an electron 

emitter, the mother must be in the higher orbit for charging and in the 

lower orbit for boosting. This necessitates 'flipping' the configuration 

between these two experiments. 

6 
GATE 

.- .. 

1, 

2, 

3, 

4, 

5, 

6, 

7, 

8, 

9, 

10 I 

MI SSI ON SEQUENCE 

I NIT1 AL DEPLOYMENT 

FLIP 8 RE-DEPLOYMENT AS REQUIRED 

ELECTROBOOST I (15 DAY)" 

FLIP & RE-DEPLOYMENT 

CHARGING EXPERIMENT I * 

(15 - 30 DAY) 

FLIP 8 RE-DEPLOYMENT 

ELECTROBOOST I I (15 DAY 1 " 

FLIP & RE-DEPLOYMENT 

CHARGING EXPERIMENT I I* 

(15 - 30 DAY) 

RE-ENTRY 

* ALSO RADAR CROSS SECTION, RADIO . PROPAGATION AND METEORITE IMPACT 

EXPERIMENTS 

7 
The detail mission sequence for a possible 90 day mission is given. 

After initial deployment, the orientation of the system will be sensed from 

the polarity of the induced voltage. If the mother is not in the down 

position, the system will be 'flipped' and redeployed. A 15 day electroboost experiment is envisioned that will increase. the orbit by 7- 10 %. The 

system will then be 're-flipped' and a charging experiment will ensue. 

After 're-flipped' 

again and a second boost experiment is planned. This will be followed by 

another 'flip' manuever and another charging experiment. 

a recharging period of up to 30 days the system will be 
(No Figure) 

9 
GET AWAY TETHER EX P ER 1 TV1 E:.! T 

(GATE) 

ELECTRODYNAMICS 
UBL for 28 degree Inclination 

,240,0 I I I I I I I 1 I I I I I I I I I I I 

OmO XI 0Qm0 16,00 Orbi t Number 

11 
- GATE ELECTRODYNAMICS 

The GATE, when fully deployed and stabilized in local vertical, is a 

conductor moving through the Earth's magnetic field, and a nonelectrostatic 

EMF is developed along the length of the tether due to the V x B * L effect. 

The length of the tether is 1000 m, its velocity will be about 7740 m/s, and 

a nominal value of magnetic field strength at 300 km altitude is 0.45 Gauss. 

Thus, when the tether orientation, velocity, and magnetic field are mutually 

perpendicular, the developed EMF has a maximum value of 350 V. However, due 

to orbital inclination and magnetic anomalies, this EMF is expected to vary 

as- much as 50%'. This is a plot of the induced potential, VBL, as a 

function of time for an orbit with a 28" inclination for the 1 km. tether of 

the GATE. 
METHODS OF PLASMA CONTACT 

o PASSIVE: PARTICLE COLLECTION 

ADV: NO MASS PENALTY 

NO ENERGY PENALTY 

LOWEST COST AND COMPLEXITY 

ADDITIONAL SURF ACE AREA INCREASES DRAG 

CURRENT DEPENDENT ON PARTICLE DENSITY 

DIS: CURRENT PROPORTIONAL TO SURFACE AREA 

o ACTIVE: ELECTRON EMISSION 

ADV:NO MASS PENALTY 

LOW COST: HIGH CURRENT POSSIBLE 

LESS DEPENDENCE ON PARTICLE DENSITY 

DIS: ENERGY PENALTY If INTERNAL POWER USED 

POSSIBLE SPACE CHARGE LIMITATION 

o AC1TVE:HOLLOW CATHODE PLASMA CONTACTOR 

ADV:LOW PLASMA CONTACT VOLTAGE DROP 

HIGH CURRENT POSSIBLE 

LITTLE DEPENDENCE ON PARTICLE DENSITY 

MASS PENALTY 

ENERGY PENALTY 

DIS: HIGHEST COST 

13 
Ts czuse =: current to flrw thil~iigh tile tether, a means musc be provided 

for electrical conduction through the ionospheric plasma. Plasma contact 

must occur at the mother or daughter subsatellites with an insulated tether. 

Three methods of plasma contact have been examined. A passive method of 

plasma contact is particle collection on the surfaces of the subsatellites. 

No internal energy or mass is expended, and this method is certainly lowest 

in cost and complexity. One of the disadvantages of plasma contact made 

solely through particle collection is the magnitude of current realizable. 

Electrons may be trapped on the surface of the subsatellites in relatively 

large to 

microamps . Larger current values are possible with larger collecting 

surface 

quantities yet the transfer of positive charge limits the current 

10 

but provision must be made to limit aerodynamic drag. 

