[1]0 8 1 2 1 5 , 3 6 5 Utfited States Patetit Office Patented Nov. 2, 1965 2 Other objects and many of the attendan t advanta ges of this inventio n will be readily apprecia ted as the same become s better underst ood by referen ce to the followin g detailed descripti on when consider ed in connecti on with the accompan yindrawings, wherein: FIG. 1 is a schemati c view showing a typical orbit alteratio n accordin g to the instant inventib nl with at accompa nying change in orbital plane; FIG. Z is a schemati c view showing a typical orbit al10 teration accordin g to the instant inventio n, without an accompa nying change in orbital plane; FIG. 3 is a graph showinthe angle bet,,,@e en the thriist axis and the orbital plane for various orbit inclinati ons of a magneti cally stabilize d satellite; 15 FIG. 4 is a schemati c view showina a represe ntative maneuv er in changin g a satellite' s orbit; FIG. 5 is a perspect ive view of the satellite used to practice the instant inventio n, showing particula rly the apparatu s used to despin and magneti cally orient the 20 satellite; FIG. 6 is a side elevati on of the satellit e used to practic e the instant inventi on' showi ng particu larly the appara tus used to alter the o;bit of th, sat,llit.; FIG. 7 is a schem atic diagra m of the propel lent feed 25 system used in the instant inventio n; FIG. 8 is a block diagra m of the electr onics in a satellit e used to Practic e the instant inventi on; and FIG. 9 is a scbem atic view showi ng the use bf spinup techni ques possibl e with the instant inventi on. 30 Briefl y, the inventi on conter nplate s the launch ing of a satellit e carryi ng an orbit impro vemen t system into the niaxim um orbit possibl e, accord ing to the capabi lities of its launch ing vehicl e. Then, alignm ent of the satellit e with the earth's magne tic field is accom plishe d by an 35 ele ctr om ag net po siti on ed ab oar d the sat elli te. Th e ma gn et has its dip ole ax; s par all el wit h the sat elli te axi s of syi nm etr y, wh ich is als o the thr ust axi s an d the spi n axi s. To ass ur e clo se ali gn me nt wit h the ear th's ma gn eti c fiel d, the sat elli te is als o de sp un pri or to acti vat ion of the ele c40 tro ma gn et by ire ans of the de spi n ro ds de scr ibe d by R. Fis ch ell in his U. S. pat ent ap pli cati on ent itle d "M ag net ic De spi n Me cb ani sm, " Se rial No . 83, 603 , file d Ja nu ary 18, 196 1, an d no w U. S. Pat ent No . 3,1 14, 518 . Th es e de spi n ro ds are ad diti on all y us ed to da mp out osc illat ion s ab out 45 the loc al ina gn oti c fiel d on ce ma gn eti c ca ptu re has oc cur re d. T h e s a t e ll it e is tr a c k e d b y a n y tr a c k i n g m e t b o d t o d e t e r m i n e it s o r b it a l p a r a m e t e r s. F r o m t h e s e p a r a m e t e r s a n d t h e V e st i n e m o d e l o f t h e e a rt h 's m a g n c ti c fi e l d , it 50 is possibl e to deter mine the time at wbich the stabiliz ed axis will have a particu lar attitud e with respec t to the earth. The import ance of this predic tability is that the accura cy of the system depen ds on the measu remen t of time which can be very accura tely measu red. 55 Once the actual orbital param eters of the satellit e are deter mined, the nature and extent of the first velocit y chang e is select ed in order to. alter either the satellit e's period , eccent ricity, inclina tion, nodal longitu de, or argurn ent of perige e or any combi nation of these param eters. 60 Howe ver, it is to be under stood that the chang e need not be made - in one step. On the contra ry, the orbit alterati ons may be perfor med in one or more stages, depen ding on its magnit ude, compl exity, precisi on, and allowa ble time for accom plishm ent. Additi onally, the chang e may 65 result in a brakin g effect upon the satellit e and cause the reentry and recov ery of the satellit e on the earth or a remov al to an interm ediate orbit from its previo us orbit. The sele cted velo city chan ge is insti g.ate d at a certa in geo grap hical and orbit al locat ion of the satell ite whe n 70 the stabiliz ed axis of the satellit e, on which a rocket motor is mount ed, is in the right positio n to give the requir ed added directi onal irnpuls e upon firing of the rocket . For 3,215,365 SPACECRAFT PROPULSION CONCEPT Theodore Wyatt, Charles J. Swet, and John Dassoulas, Silver Spring, Md., assignors to the United States of America as represeiited by the Secretary of the Navy Filed Apr. 30, 1963, Ser. No. 277,671 9 Claims. (Cl. 244-1) This invention relates to satellite orbit alteration techniques and, more particularly, to an adjustment technique which makes a plurality of successively more refined changes in a satellite's orbit until the exact orbit required is attained. Attaining an intended orbit about