The Use of Electric Microthrusters for Orbit Correction of Geostationary Satellites

THE USE OF ELECTRIC MICROTHRUSTERS FOR ORBIT CORRECTION OF GEOSTATIONARY

SATELLITES

P . LUQUET (CNES - FRANCE)

CONTENTS

1 - INTRODUCTION

1 - INTRODUCTION

In order to maintain in position (stationkeeping) geostationary satellites planned for missions lasting 5 years and lon ger, it has been necessary to develop thrusters designed to suppl y very small thrust levels, on the order of a millinewton. These microthrusters must also possess an operating service life of several thousand hours (1).

2 - UTIL IZATION OF MICROTHRUSTERS FOR THE CORRECTIONS I N THE POSITION OF GEOSTAT IONARY SATELLITES 2. 1 - Perturbations in orbit. 2.2 - Necessar y corrections - Performance and t ypes of corresponding thrusters.

A further design-goal is to obtain high specific impulses ; this is important to minimize the mass of on-board propellant, and therefore of the complete control system.

2.3 - Number of thrusters required. 2.4 - Mass of stabilization devices.

~

Within the framework of its researches, the CNES Program and Plan Authority has endeavoured to develop amnonia electrothermal and ion microthrus ters which are solving, \~i th increasing effectiveness, the various problems present in this new field.

- THRUSTER OPERATION CHARACTERISTICS 3.1 - Ammonia heated thrusters. 3.2 - Ion thrusters .

The scope of this summary is as follows :

4 - TECHNICAL STUDIES AND DEVELOPMENTS IN FRANCE

a) the feasibility of using these microthrusters for the mission contemplated

5 - CONCLUSION

b) their characteristics (principle of operation and performance) c) the present state of their development

BIBLIOr,RAPHY

d) their prospects for the future space programs .

FIGURES

382

:' -

PT TV

ZATTON OF

Ml(~R()TH~ U STERS

FOR THE CORRECnO NS TN THE POS lTIO N OF r:EOSTATTO NARY SATELL TTES.

A .<,eos 1al ; onarv sa1elli. '·e is su :; .iected eo summari ze d in the table below :

rl.~ ree

classes of

per r ur~ati.ons

.--- ---------------------------~------------------------------

Cause of the perturb a ti on

Effects on the satellite position

- Radiation pressure - Error of alignment of the thrust vector s wi th th e center of gr avity - Spurious impulses (valves, etc . .. )

Attitude drift

Tri axia lity of the Earth (lack of sphericity l eadin g to variations in the terres trial gravi t y potent i a 1. )

Drift in longitude East or Wes t , towards the nearest stable point (75 ° E or 105 ° W) at the rate of 0.44 ° / day (maximum) .

Moon - sun attractions

Drift in l at itude (North - South) of the orb ital plane with res pect to the equatoria l plane, from 0.75 to 0.95° / year, depending upon the value of the lunar o rbital inclination

(rotation around cente r o f grav ity )

(2)

2.2 - Necessary corrections - Performance and t ypes of correspon ding thrusters.

..

their effects are

--------------------------, Cost o f the c o rrection expressed in "velocit y increment o.V" ( o.V = velocity variation which has to be imparted to the satellite to cancel the drift)

<. lm / s per year (value usually given : 3m/s over 5 years).

5.5 m/ s per year (maximum)

(2)

- function of satel lite's pointing accuracy (ali gnment) - on the order of 5Om/s per year for a pointing accurate to 0.15°, ap proximately, for the corrections applied in the immediate vicinity of each node (optimum operation from propellant consumption standpoint. See 2.2.1

required direction . in velocity on its trajector y (orbital control in latitude and in longitude), to maintain it in a fixed position wi th respect to Earth, in the des ir ed poin tin g direction.

As a result of the perturbations pre vious l y described, th e stab ilization sys tem must allow the satellit e to be contro l led :

Each t ype of mission (or here, of correc ti on) i s cha racterised by :

- in pos ition arou nd it s centre of gravit y ( a ttitude contro l ), in order to orient the various elements (anten nas - so lar panels, etc .. . ) in the

a) the tot a l impul se to be supplied : I t = F.t = M. D. V.n in N.s.

383

F t

thrust level in N. total mission duration in· s. M mass of satellite in kg. AV velocity increment in m/s per year n = numbers of years of satellite operating life.

system must exert a thrust tewards the North when the satellite passes through the ascending node, and the other towards the South when the satellite passes through the descending node. If 2~ is the angle during which each of these corrections is made, the velocity increment required varies in accordance with the law :

This relation is valid as a first approximation on the assumption t.hat M is constant during the whole mission duration (consumption of propellant negligible compared to the satellite mass. Base of high specific impulses).

0(

sincx

Mit1V[1-~e~+i(~~)2

J

initial satellite mass

Ve

velocity at which the propellant fluid particles are ejected.