An alternative to collection of positively charged ions is the emission 

of electrons. Thermionic emission is a proven technology and has no mass 

penalty. A high current is possible with such a configuration, but there is 

an energy penalty for heating the filament. Either internal electrical 

energy must be expended, or if solar energy may be harnessed, provision must 

be Previous experiments 

with electron emitting spacecraft such as ATS-5 have shown that for this 

method of plasma contact to be effective, the charging of the satellite 

11 surface must be taken into account. . 

made to direct the collecting surface at the Sun. 

Perhaps the most effective method of plasma contact is the hollow 

cathode . The hollow cathode creates its own plasma and clamps the potential of the subsatellite to that of the surrounding environment. High 

currents are possible, and dependence on ambient ionization lex7el is minimized. The disadvantages to the hollow cathode are three-fold: higher cost, 

mass penalty, and energy penalty for heating the ionization chamber. 

12 
GATE ELECTRODYNAMIC 

EQUIVALENT CIRCUIT 

I vc 

EMF 

I 7 

I 

IE 

' VE 

RT = TETHER RESISTANCE + LOAD RESISTANCE 

EMF = (V X B L)+ APPLIED VOLTAGE 
n--- -1 ---- -...-&--A I--- L --_- WLbbC pAaauia LuiiLaLL ~ma ueeti established, the GAiE wiii execute two 

electrodynamic experiments: electroboost and system recharge. The simplified equivalent circuit of the tether-plasma system has a series resistance, 

a source of EMF, the nonlinear plasma contact relationships, and a circulating current. This circuit can be used for either the Alven generator or 

Alven motor configuration simply by changing the EMF source. In the generator configuration, the EMF source is only the induced VBL. In the engine 

configuration, the EMF source is Vapplied - VBL. 

For the GATE to be configured as an Alfven generator, the series 

resistance would be the sum of tether resistance, a load impedance, and the 

impedance of the plasma between the two contactors. * 

L and the power extracted from the orbit is the product of the current 

flowing in the tether and the induced EMF. The product I* RL represents the 

The EMF is just v x B 

conversion of part of the orbital energy into useful electrical energy. 

The other configuration for the GATE is the Alfven Engine. Stored 

electrical energy is expended to increase orbital energy and, thus, increase 

the orbital altitude of the GATE. In this case, the total EMF of the system 

is the sum of V x B . L and an applied voltage inside the mother subsatellite. The polarity of the applied voltage must be opposite to the induced 

EMF and its magnitude must be greater for electroboost to be successful. 

The series resistance of the applied voltage source replaces the load 

impedance in the electroboost configuration and there power gained by the 

orbit would be the power expended by the voltage source minus the losses in 

the tether and the plasma contact. 

16 
GATE.ELECTRODYNAMICS 

o EMF DEVELOPED ACROSS LENGTH OF TETHER = C x E * Tt 

o GATE VELOCITY = 7740 rn/s AT 300 km ALT. 

o MAGNETIC FIELD MAGNITUDE APPROX. 0.45 GAUSS 

o ORBITAL INCLINATION AND MAGNETIC ANOMALIES VARY 

HORIZONTAL MAGNETIC COMPONENT BY MORE THAN 25% 

o PLASMA CONTACT AT ENDS OF TETHER: 

CIRCUIT COMPLETION THROUGH PLASMA NOT WELL DEFINED 

o EFFICIENCY OF PLASMA CIRCUIT CONTACT DEPENDS ON 

IONIZATION LEVEL AND PARTICLE TEMPERATURE: 

BOTH VARY WIDELY FROM DAY TO NIGHT 
It must be stressed that the circuit completion through the plasma is 

, still not well defined13. Many studies on the subjects of particle collecI 

tion and plasma impedance l3-I8 have been done in the last two decades but 

little experimental data has been collected. It is known that the magnitude 

of the current in the GATE system will be a nonlinear function of: Ionization level of the plasma; Thermal energy of the ions and electrons; Magnetic 

field strength and orientation; Tether velocity and orientation; Total 

series impedance of the tether-plasma circuit; and, Plasma contactor 

potential and temperature. The ionization level of the plasma and thermal 

I 

energy of ions and electrons in the ionosphere vary from day to night and 

with level of solar activity. It has been hypothesized that the plasma 

impedance between the contactors will be very small: on the order of one 

ohml8. 