the earth for a satellite is a complex problem. The problem is further complicated by the use of a launch vehicle at its niaximum wei.alit lifting capacity. Occasionally, a rocket develops less total impulse than expected and/or exhibits guidance deficiencies, threatening the attainment of the satellite's intended orbit, and the entire usefulness of the satellite's mission. The orbit deficiencies may appear as a difference in the orbital period, eccentricity, inclination, nodal longitude, or argument of perigee, or as any combination of such differences. They may be fairly small, reflecting only the dispersion about nominal booster performance, or they may be substantial, caused by some unexpected aberrance. The orbit deficiencies may appear as a recognized altitude penalty imposed by inadequate booster impulse, or they may be deliberate where the mission calls for a succession of different orbits or a final de-orbiting nianeuver. Clearly, many planned and unplanned situations can arise which call for orbit alteration, the precise nature and extent of which cannot always be determined by prelaunch analysis. The instant invention has been designed to operate in combination with a satellite magnetically oriented within the earth's inagnetic field. The predictability of the earth's magnetic field is used as a means to determine the attitude of the satellite with respect to its orbital plane and to provide the correct combination of attitude and position for adding an additional irnpulse to shift the satellite to a new orbit. A considerable amount of original research effort has been expended in developing a model of the earth's magnetic field. This work bas successfully reached a point where it is possible to predict the pitch attitude of a magnetically oriented satellite in a given earth orbit. The pitch attitude is the angle in the orbital plane between the stabilized axis and a tangent to the instaintaneous fli.-ht path. The Vestine model of the earth's magnetic field bas been accepted by knowledgeable scientists as an accurate model. Therefore, it is felt that it isn't necessary to completely describe how the satellite's attitude changes as it orbits the earth. However, a basic explanation is given hereinafter. Since the invention is designed to operate within a predictable magnetic field it is not lirriited to the earth's magnetic field, but will function with any celestial body having a determinable magnetic field. One object of this invention, therefore, resides in the provision of a metbod for satellite orbit alteration which will operate within a known or determinable inagnetic field. A further object of this invention is to provide a method of satellite orbit alteration for manetivering a satellite in space within a given orbit or into new orbits. A still further object of this invention is to provide a method of satellite orbit alteration that does not destroy the compatibility of the satellite with its launching vehicle, and which will increase the efficiency of the entire launching system.
[2]3 example, the application of an added impulse at an oblique angle to the plane of the orbit results in a change in altitude, period, eccentricity, inclination and nodal longitude, or a combination <)f less than all, depending upon the angle between the path of flight and the impulse axis as seen in FIG. 1, while an application of impulse witbin the plane of the orbit results in a change in altitude, period or eccentricity, as can be seen in FIG. 2. FIGS. I and 2 are schematic representations of an orbital path of a satellite circling the earth I in an orbit 2 and the velocities associated with the moving satellite. Vo represents the steady state velocity of the satellite, AV Tepresents the velocity increment added by use of the instant invention, resulting in a new satellite and velocity Vl. However, when selecting the point at which an impulse will be added, it is important not only to determine the pitch angle as hereinbefore mentioned, but also to determine the out-of-plane relationship of the magnetically oriented satellite in respect to its orbit. FIG. 3 shows the various out-of-plane angles between the thrust axis Of a satellite and the orbital plane, for diverse angles of inclination. For example, a satellite in a polar orbit, having an inclination of 90 degrees, has its thrust axis vary approximately 25- degrees from coplanar. That is, as the satellite passes in a complete orbit, it appears tO wobble about its ma.