I~

b) the mass of propellant required It M. 6V. n m in kg Isp.g Isp.g Isp g

=

Specific impulse in s

o

H SOm/s per year (3)

The minimum consumption of propellant corresponding to o.V mini = AV = SOm / s per year is therefore obtained for ~irings (impulses) of very short duration ( 20( minimum) on passing through each node; in consequence, the thrust level must be high in order that the value of the product : I = F x t may be obtained. Chemical propulsion is perfectly suitable for this type of impulsional operation.

.....

M.

c:..V

Figure 1 gives the value of 6V with respect to the durations of each correction around the corresponding node.

A rigorous relation linking the total impulse to the velocity increment is given in reference 7. It

with

Apart · from its lack of finesse, the mode of correction suffers from the major drawback of imparting violent mechanical shocks to the satellite ; this in turn means that t.here is a danger that the deployable element likely to be used on these satellites (gravity gradient stabilization booms - antennas flexible structure solar panels, etc.) may be subjected to large-scale vibration, and may even break off. Also, owing to the fact that the vector thrust may not pass sufficently close to the center of gravity, the satellite may tilt or swing over.

Ve g

acceleration due to gravity at ground level = 9.81 m/s 2

The North-South control is by far the most costly in propellant mass, owing (0 the high level of velocity incre ment required ( § 2.1 )

For this reason, preference is being given at the present time to the operation in a continuous regime, utilizing extremely low thrusts on the order of a millinewton, and functioning over long periods of an order of magni tude of several thousand hours.

These are therefore the corrections which condition to a very great extent the mass of the control system, and which justify : - the search for high specific impulses - the optimisation of the mode of application of the thrust (level and duration) .

Chemical propulsion cannot be used for this mode of operation ; on the other hand thrusters of the following types are eminently suitable :

2.2.1 - Control in latitude (or NorthSouth control)

- ammonia electrothermally heated microthrusters, often called "ammonia resistojets" (thrust: a few millinewtons ; service life: a few thousand hours).

Each North and South correction must be carried out around the corresponding node (point at which orbital plane intersects the equatorial plane - see figure 1) in such a way as to cancel the satellite's velocity component perpendicular to the equatorial plane.

- Cesium contact ionization (thrust equal to, or less than, 2 m N approx. - service life on the order of 20,000 hours). a) When the specific impulse is high, as in the case of ion propulsion (Isp = 4000 to SOOO s), very little propellant is consumed, whatever be the value of the velocity increment (about 1.S kg for a S year miss ion) in this case al most

This requires the use of two propulsive systems able to operate at l2-hour intervals:one

384

pass through the satellite's center of gravity in order not to perturb the attitude.

continuous corrections may be envisaged (10 hours/day arou~~ each node with a thrust of 0.5 x 10 N for a 200 kg satellite (4) permanently opposing, or nearly permanently, the satellite drift, and thus maintaining it practically fixed with respect to Earth. This entails a thruster life on the order of 18,000 hours .

One sole thruster would in theory suffice, but in practice it is worthwhile providing at least a second one, diametrically opposed to the first one, and thrusting in the reverse direction, to allow of a more flexible control.

b) In the case of heated ammonia propulsion, where the specific impulse is distinctly lmo}er ( :=:. 200 seconds) (5), such an operation would be much too costly in additional propellant m,;ss : the duration of the corrections must therefore be reduced, with a consequent increase in the thrust level, in order to obtain the total impulse required : It = F , L

Since the thrust levels are in this case much greater than what is actually required, the East-West corrections : are much shorter (in the ratio of the velocity increments for equal thrust)

This involves letting the satellite drift in the tolerated range, and to reset it when it reaches the boundary limit of the drift zone .

may be performed outside of periods of North-South controls, making use of the same source of power.

Thus, to align a 200 kg satellite within 0.2 degree under the optimum conditions the North-South corrections must be carried out by means of an 8 . 10- 3 N thrust (4 thrusters of 2. 10-3 N) applied for 10.5 hours around each node, over a period of 4 consecutive days, and with a period of 84 days (6).

2.2.3 - Attitude Control Since the cost of attitude control is very small ( ~V ~ 1 m/s per year) this mission does not justify the researches to achieve high specific impulses; with a level of 70 seconds typical of cold gases (nitrogen-ammonia), and the sublimable solids, the consumption of propellant would in fact be limited to about 0.3 kg/year for a 200 kg satellite.

Under these conditions : - the mass of propellant consumed is on the order of 28 kg for a 5-year mission (It = 55 600 N.S. - Isp = 200 s). -

It is preferable, in the interest of the homogeneity of the complete stabilization system, to use thrusters similar to those serving for the North-South controls.

the overall time of operation of the thrusters is about 900 hours.

Nevertheless, when a system of orbital correction by ion thrusting is already mounted on-board the satellite, this system offers particularly interesting possibilities since these same thrusters may be utilized for attitude control by electrostatically deflecting their beam (§ 2.3) (7).

On this scale of miniaturization, the electrical powers required i.e. - 65 W/ mN for the ion thrusters 6 W/mN for the resistojets

The control logic is in this case more complex (deflections of ion beams slaved to attitude sensors), but may form the basis of a particularly advanced 3-axis stabilization system. (8)

remain compatible with a reasonable solar panel mass ( :;,: 0.140 kg / w at present and 0.020 kg/w from 1980 with flexible structures). 1.2. - Control in longitude (or East-West controll

Should resistojets be used for orbital corrections, a small fraction of ammonia may be taken from the main tank to feed the cold gas jets intended for the attitude control.