18 
V 2 0: J = K T2 exp[ -(4+eV)/kT] 

V < 0: J = K T2 exp[ (-4+Je3E/4xc)/kT] 

I E -0.00589 + 0.03884 EXP(-0.07541 V ) 

( Source: experimental data from NASA contract TM-88850, 

M. Patterson and P. Wilbur ) 

For negative collector potential: 

1 - A [Ji + Je exP(-INepI)I 

For positive collector potential: 

10 A [Je + Ji ex~(-IN,pl)I 

Dobrowolny - Janve Model: 

I 0.6410 ( (lelvep/k Te) (Ld/ 

Parker - Murphy Model: 

I - I~(I + 2 Jvep / v0 ) 

)4/3 ) 6/7 

= potential of eletrode with respect to plasma 

= dimension of collecting electrode 

- surface area of collecting electrode 

veP where : 

a 

A 

Te - electron temperature 

k = Boltzmann's constant 

Ne - electron density of plasma 

Vthe = electron ther a1 velocity = 42 k Te / me 

K - 120 ampere/cm'oK2 

T - filament temperature, "Kelvin 

4 = emitter work function 

V - potential of filament 

e - electron charge, 1.602 x lo-'' C 

e - permittivity of region around filament 

' E - electr'c field at filament (V / rfil) 

Vo = 

3 178 a (B / 0.45) 

Nep IelVe /k Te 

Ji - 0.5 1Vi Zi le1 v / A 

J, = 0.25 Ne e Vthe 

19 
For the purposes of the missions of the GATE, two models of particle 

collection will be considered. The first by Parker and Murphy12 will be 

considered as the worst case (lowest current) and a model based -on 

Dobrowolny and Janve13 will be considered as the best case. The use of 

these two models will bracket the performance of the GATE. 

In addition, both the use of thermionic emitters and hollow cathode 

devices was considered. The non-linear relationships for these two devices 

are given. The relationships for the passive collection models are also 

given. 

To illustrate the electrodynamic capabilities of the GATE, next 4 pages 

show the results of a digital simulation of the tether-plasma circuit in the 

Alfven generator configuration while the fifth page shows the results for 

the Alfven motor configuration. Passive electron collection is assumed at 

the anode, or mother subsatellite, and active thermionic emission is assumed 

at the daughter subsatellite, the cathode of the system. For simplicity, 

the anode is assumed to be a metallic sphere. The tether resistance is 

fixed at 100 n and the total circuit ohmic resistance at 1000 n. In each 

case, an ion density of 2 x 10 "/m3 has been assumed, although densities up 

to 20 x 1O1'/meter3 have been recorded 13-18 
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It is interesting to compare the simulation results using a hollow 

cathode model versus a thermionic emitter model. This is done in the 

figures on the following three pages. With the same electron collection 

model, there are virtually no differences in the current voltage relationships for the GATE, the power - voltage relationships, or the power - 

resistance relationships. The only notable difference occurs for resistance 

of less that 100 0 with the Dobrowolny-Janve collection model. These data 

imply that the current in the GATE is strictly limited by the passive 

collection of electrons by the mother satellite. 

All of the accompanying figures are plots of tether current versus 

induced voltage. (a) uses the Dobrowolny-Janve model with a thermionic 

emitter while (c) uses the same model with a hollow cathode. (b) uses the 

Parker-Murphy model with a thermionic emitter while (d) uses the hollow 

cathode device. It is interesting to note that when the same collection 

models are compared (a-c or the results are virtually identical with 

both a thermionic emitter or a hollow cathode device. The only differences 

occur with small values of total circuit resistance (R I 1000) with the 

hollow cathode device. Using a thermionic emitter, no differences occur. 

These results indicate that GATE is current limited strictly by electron 

collection. 

b-d), 

Comparing between particle collection models (comparing a with b and c 

with d), significant differences are encountered. The Dobrowolny- Janve 

model results in an increase of almost an order of magnitude more current 

flowing in the tether. These results occur with either electron emission 

device. 

22 
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23 
The accompanying figures all show the power converted from the tether 

current to batteries as a function of total induced voltage. Once again, 

figures a and c compare the thermionic emitter with the hollow cathode for 

the Dobrowolny-Janve model while figures b and d do the same for the ParkerMurphy model. Figures a and b compare the power converted with a thermionic 

emitter for the two particle collection models. Figures c and d perform the 

comparison for the hollow cathode device. Two observations can be made: 

Virtually no differences occur as a result of the electron emission process; 

and, .the Dobrowolny-Janve model results in a prediction of about five times 

the converted power. 

24 
II 
A comparison of the power converted as a function of total circuit 

resistance is given. Figures a and c compare the two types of electron 

emission devices with the DJ model while b and d compare makes the comparison for the PM model. With the 

PM model, no maximum power level is achieved up to 10,0000. With the DJ 

model a maximum power conversion is achieved between 2000 - 

All plots are done using a VBL of 350V. 

30000. 

26 
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W 

111111111111 

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t/ I 55 

ll- W 

27 
The tether current, electro-drag force, orbit degradation, and power 

generated are given versus orbit number. The EMF source for this simulation 

is the induced VBL of page 11. In each figure, the results are shown using 

both the Parker-Murphy (PM) model and the Dobrowolny-Janve (DJ) collection 

models. The method used for electron emission was thermionic emitters. For 

the DJ model the charging current averaged 53ma; the drag force averaged 

2.25mN; and, the power generated averaged 9W. For the DJ model the altitude 

was reduced approximately 5km per day. The results using the PM model were 

approximately an order of magnitude less than those obtained from the DJ 

model. Variations of the colatitude of the system with geomagnetic north 

caused the currents, forces and power to vary by about 15%. 

28 
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29 
These figures show the tether current, electro-boost force, power 

supplied by the batteries and orbit increment versus orbit number. The EMF 

source for this simulation was 700V supplied by a DC to DC converter less 

the induced VBL of page 18. The satellites were 'flipped' for this simulation so that current would flow down the tether. The total series 

resistance was 200n (100 in the tether and lOOn DC-DC convertor output 

resistance). Once again, the performance was bracketed by use of the two 

collection models and thermionic emitters were considered. The current 

averaged 135ma using the Dobrowolny model; the thrust averaged 5.5mN; the 

orbit was boosted by 12km per day; and, an average power of 95W was expended. Using the Parker- Murphy model these results were reduced by an 

order of magnitude. The currents, forces and power were not constant due to 

the variations of the colatitude of the GATE with respect to geomagnetic 

north. 
ELECTRODYNAMIC CONCLUSIONS 

------ TETHER CURRENTS 1OX USING 

DOBROWOLNY MODEL 

MODEL CV,; 

FOR DOBROWOLNY 

= 350V) 

YL 

------ 6,25W TO LOAD FEASIBLE 

------ NO FUNCTIONAL DIFFERENCES 

BETWEEN THERMIONIC EMITTER 

AND HOLLOW CATHODE OVER 

GATE RANGE OF INTEREST 

31 
In summary, the tether currents are approximately an order of magnitude 

greater using the Dobrowolny-Janve model over the Parker-Murphy model. With 

the DJ model the power converted to storage is maximized at a load 

resistance of 3kn while using the PM model a maximum was not In 

addition, 9W to the load appears to be feasible in a charging cycle. The 

observed. 

most surprising result was that no function differences occurred between 

thermionic emitters and hollow cathode devices. 

32 
(No Figure) 
- COORDINATE SYSTEM 

- SEPARATION AND STABIiIZATION ALObIG LOCAL 'JEWXciL 

- ORIENTATION (ATTITUDE) REVERSAL 

34 
M7 

EARTH 

CENTER 

35 

Y/ I I. \/ Y 
A dynamic model of the GATE system was developed and implemented to 

allow deployment and attitude control simulations. Out-of-plane motion was 

not implemented due to the preliminary nature of this work. The planar 

coordinate system used in the simulation is given in the opposite figure. 

The X-axis is defined to be tangential to the orbit plane, while the Y-axis 

is directed away from the Earth's center. 

and M2 represent the masses of the two subsatellites. 

The origin is located at M1. M1 

All planar motion is 

defined relative to mass M1. The angle theta is a measure of M2 relative to 

local vertical. L and L are the deployed length and length rate. 

When the satellite system is ejected from the GAS can the motherdaughter subsatellites are connected by a locking mechanism. This locking 

mechanism is released and the springs give an initial separation velocity 

between the mother-daughter pair. As separation continues the the gravity 

gradient forces and the diffrerential centrifugal forces cause the separation to accelerate. Initially, the friction wheel deploys 300 m of tether. 

The second 700 m is deployed by the motor driven reel in the mother. During 

this deployment phase, an active control system regulates the length, length 

rate, and necessary damping until the system is fully deployed (1 km) and 

stabilization is achieved. 

36 
r 8 

su 

u; 

c3 
Since the GATE system has no thruster reaction control system, explicit 

control over translation and rotation is limited. Therefore a more subtle 

approach is necessary. A control law was developed by Rupp5 that only 

requires the amount of tether deployed, the rate of tether deployment, and 

the commanded length. This control law assumes one of the subsatellites is 

much more massive than the other. Although this is not the case in the GATE 

system, this presents only minor modifications in the control algorithm. 

The opposite figures (a-d) show a typical deployment sequence using the 

control laws mentioned where the satellites are initially aligned along 

local vertical (60 - 0'). Figure (a) shows deployed length, L; (b) shows 

length rate; (c) shows angle from local vertical, 6 and, (d) shows 

tension in the tether. As the deployed length exceeds the commanded length, 

the tension in the tether increases to return the deployed tether to lkm. 