The maintenance in longitude is about 10 times less costly in propellant consumption than the North-South control (ratio of velocity increments, for the same specific impulse ). It requires therefore on the order of : - 0.150 kg of cesium with ion thrusters.

This method offers the advantage of an overall homogeneous stabilization system, using the same propellant for all thrust controls. 2.3 - Number of thrusters required

- 3 kg of ammonia with resistojets.

An active, 3-axis stabilization system (orbital and attitude corrections) of a geostationary satellite requires a large number of thrusters ; the directions in which the perturbating forces may be exerted, as described in the previous paragraphs, entails the

The thrust should - be exerted in the orbital plane, and tangentially to the trajectory in order to brake or accelerate the satellite, depending upon the direction of drift.

:;85

mounting of at least 9 thrusters for the ~orth corrections for the South corrections - 1 for the corrections in longitude (Eas t or Wes tl - 6 for the . attitude corrections (2 thrusters per axis to allow the restoring torques in yaw, roll and pitch.

tion very tion logy

of hydrazine notably), which forms the latest technology at present in applicain the Uni ted States ; this same technomav shortly be adopted in France.

In actual fact, for reasons of redundance and to be able simultaneously the thrust in two parallel directions, and in the same sense, in order to avoid torques causing the satellite to tilt (case of North-South corrections in particular), one can readily see that 20 thrust axes are needed (figure 2).

The difference noticed on figure 3 between these two types of propulsion should be reduced still further; in actual fact, al th0ugh the specific impulse given for the resistojets (175 seconds) is rather on the pessimistic side (200 seconds is a reasonable limit at the present time), that envisioned for the monopropellants (230 seconds) is on the other hand slightly optimistic : 215 to 220 seconds is a more probable order of size in rhe micropropulsion field, owing to the 101V efficiency of the nozzles.

By designing multijet assemblies using ammonia thrusting, a close grouping of the thrust axes may be achieved; at the limit, 4 assemblies of 5 jets each (9), arranged in accordance with the drawing of figure 2, may form an interesting solution to the problem.

The stabilization system would therefore lVeigh on the order of 40 kg for a 200 kg satellite, and a 5-year mission duration. This result agrees fairly well with the estimates which it has been possible to make more recently (12).

In the case of ion propulsion, the orbital correction thrusters alone may ensure the complete satellite stabilization, by the electrostatic deflection of their beam.

The 20 % of the satellite mass thus assigned to an ammonia system for a 5-year utilization reduces to about 5 % lVith ion propulsion. This percentage mass reduction, already considerable, continues to increase quite sharply lVith an increase in mission duration.

Disregarding all questions of reliability, the control system may then be reduced, at the limit, to 3 thrusters : - 2 lamellar beam thrusters (type single-slit § 3.2.2.-2°) for the latitude (North-South) and roll controls (deflection of beam around a single plane). - 1 revolution beam thruster (type mono button) for the longitude (East-West), yaw and pitch controls (deflection of beam around two perpendicular planes). The use of thrusters with 2 lamellar orthogonal beams (10) the development of which is contemplated by "HUGUES Research Laboratories" 1V0uld comprise a more advanced solution al10IVing : -

the corrections to be carried out around two perpendicular planes at a thrust level higher than that used with single-aperture thrusters (§ 3.2.2. _ 2°)

- an increase in the redundance of the stabilization system. 4 - Mass of stabilization devices The graph of figure 3 (11) shows at a glance the comparative masses of the stabilization systems with respect to the various types of propulsion treated up to now. The benefit of seeking high specific impulses for long duration missions is clearly apparent on this graph. The ammonia resistojets are classed in the field generally designated as chemical monopropellant propulsion (catalytic decomposi-

The almost horizontal shape of the curve relative to ion propulsion is due to the fact that the mass of propellant only forms a small part of the overall assembly mass. The main item in the overall mass budget is the power source (solar panels - conditioner) the mass of which is constant whatever be the duration of the mission.

3 - THRUSTER OPERATION CHARACTERISTICS 3.1 - Ammonia heated thrusters (27) 3.1.1 - Principle of operation The ammonia thrusters function following the conventional principles of jet propulsion. The gas is heated by convection in a tubular assembly raised to a high temperature ; the thermal energy is then converted into kinetic energy by expansion in a convergent-divergent nozzle. In the case of resistojets (electrothermal heating) the calorific energy is produced by Joule effect. The possibility also exists of heating the gas radiothermically by means of encapsulated isotopes ; these thrusters are then called radioisojets. The resistojets are split up into three groups, depending on their design-concept and the mode of operation. These groups are as follO\"s : a) LolV thermal inertia or fast heat up

program, to perform the attitude and longitude corrections of the ATS 4 and 5 satellites (M ~400 kg) stabilized by gravity gradient (respective launches : September 1968 and August 1969) (13) -(14).

thrusters (type AVCO Corporation) (9) for which the gas circulates in a straight capillary tube extended by the nozzle, and in which an electrical current simultaneously flows. raising the gas to a high temperature (figure 4).

For this application their performance data is as follows :

b) Thermal storage thrusters (type General Electric) (9) for which the gas takes only a small fraction of the thermal energy stored in an oven of very high thermal capacity (fi gure 4)

- thrust ~ 0.5 mN - total power : 10 W of which 5 of the gas heating (T: 900 0 K) - specific impulse ~ 150 seconds.

These thrusters operate by impulsions ; the frequency of the impulsions should be such that the oven's thermal equilibrium be little disturbed, in order to :

b) The thermal storage thrusters are on the contrary suitable for high thrust levels. Owing to their high thermal capacity, several hours may be required before the nominal operating temperature is reached (~ 1400 0 K) with the few score of watts available on board the satellites.

- obtain higher levels of specific impulse - minimize the consumption of power necessary to maintain the oven temperature. c) Continuous operation thrusters (type SNECMA § 4, fig. 5), which form a solution intermediate between the two previous ones ; their thermal insulation must be more efficient than that of the first type to limit the consumptions of power, but may be less complex than that of the second type since the gas absorbs at each instant all the energy available in the oven.

However, this response time does not necessarily form a drawback for orbital correction missions since these corrections can be predicted a very long time beforehand (§ 2.2.1-b) The General Electric Company has developed this type of thrusters (resistojets and radioisojets) to the extent that their performance capabilities are now sufficient to allow the control of heavy capsules, or the placing in orbit of smaller type satellites

The radioisojets operate on the principle of thermal storage thrusters.

(9)

All these thrusters may be grouped, as mentioned in § 2.3, in multijet assemblies.

- thrusts : 90 and 200 mN - power : 30 to 40 W - specific impulse : 230 to 265 seconds

The ammonia is easily stored in the liquid state, self-pressurized under the effect of the vapour pressure (10 bars at 25 °C) ; a better method is to store the ammonia in agelatinous state, eliminating the problems of sloshing in the tanks (6).

c) The level of thrust obtained from continuously heated resistojets is perforce much lower, in order to keep the consumption of power to a minimum.

These possibilities are advantageous with respect to chemical propulsion, which often require annex pressurization systems.

They can be miniaturized down to about 2 mN; this limit is imposed by the efficiency of the nozzles which falls of rapidly beyond this threshold value (problems linked to the roughness of the side-walls).

3.1.2 - Performance

1.

:

Their performance is an optimum for slightly higher thrust levels, on the order of 4 to 5 mN, suitable for orbit corrections of satellites of average mass (200 to 500 kg, for example).

!~:~~~_:_e~~::_:_~e::~~~~_~~e~~~:

a) The design features which characterize fast heat up resistojets (low inertia and thermal insulation) inevitably entail their miniaturization, having regard to the powers available on board the satellites.

The experimental thruster developed by SNECMA (figure 5) has the following characteristics (20 - b) :

These same features also make the res istojets suitable for pulsed or semi-continuous operation, by offering relatively low response times (a few tenths of a second).

thrust: 5 mN power 20 W specific impulse: 170 seconds. This value is low ; having regard to the average gas temperature reached (1100° K). It should be possible to attain a level of 200 seconds, through studies aimed at obtaining a higher nozzle efficiency.

These thrusters have been developed by AVCO within the framework of the A.T.S. (Applications Technology Satellites)

38 7

(~resent

efficiency

mass length diameter

2.

0.60)

60 g

Two ionization processes are used.

55 mm 37 mm

Endurance ..--------

The ammonia resistojets are based upon a relatively simple technology, which does not set any f~ndamental problems of service life; the most fragile parts are certainly the heater elements, but this difficulty has been overcome to a large extent, in the electronics industry in particular (endurance of emissive filaments).

In the first case, the propellant vapours Are introduced in a cylindrical chamber ; they are there bombarded by electrons at a level of energy higher than that of the threshold of the first atom ionization (10.4 eV for mercury - 3.9 eV for cesium) ; the electrons emitted by a cathode are gathered on a cylindrical anode placed at the periphery of the ionization chamber. An axial magnetic field is superposed on the electrical discharge field to lengthen the electronic paths, thus increasing the collision probabilities.

The problems ".hich arise wi th respec t to the propellant supply devices are similar to those of every conventional propulsion system; they concern in particular the solenoid valves~ and the dangers of leaks at the junctions. These difficulties should not be made more complex by the use of ammonia since when cold, and in the anhydrous state, this propellant shows an excellent compatibility with most of the conventional materials. .2 -

These sources of ions are called ~aufman sources, in honour of the inventor. In the second case, the cesium vapor is injected through a porous tungsten wall, ionization takes place, said to be "by contact", owing to the fact that the cesium atoms are carried to a level of energy, characterized by the work function of the material (4.5 eV), superior to their threshold of first ionization 0.5 eV).

Ion thrusters (7)

3.2.1 - Principle of £reration (figure 6) 1.