The in-plane angle from local vertical swings from 0" as the tether deploys 

to about -45" and settles out to local vertical. The complete deployment 

and settling requires less than 2.5 orbits. 

38 
39 
The opposite figures (a-d) show a typical deployment sequence where the 

satellites are initially aligned along local horizontal (Bo = 90"). This 

represents the worst-case in-plane angle from which deployment is initiated. 

Figure (a) shows deployed length, L; (b) shows length rate; (c) shows the 

angle B from local vertical; and, (d) shows tension in the tether. As in 

the the tension control law stabilizes 

the system to a full length of 1 km. The in-plane angle starts from 90" , 

and as the tether deploys, swings to a maximum of -60" before settling out 

to local vertical. The complete deployment and settling requires less than 

2.5 orbits. 

previous deployment case (Bo = O"), 

40 
GATE Orientations 

A B C D 

A. GATE system in "charging" orientation 

B. Tether is partially reeled in giving a change in 

the system inertia which causes a torque Tfabout 

the CM 

C. 

D. GATE system in "electro-boost" orientation 

Torque Tfis used to fully re-deploy the system 

I 

41 
As a strictly collection satellite-emission satellite configuration is 

envisioned, there are two attitudes needed along local vertical. Therefore 

a method for reversing the orientation of the system is necessary. This 

"flip" maneuver about the system's CM may be accomplished using the same 

control law as was used in initial deployment. When the decision is made to 

reverse the orientation of the system, the controller will begin tether 

retrieval. The change in moment of inertia causes the spin rate to increase 

from the normal orbital rate to about twice orbital rate. This manuever is 

depicted in the figure. In preliminary dynamic simulations the retrieval 

length was 300 meters which produced enough angular rate to spin the system 

past local horizontal. Once the system rotated beyond local horizontal the 

control system began re-deploying until stabilization along local vertical 

was again achieved. Simulation has shown that this control method is 

feasible in accomplishing the orientation manuevers. 

42 
00 

. 

0 

m' 

m 

m - 

X 

0 

. 

0 

8 

m 

I r4 

I- - 

Q) cu 

m 

cp 

m CSNOIM3N> NO ISN31 

-- 

c.:.. ___ .. ... I> --.> 

-- .. .. ._ -. . .. ._ - . . . . . . . ~ 

43 
Figures (a-d) show a 'flip' manuever. As in previous figures, (a) 

shows deployed length; (b) shows length rate; (c) shows angle from local 

vertical; and, (d) shows tether tension. After initial deployment and 

stabilization, the tether is rapidly reeled in to a length of 300m causing 

the system to go unstable and the in-plane angle, e, to rapidly change. Once 

the system crosses local horizontal, the control law again initiates the 

deployment sequence. The tether tension is once again modulated using the 

length and length rate control laws ,whereby, oscillations in in-plane angle 

and length are rapidly damped. 0 changes from 0" to 180" within approximately one-half orbit. 
.. 

i 

CONCLUSION 

- DEPLOYMENT AND STABILIZATION OCCUR WITHIN 2.5 ORBITS 

- ORIENTATION REVERSAL AND STABILIZATION REQUIRE 2.5 ORBITS 

AND AN ENERGY REQUIREMENT OF APPROXIMATELY 3Q JOULES 

- CONTROL ALGORITHMS PROVED ADEQUATE IN PRELIMINARY ANALYSES 

I 

- _____~~ ~ ~~ 

45 

~~ 
As seen in the previous figures of deployment and attiude control, 

impulsive type tension spikes occur during deployment when the commanded 

length is initially achieved. Although the tension spike is only on the 

order of 1 Newton for GATE, it could excite tether modal vibrations and 

librations. A smoother control strategy could minimize these effects. Ongoing work involves the study of uniform and exponential deployment rates 

such as those 

application of 

proposed by Misra and Modi6 , 7. Control laws 

Linear Quadratic Regulator theory using length, 

based on the 

length rate, 

pitch and pitch rate are also being investigated a . In conclusion, the 

simulations show that deployment, stabilization, and attitude control is 

feasible for the GATE system. three 

dimensional dynamic simulation incorporating aerodyamics, electrodynamics, 

vibration effects, and alternate control strategies for deployment, stationkeeping, and attitude control. 