These are

mercury or cesium ionization by electron bombardment cesium contact ionization.

~~E~~E~~~_~~_E~=_E~:~~E

The ionizer should be heated to limit the cesium coverage rate at its surface in order to maintain a sufficient level of overall output potential (wet ionizer) and in this way to obtain a good ionization efficiency. This temperature is on the order of 1300 0 K for the corresponding flows of cesium at thrust levels measured in millinewton.

The thrust is produced by the electrostatic acceleration of positive ions created from the propellant atoms (mercury or cesium) : Only the ions are ejected at a high velocity (40,000 to 50,000 m/s) ; the electrons are collected in an internal circuit, the ground terminal of which is formed by the satellite structure.

3.2.2 - Performance

An equivalent charge of electrons must then be injected in the ion beam to neutralize it (cancellation of the space charge), whilst at the same time discharging the satellite s truc ture.

1.

Electron bombardment thrusters

The limit to which electron bombardment thrusters may be miniaturized is imposed in particular by the rate of recombination of ions, which increases rapidly when the volume of the ionization chamber diminishes.

An ion thruster comprises therefore three basis component systems, i.e.

This results in a poor ionization efficiency ( .:::: 50 % for the small thrus ters) ; the value of this parameter has a direct effect on the thruster service life. The presence of atoms in the ion beam is in fact at the origin of the formation of secondary, charge exchange, ions (collisions between slow atoms and rapid ions) which subject the accelerating electrode and the neutralizer to an intense bombardment.

the source of ions the ion optics (focussing and ejection of ions) the neutralizer. The propellant supply system, associated with the actual thruster, comprises: the tank allowing the fluid to be stored in a zero gravity state the vaporizer which supplies the necessarv electron flow to the source' of ions

The ion bombardment of the ionization cathode is another factor curtailing the thruster operating life. The mercury microthruster developed by ONE RA (figure 7) may be placed at the reasonable limit of miniaturization:

This propellant assembly is driven by the electrical power supply system which comprises the power and voltage conditioner the means of checking and of regulating the operation

diameter of beam thrus t

388

40 mm 2m~

- power

: 100 W

In the fiel d of high thrusts (a few tens of

mN), the multi-slit and multi - button th r us -

With the cathode technologies at present a va i lable for this range of thrusters (oxide cathodes), the operation life is limited to a few thousand hours .

ters (fig. 9) have been rapidl y supp l anted by mercury bombardment thrus t ers . On the ot h er hand the single-button thrusters developed by Electro Optical Systems (thrust between 10 - 5 and 10 - 4 N) and single - slit thrus ters deve l oped by Hugues Researc h Laboratories (thrust between 0.5 and 2 mN) are based on tec hno l ogies a l lowing the control of geosta tionary sa t ellites of masses limited to about 1000 kg (7) with all the inte r esting applica tions detailed in the previous chapters.

This figure should be greatly improved by the use of ho l low cathodes, the possibilities of miniaturizat i on of ,,,hich are at present being studied by NASA (L . R.C . ) (15). A much greater degree of miniatur i zation is obta i ned wi th cesium, owing to its easier io nization ; El ectro Optical Systems has deve loped thrusters of 12.5 mm in diameter, sup plying thrusts on the order of 5 . l0- 5N (16).

The elect r ostatic beam of these thrusters is deflected by applying a differential voltage on the elements of the acce l erating electrode a deflection angle of ± 10 " seems to be reaso nable l imi t ,,,hich must not be exceeded in or der not to de g rade the operating life of the ion optics ( 18) and of the neutralizer.

These are only experimental models, hm"ever , which do not appear to have been subjected to elabora t e endurance tests. The e l ectronic bombardment thrusters, on the other hand, are much more suitable for obt~i ­ ning high t hrust levels. A great effort has been made by NASA (L . R.C.) to develop mercury t h rusters for primary propulsion missions (space probes) Researches are at present being carried out over a very wide range, from diameters of 15 mm (thrust ~ 20 mN - power : 1000 W) up to a diameter of 1 . 5 m (F ~ 2N P '::" : 100 KW ( 15).

The Single - button thrusters have been deve l oped to an operational sta te in the United States within the framework of the A.T . S. pro gram. Th ese thrusters - l i ke the ammonia resistojets § 3 . 1.2. _1 ° ) have been successfully flight tested durin g the s hortened fli ght of the A.T . S . 4 satellite (19), which was placed in a too ldw orbit as a re su lt of a launcher m~l­ function. A complete experiment (attitude and longitude corrections) is being flown on board the the A. T . S . 5 satellite stabilized by gravity gradients recently launched (Au g ust 1969).

The SERT 11 experiment (change of orbit by mercury i o n thr uster - thrust: 30 mN - power 1000 W), expec t ed to take p l ace in ea r ly 1970, s ho u ld fina ll y prove the feasibility of this type of thruster for mission durations on t h e order of 10,000 ho ur s . 2.

Ces i um contact i onization thr usters

4. - TECHNICAL STUDIES AND DEVELOPMENTS IN FRANCE

The degree of miniaturisation of this type of thruster is not limited by any phys i cal phenomenon ; the s h ape and dimensions of t he th r usters a r e di rectly linked to those of the emissive surface - area (porous tungsten). Figure 8 shows the var i ous types of t h rus ters one can envi sage (lamellar or cylindri cal beam t hr usters) .

The Program and Pl an Authority of the CNES has o riented its researches to the systems of micropropulsion capable of being used on the comin g generations of sa tellites, the mass of which should vary between 200 and 1000 kg. It has as a result taken two options in prepar i ng the Ind u str y for the development of :

The pe r formance is essentially a function of the techno l ogica l characteristics of the io nizer, wha teve r be its size. With t h e pr ogress made in the manufacture of tungsten powders an d i n the sintering methods, it is possib l e :

continuous l y heated resistojets for a possib l e s hort nr medium-term ut i l i zation (beginning of 6th p l an) ; cesium contact ionization thrusters (sin gle - slit type) for applications in the mo re distant future (from 1975 onwards) .

- to obtain ionization efficiencies of more t han 99 % - to est i mate an operating life on the order of 20,000 hours for the ionizer i tself (prob l ems of ageing at a hi gh temperature) and for t he ion optics (very small frac tion of s e condary, charge annulling ions i n the beam. )

4 . 1 - Ammonia Resistojets The work carr i ed out from 1967 to 1969 at SNECMA (20) in coll abo ration wi t h t he Grenob l e Nuclear Study Center (Studies of the forced convection of ammonia at light f l ows) (21) have allowed : experimental thrusters to be developed, the c haracterist i cs of which were detailed in " 3.1.2.

The endurance problems concern therefore, above all, the neutralizer; the results already obtained with thoriated tungstenc a rbide filaments (a few thousand hours without erosion) are however very promising

the acquisition of a large store of theore tical and experimental know - how, g i ving France the capacit y to develop an opera -

{1n .

389

tional system.

5. - CONCLUSION

Furthermore, researches have also been carried out on radioisojets by the SNECMA Missiles and Space Division (22), and by the SE PR (23) (mer. ged in October 1969 under the name "Societe Europeenne de Propulsion",) in collaboration with the C.E . A. - Service des Radioelements (24). The aim of these researches is to check to what extent this type of propulsion may be envisaged for such space applications. The re sults have been encouraging. Data has been acquired on the type of technologies that will have to be developed to realize isotopic capsules.

Ammonia electrothermal propulsion is a relatively simple but advanced technology which has the advantage over chemical propul sion of being capable of supplying the low levels of thrust, if an order of magnitude of a millinewton, and the long operational lives (several thousand hours) sought for the control of geostationary satellites . French industry possesses the necessary capabilities to develop this type of propulsion for a medium term application. This technology could therefore serve as a tran sition between the use of chemical propu1 sion and that of ion propulsion, which ap pears to form the most advanced mode of propulsion which it is possible to envision at the present time for a development in the relatively near future.

4 . 2 - Cesium contact ionization propulsion In parallel with the work performed at ONERA (25) on this type of propulsion from 1968 onwards (engineering of experimental mo dels), basic studies concerning the operation of 3 main thruster systems (ionizer - ion optics - neutralizer) have been entrusted at the beg inning of 1969 to the Electronics and Applied Physics Laboratories (L . E.P.) (26).

Fesearches have been undertaken in France in this field with the aim of disposing of an operational system from 1975 onwards.

The L . A. A.S. "Laboratoire d'Automatique et de ses Applications Spatiales", has also zollaborated in these researches on the working theme - electrical converter and control logic (8).

In t~e United States, the feasibility of these advanced micro thrust systems (ion resistojets and thrusters) was proved in 1968 at the time of the A.T.S. 4 space experiment. The satellite A.T.S. 5, launched in 1969, should confirm the validity of these thrusters.

Having regard to all the know-how which will thus be acquired, one may estimate that an ad vanced thruster model, conforming to precise specifications, could be developed from 1971 onwards ; the design-goal is a first space experimentation round about 1975.

DISCUSSION (~.

What is the specHic impulse of the ionic engine?

A.

5,000 s.

Q. What type or electrical generator do you propose for this type of motor? C~les

A.

The electrical power

Q.

Will this power he sufricient?

from the solar panels of the satellite.

A. A 200 kgs telecommunications satellite will have an installed power of 500 IV and the power required for the propulsion system will be in the order of 50 \1'.

390

BIB L lOG RAP H Y

9. - WHITE A.F. (NASA

1. _ J.P. PUJES (CNES) Xx

C.S.F.C.)