Future work includes development of a 

46 
---- - .. ... .. -. __ .- 

t 

W 

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>- 

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IZ 

W 

E 

W 

m 

a 

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E 

to 

a_ 

r 

a 

H 

\ 

7 

0 

cn 

7 

W + 

W 

la 

c3 

H 

I- 

- 

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a 

1 

I I I 

47 
GATE TENSION/IMPACT DYNAMICS 

The GATE Tension/Impact Measurement System is shown in Figure 7. The 

system provides two signals; The tension signal 

is used in the feedback loops of the tether controller. The impact alert 

signal is used to alert the dormant computer that an impact has occurred and 

tension should be monitored for experiment analysis. 

tension and impact alert. 

A number of sensors have been evaluated including piezoelectric, displacement tranducers such as variable resistance/reactance transducers and 

strain gauge sensors. Semiconductor strain gauge sensors appear to be the 

best candidates for meeting the dual goals of the system. Environmental 

compensations, calibration, and motion errors are the areas requiring 

additional work. 

48 
h 

E 

W 

I 

IO 

E v 

- 

0 

Z 

n 

a 

0 

z 

z 

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0 

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a 

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w 

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aa 

t: 

m 

t 

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I- W 

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>- 0 

cn a 

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cn 0 

c3 

z 

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49 
cGiTc"K$j S-I-STm 

The command and control system is responsible for both spacecraft and 

experiment management. Performance, reliability, and resource optimization 

are key objectives. Pre-deployment status, deployment, station keeping, and 

experimentation are the primary tasks of the command system. The proposed 

subsystems are sensor, power, communication, and tether. Coordination and 

control of all subsystems will be provided by a central microprocessor. The 

microprocessor duty cycle will be divided between spacecraft operations such 

as attitude control as well as the scheduling, execution, and data handling 

of the experiments. 

The command system will also be responsible for managing all aspects of 

experimentation. These include experiment scheduling, execution, and data 

I handling. This will require knowledge of both system and environmental 

I 

, 

constraints. High data acquisition and handling rates are not anticipated 

because of the frequency characteristics of the experiments. Spacecraft 

1 

1 

management will include error checking, deployment, attitude control, and 

stationkeeping. Stationkeeping involves maintaining a local vertical orientation in the face of tether librations and disturbances. 

I Spacecraft and experimental activities are loosely synergetic: no 

I 

I experimentation will occur during spacecraft system operations such as 

initial deployment and orientation maneuvering. Once the system is aligned 

along local vertical (nominal or electroboost orientations), the primary 

efforts of the command system will be devoted to experimentation. Occasional stationkeeping manuevers will be necessary to maintain the desired 

attitude. Recause cf the frequency of spacecraft system operations as weii 

as low data rates, the total number of operations is well within the 

capabilities of a central low power microprocessor. 

50 
\, 

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(v 

0 
THERMAL PROFILES 

Two cases of the thermal properties for the GATE were studied: a low 

emissivity case and a high emissivity case. In each case the satellites 

were properties 

given in the Physical Layout section. In the low emissivity case, The 

satellites were considered to be covered with evaporated aluminum. In the 

high emissivity case, the covering of the satellites was considered to be 

dull stainless. The accompanying figure shows the thermal histories for the 

two satellites over a two orbit period in the steady state. For the high 

emissivity case, the daughter maintains a temperature close to room 

temperature; the mother changes a total of about 30°C with an average of 

265°C as it passes into and out of Earth's umbra. In the low emissivity 

case, both satellites heat up to over 350". 

assumed to be spherical for ease of analysis with the mass 

52 
Table 3 

GATE POWER CONSUMPTION 

Sub s y s t em 

1. Electroboost 7w x 24 hr/day = 

1 mission @ 15 day 

2520 w-hr/mission 

90 A-hr @ 28V 

2. Mother 

A. Computer & Electronics 

B. Communications 

C. Control 

D. Other 

TOTAL 

90 day mission - 3375 w-hr 

120 A-hr @ 28V 

3. Daughter 

Sun Pointing 

90 day mission = 1035 w-hr 

40 A-hr @ 28V 

W - Hr/Day 

168.0 

6.5 

3.5 

17.5 

10.0 

37.5 

- 

11.5 
-- GATE POWER 

The projected GATE power consumption is given in Table 3. Three 

power systems are envisioned: 1. electroboost; 2. Mother; and, 3. 

Daughter. One or more of the systems will be recharged using the induced tether current allowing for reduced battery capacity. Total 

required power without recharge would be 210 A-Hr @ 28V for the mother. 

From space and weight consideration, 200 A-Hr @ 28V of silver-sinc 

batteries can be incorporated into the mother satellite. Additional 

storage of 40 A-hr can be achieved in the daughter. 

54 
TABLE 4 

GATE POWER REQUIREMENTS WITH A 30 DAY 

9W CHARGING CYCLE 

Subsys tern w/charging w/o charging 

1.Electric Boost (2 missions: 15 days each) 

Required: 5040 W-Hr 5040 

Supplied: 2240 W-Hr 0 

Net: 2800 W-Hr 

28V @ 100 A-Hr 180 

2 .Mother 

Required : 6750 W-Hr 

Supp 1 i ed : 4240 W-Hr 

Net: 2510 W-Hr 

28V @ 90 A-Hr 

3.Daughter 

Required: 

Supp 1 i e d : 

Net: 

28V @ 

2070 W-Hr 

605 W-Hr 

1465 W-Hr 

52 A-Hr 

55 

6750 

0 

- 

6750 

210 

2070 

0 

2070 

74 
The projected power requirements with a charging current of 30ma 

generating 9.0 W of usable energy for a 30 day charging period are given 

in Table 4. Under these apparently reasonable conditions, two electroboost experiments could be accomplished with the available energy 

storage capacity by shifting a small fraction of the total stored energy 

from the mother to the daughter. After the second boost experiment, 

the satellites would be returned to the charging configuration for reentry into the atmosphere. 

56 
GATE SUBSATELLITE (DAUGHTER) 

PLAN VIEW I 

GATE SUBSATELLITE (MOTHER) 

PLAN VIEW 

r--- 

I 

.I. 

A 

w 

TENSION CONTROL 

SENSOR ELECT. 

I------ 1 --- 

I 

.I. 

A 

w 
PHYSICAL LAYOUT 

The first figure shows the plan view of both the mother and daughter 

satellites. A solar collector is shown in the daughter for the heating of a 

thermionic emitter. The control electronics and microprocessor are for the 

pointing of this collector. If plasma contactor is used, the daughter would 

contain only the plasma source an control, batteries, friction reel and 

separators. It is envisioned that about 300 m of tether will be deployed 

from the friction reel. 

The mother satellite contains most of the batteries and virtually all 

control and experimental apparatus. It also contains a reel with 900 m of 

tether. reel is actively controlled by a motor driven by the control 

laws. A tension meter is envisioned for motor linearization and an odometer 

for the tether control algorithms. A DC-DC converter is shown for recharging and electro-boost experiments and loading coils for LF propagation 

studies . 

This 

58 
J, 

I 16 

L I \ - 

lk 

T 

2.5 

GATE SATELLITE SYSTEM 

SECTION VIEW A-A 

DEPLOYMENT 

#?M 

REFLECTOR 

EMITTER 

LATCH01 FEMALE - SEPARATORS(4) 

Tl 

FRICTION 

JOINT LATCH(4) MALE 

SENSOR - 0 THERMALLY CONOUCTIVE f ILLER 

ANTENNA REEL MOTOR 

XMITTER 

AND h h EXP. 

GEARS ELECT SENSOR 

BATTERIES 

3 

2088 IN 

FIBERGLASS SHELL WITH 

CONDUCTIVE P AINT 

v l9 -----+I 

ALL DIMENSIONS IN INCHES 

59 
As seen in this figure, the mother is two layered with batteries on the 

bottom layer. This design keeps the C.M. of the system low relative to the 

GAS ejection plate. Approximate weights from the system are: mother, 120 

lbs; daughter, 30 lbs. Prior to separation the deployment arms fold and the 

daughter nests into the top of the mother. Latches maintain integrity 

during launch and ejection from the GAS canister. Separation springs 

provide the initial velocity to deploy the satellites away from each other. 

60 
b 

61 
Measurement of the radar cross section of the long wire will require 

rendezvous of the orbiter with GATE at least one day after deployment. The 

orbiter should station keep at a range of 2 km to allow radar operations to 

begin. The measurements will consist of commanding the radar (with the 

orbiter-.in an appropriate attitude) to point to the approximate location of 

the center of the tether. The radar is then commanded to search while the 

signal strength (AGC) is monitored on telemetry. The search pattern is a 

spiral motion and will intersect the tether at various angles. 

- GATE PROJECT SCHEDULE 

The GATE project schedule is shown in Figure 7. Preliminary design 

will 

mence on Oct 87 lasting 1 year. 

starting Oct 87. 

start Jan 87 continuing through Sept 87. Design is projected to comDevelopment and testing will last 18 months 

The flight will last approximately three months. 

Possible PSN involvement would be in the areas of electrodynamics 

and controls. Specifically, PSN has expressed interest in to studying 

plasma contactors, the charging and switching systems and the electroboost tradeoffs. Joint efforts will be conducted in the control and 

power systems. Final decisions relating to cooperative efforts will be 

made at appropriate levels. 