Electrothermal microthrust systems AlAA Paper 67 - 422 - July 1967

Orbit and attitude control systems for synchronous satellites. Automatic control in space - 3rd I.F.A.C. Symposium - Toulouse, March 1970

10. - J.R. ANDERSON (Hughes Research Laboratories) R.S. CYBULSKI (AVCO CORPORATION)

2. - J.C. BLAIVE (CNES) x

Status and application of low thrust electric propulsion systems AlAA Paper 66-578 - June 1966

Stationkeeping of a 24-hour equatorial satellite Note JCB/CD/q - 157 - April 1969 CNES - CB/MT Division - B.P. N° 4 91 - Bretigny-sur-Orge

11. - C.R. BREWER - J.H. MOLITOR (Hu ghes Rese a rch Laboratories)

3. - J. CAUBEL (CNES) x

Timetable for ion propulsion Space and Aeronautics - June 1967 -page 92

Latitude corrections of a geostationary satellite - Comparison between the impulsional and the continuous modes. Note PR / AM N° 41 - January 1968 CNES - PR/AM Division - B.P. N° 4 91 - Bretigny-sur-Orge

12. - SNECMA (Missiles and Space Division - 3 3 Blanquefort) x Symphonie satellite - Proposal for the supply of an attitude control and stati o nkeeping system using ammonia. Document n O 145 / r - April 1968

4. - J.H. MOLITOR (HUCHES Research Laboratories) Optimization of ion engine control systems for synchronous satellites A.I.A.A. paper N° 63-273 June 1963

13. - C.D. BULLOCK (NASA) A.T.S. - Program Summary - April 1968 NASA C.S.F.C. - Greenbelt - Maryland - lISA

5. - J.P. PUJES (CNES) x 14. - R. SHAW - T.K. PUGMIRE (AVeO Co.) R.A. CALLENS (NASA G.S.F.C.)

Utilization of ammonia as propellant Note PR/AM N° 941 - August 1967 CNES - PR/ED Division - B.P. N° 4 91 - Bretigny-sur-Orge

Ammonia resistojet station keeping subsystem aboard Applications technolog y satellite (ATS) IV AIAA paper 69-296 - March 1969

6. - R. MORIN - C. MARION (SNECMA - Centre de Blanquefort _ Cironde) x

15. - E.A. RICHLEY - W.R. KERSLAKE (NASA L.R.C.)

Ammonia resistojet - International conference on the changes of attitude and the stabilization of satellites - Paris - October 1968 - Desgrandchamps printers - 161, Bd Brune - Paris 14°

Bombardment thruster investigatio~s at the Lewis Research Center AIAA paper 68-542 - June 1968 16. - G. SOHL - V.V. FOSNICHT - S . .!. GOLDNER R.C. SPEISER

7. _ P. LUQUET (CNES) x

Cesium electron bombardment ion microthrusters - Journal of Spacecraft and rockets, Vol . 4 - n ' 9 September 1967

Application of ion propulsion to the corrections in the position of geostationary satellites L'Aeronautique et l'Astronautique - n O 8 and nOlO - 1969 8. - C. DURANTE - A. COSTES - JC. LAPRIE

17. - J.R . ANDERSON - S.A. THOMPSON ( Hughes Research Laboratories) .

Xh

Development and long life performance of ion en g ines for satellite control . A.I.A.A. Paper 66 - 234 March 1966

A suboptimal system for the attitude control of an ion propelled satellite Automatic control in Space - 3rd IFAC Symposium - Toulouse, March 1970

391

24. - a) R. SAUVAGNAC x

18. - J .R. ANDERSON, G.A . WO RK (Hughes Resear ch Labo rat o ri es)

Protection aga inst the radiation of an iso topic so urce of 20 W th. in Ac 227 - Pu 238 - Pm 147 Note CEA DR / PR n O 271 - Ap r il 1968

Ion beam deflection for thrust vector control AlAA Paper 66-204 - March 1966

b) M. COHEN x

19. - R.E . HUNTER - R.O. BARTLETT ( NASA G.S .F .C.) R.M. WORLOCK - E.L JAMES (Electro Opt ica Sys terns)

Technical note on Actinium No te CEA DR/PR - April 1968 C.E.N . - Saclay

Cesium contac t i on microthruster experi ment aboard App lic at i ons Technology Satell ite (ATS) IV - AlAA Paper 69-297 - March 1969

(D~partement

des Rad i o-

~ l ~ment s)

B.P. n O 2

~

9l - G1F-sur-YVETTE

25 . - J . SURUGUE - J. FABR1 - E . LE GR1VES x 20 . - R. MORIN - C. MAR~ON - P. GUER I N M. LAV 1LLE - C. COUDURIER x

Prelimi nary Design Report - Cesium Con tac t ionization electrostatic thrusters No te ONE RA (Direction En erg ie et Propulsion) n O 7/6028 EY (contract CNES 68 - 244) - June 1969

Ammo nia re s istojet (S NECMA - 33-Blanque fort) Synthesis report a - n O VERG - 31884/48 (conaact CNES 67-282) - Sep t ember 1968 b - n ° VERG - 69 - 8038 (contract CNES 67 - 282 / 1) - Jul y 1969

ONERA - 21 , Av. de l a Division Le clerc 92 - Chati ll on

21 . - J. REBIERE - J.F. BARTHELEMY - L. GRUME L J . PETRES x

26. - G. ESCHARD - C. LOTY - A. PEL1SS1ER

- Ammonia resistojet - Study o f the dis soc i ation and of the thermal exchanges in ammo ni a durin g l amel l a r flows.