62 
CONCLUSIONS 

------ NO MAJOR TECHNOLOGICAL 

STUMBLING BLOCKS 

------ FREE FLYER ADVANTAGES 

------ 60 DAY MISSION POSSIBLE 

------ DEMONSTRATE ELECTRO BOOST 

AND ORBIT CHARGING 

------ RECORD METEOR IMPACTS 

------ DEMONSTRATE CONTROL STRATEGIES 

------ FAST TURNAROUND FOR 

POSSIBLE FUTURE WORK 

63 

- 
CONCLUSIONS 

The four experimental objectives of GATE appear feasible from our 

preliminary analyses. In addition, the total system also appears feasible. 

No significant management problem or engineering impediments are foreseen at 

this time. Technology development's expected in the meteor impact system 

and the control laws of the command and control system. Current technology 

will be used in all other areas and many "off-the-shelf" components will be 

used. 

The GATE will add to the body of knowledge of plasma dynamics, tether 

charging and electroboost. The results will complement the work of TSS-1 

and may help plan the second TSS electrodynamic mission. In addition, the 

GATE will break new ground in the study of micrometeriod hazards to space 

tethers. GATE can also provide for more than one set of experiments since 

it appears to be relatively inexpensive to build and low in integration 

time. As GATE is a free flyer, safety concerns are minimized. 

64 
REFERENCES 

1. Applications of Tethers in Space, Vol 1. Proc. of a Workshop in 

Venice, Italy, Oct 15-17, 1985. 

2. Martin Marietta Corp., "Shuttle/Tethered Satellite System (TSS)", 

Doc# MCR-78-576, Contract NAS8-32854, June 1978 

3.Stluka, E.F., "Marshall Amateur Radio Club Experiment (MARCE)", Get 

Away Special Experimenters Symposium, Goddard Space Flight Center, Oct 

7-8, 1986 

4."Shuttle Tethered Satellite System Test on Ku Band at White Sands 

Missile Range, NASA TM# JSC 22242, July 1986 

5. Rupp, C. C. "A Tether Tension Control Law for Tethered Subsatellites 

Deployed Along Local Vertical", NASA TM-64963, 1975. 

6. Modi, V.J.and Misra, A.K.,"Deployment Dynamics of Tethered Satellite 

Systems," AIAA/AAS Astrodynamics Conf., Palo Alto, CA, Paper# 78-1398, 

1978 

7. Modi, V.J.and Misra, A.K., "On The Deployment Dynamics of Tether 

Connected Two-Body", Acta Astronautica: - 6, pp1183-1197, 1979 

8. Bainum, P.M. and Kumar, V.K., "Optimal Control of the Shuttle - 

Tethered - Subsatellite System," Acta Astronautica: 1, pp1333-1348, 

1980 
9. Martinez-Sanchez, M. and D. E. Hastings, "A System Study of a 100 Kw 

Electrodynamic Tether", NASA/AIAA/PSN International Conference on 

Tethers in Space, Sept. 1986. 

10. Williamson, P. R. and P. M. Banks, "The Tethered Balloon Current 

Generator: A Space Shuttle-Tethered Subsatellite for Plasma Studies and 

Power Generation", Final Report NOAA Contract 03-5-022-60, 1976. 

11. Olsen, R. C., "Modification of Spacecraft Potentials by Thermal 

Electron Emission on ATS-5", Journal of Spacecraft, Vol #18, pp. 527- 

532, Nov. 1981. 

12.Parker, L. and B. Murphy, "Potential Buildup on an Electron-emitting 

Ionospheric Satellite, Journal - of Geophys. Research, Vol 72, p 1621, 

1967. 

13. Dobrowolny, M. and A. V. Janve, "Space Charge Around Conductors at 

High Potential in Magnetoplasma", Istituto Fisica Spazio 

Interplanetario, CNR Frascati, Italy, 1984, IFSI-84-4. 

15. Langmuir, I. and H. M. Mott-Smith, "The Theory of Collectors in 

Gaseous Discharges", The Collected Works of Irving Langmuir, Vol 4, G. 

Suits, editor, Pergamon Press, N. Y. 1961 

16. Al'pert, Y. L., A. V. Gurevich, and L. P. Pitaevskii, "Space Physics 

with Artificial Satellites, Plenum Press, N. Y. 1965 

17. Linson, L., "Current-Voltage Characteristics of an Electron-Emitting 

Satellite in the Ionosphere", Journal .f Geophysical Research, Vol 74, 

p. 2368 1969 

18. Beard, D. B. and F. S. Johnson, "Ionospheric Limitations on "Attainable Satellite Potential", Journal of Geophysical Research, Vol 66, 

p. 4113 1961 

66