Genera l studies on the systems of a cesium contact ionization thru ster Monthly notes - 1969 (contract CNES 69 - 212) L .E. P. - 3, av . Descartes

- No tes : CEN - G/ A.S .P. n OOl / 69 - January 1969 (contract CNES 68 - 243) CE N- G/A.S.P . n 069 / l 6 - July 1969 ( contract CNES 68 - 243/1) - C.E. A. CENG - 38- GRENOBLE 22 . - R. MORIN - G.L. RIDEAU - J . C. GROSSET1E

x

94 - Limeil-Br~vannes

27. _ P. LUQUET x

x Utilization of ammonia microthrusting for the s tabili zing of satellites.

a - Comparative st ud y of four radioiso topic sou r ces Note SNECMA n O VER- 318 / r ( contract CNES 68-245) October 1968

Note CNES n O PR / ED / PS - 69 -T - 66 October 1969

b - Proposal for the manuf ac ture of an ammonia radioisojet prototype using actinium 227 as the source of h eat (RAMAC) Not e SNECMA n O VER-69-80l4 - Ap ril 1969 SNECMA - Missiles and Space Division 33 - B1anquefort 23 . - A. PLEDET - R. BESSE - R. CARON P. ADL1N - J . HACHE x Radiothermal thruster - Comparative s tudy of certain radioisotopes wi th a view to a satellite stationkeeping mission by ammonia radioisojet Note SEPR n O 11 89 - 68 V (contract CNES 68-266 - Nov ember 1968 SEPR - 3 aven ue du G~n~ r al de Gau1 1e 92 - PUTEAUX

392

x

in French

xx

in French and i n English

I>,'

{,

,;:. c::·, uir'J ,; · .:iu

Orbit: radius

42 ,165 km; period G6 ,164 si

fll;

( , J,r;cr,ej ill ~ "C'df~

Velocity: 3.075 km/sec •

.2. d... : angle corresponding to the dUration of each North and South correction

t.V

<:Jnnual

(m/s)

8c

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Large

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z0..Vo

50m/s/year

d.. sin 0{

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70

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f:!,

.(0

t

{4 (hours/day)

Fi O. I - L<:Jtitu·k correctiollS : necessar-y ve :ocity increments, with respect to the (jur-crt ion of tho North and South correctioll (3)

I

I

WEST

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__ _-;L!._~9

~ 1~ - __~d~

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~-

J

~(

stabili ze,t ion by amn-onia thrusters

t ,

t~ORTH a_

Stab! lization by ion thrusters a - single-si it thrusters b - single- \:)""'\-\-0,\ thruster

Fi g . 2- Sta b i! ization dovices -- - - (3- ax i s orbit ilnd attitude corrections)

394

U1

V)

'"

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Sub l irroblc (I sp = 50 sec.) 50 1 id

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~

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Honopropellant (Isp

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230 sec.

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Bi-propellant

~ =

I on thru ster (Isp

(lsp

300 sec.)

= 5,000

sec.)

Fi g . 3 ~~S 5 of c r b it cor r ec tion system of a sate l I ite for various types of - - - pro pu I ", i OI1 1:,0 5'-, or ~;'J'ie l! ite t150 kg Total i mpuls i o n 27 ,500 N.s/year F'OI'.'0 1" :';"' ,;; ' (:0

( Ll V = 61 m/ s ( so l <:: I' pa nels 0 .1 36 kg/ watt)

39 5

x yea r )

Orb it ·

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!.U

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Thenna I . .!>rO(o.j~ ,res i st oj et

Fig. 4

- Anmonia resi stojct s - Pri nc i p l e of o:)cra tion

396

(6)

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.

.--~

,

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ammonia resistojet (20)

Continuous operation Thrust : 5 m N POl'-or : 20 W Specific impulse expected

397

200 s

- Princi p le of ope ration of an ion thruste r I.propellant tank: 2 .Vaporizer 3. Source of ions; 4. acce lerating grid (H.T.

<0);

5. neutra I i zer'

- Ionization by electron

- Cesium contact of ionization thruster

bombardment

I. ano~ 2. cathode; 3. magnetic coils or perma nent magnets: 4. screen-grid; 5. accelerating grid; 6. Hg or Cs vapors

Ion thruste rs processes (7)

I. 3. 4. 5. 6. 7.

Cesium storage ; 2 . vaporizer; porous tun gsten ; focussing el ectrode; accelerating el ectrode; dece le ratin g e lectrode neutra I i zer

principle of operation and

398

ionizing

Fig. 7 - Ion thruster by me rcury

diameter 50 mm; thrust

399

elect ron 2 mN (oNERA)

bomba rdmen t ;

MultLbu..trol"l thruster (type Sastrugi) I. Ionizer; 2. acceleration grid

Multi-si it thruster; I. Accelerating electrodes.

Single-51 it thruster (electrostatic deflection of beam towards A or A')

Sing Ie- b",,1"ron thruster (electrostatic deflection of the beam toward A, AI I B or B') I. Acceleration electrode

Fig. 8 - Diagram of various types of ceslum contact ionization thrusters (7)

400