-netically stabilized axis. However, a satellite launched intb an orbit with 60 degrees inclination will wobble only on one side between approximately 5 and 55 degrees. Since the thrust which alters the inclination and nodal longitude varies with the cosine Of the out-of-plane angle, while the thrust which alters the period, altitude and eccentricity varies as the sine of the out-of-plane angle, the correct out-ofplane angle must be selected during the orbit which has a sufficient in-plane component to attain the new desired orbit. From FIG. 3, it cari be seen that bighly inclined orbits above approximately 67 degrees provide a coplanar thrust angle corresponding to the desired direction of the velocity increment to change the attitude or period of the orbit alone. For satellites having lesser angles of inclina-tion, the in-plane component will alter the altitude period .and eccentricity, while the out-of-plane component will alter the inclination or nodal longitude. However, since these orbits never have a coplanar component, it is not possible to change only the period, altitude or eccentricity alone, bilt only with an associated chan,-e of nodal longitude and inclination. Highly inclined orbits of greater than approximately 67 degrees are most likely to provide pitch angles corresponding to the desired direction of the velocity increment, to change the altitude or period of the orbit only. The position of a specified point in the orbit with respect to the earth, such as perigee or apo.-ee, chan.-es continually due to the precession of perigee. Thus, the simultaneous satisfaction of the above-mentioned three conditions ofout-of-plane angle, pitch angle and point within the orbit can be obtained over a finite range of values for a given orbital inclination by selection of a calendar date and a time on the date. Since perigee precession is fairly slow, the ri.-ht conditions will recur for several times during consecutive orbital revolutions. Conversely, it may be necessary to wait for a considerable 'period for th@, desired amount of perigeeprecession to occur. This disadvanta.-e is removed by spinning up the satellite having the proper out-of-plane angle and pitch angle under magnetic control, thereby preserving these ,angles by a gyroscopic stabilization until the perigee position or other proper point in the orbit is obtained. After a time, the spin would be damped out by the electromagnetic despin technique and magnetic attitude control would be restored, and, if desired, the process could be repeated as often as required to attain the proper combination of position and attitude. The major advantage of this temporary "freezing" of a magnetic attitude and of the subsequent transferrin.- of the attitude to another point in the orbit for application 3,215,365 4 of a velocity increment is that time necd not be wasted in waiting for the proper point in the orbit to precess around to the proper attitude for rocket firing, but, rather the attitude can be carried around to the firing point. 5 Referring to FIG. 4, there can be seen a schematic representation of a magnetically oriented satellite 3 orbiting the earth. A method for magnetically orienting a satellite is completely described by Fischell et al. in their U.S. patent application entitled "Magnetic Attitude Conio trol," Serial No. 99,644 filed March 30, 1961, and now U.S. Patent No. 3,118,637. It is sufficient for the understandin.- of this invention @o briefly describe magnetic orientation by stating that the satellite carries magnets which align themselves, and thereby the satellite, with the 15 earth's magnetic field, and @therefore cause the satellite to tumble in the plane of the magnetic field. The angle between a tan,@ent to the orbit and a line through the thrust axis is called the pitch angle. The satellite 3 has been launched into a low orbit des20 ignated by a line 4. This orbit is shown as a circular orbit, but any orbit configuration would work equally as well. Additionally, the satellite 3 is shown in several different positions demongtrating the above-mentioned tumbling effect of the magnetically stabilized satellite 25 as it aligns itself with the earth's magnetic field represented by a dashed line S. Prior to attaining a magnetically oriented satellite, the satellite must be despun. A suitable method for despinning such a satellite is described by R. Fischell in his 3o abovementioned U.S. Patent No. 3,114,518. It is sufficient for the understanding of the instant invention to indicate that a plurality of permeable rods are mounted on or within the satellite so that they lie in a plane or planes that are perpendicular to the magnetic axis of 35 stabilization. These rods may be grouped, with various groups laying at right angles to each other. The permeable rods will dissipate the rotational energy by providing a type of magnetic braking effect. A satellite located in the orbit 4 may have been in40 tended to orbit at a higher altitude which it failed to attain. As previously described, when the correct pitch angle and in-plane component of the thrust vector are pres,nt and the satellite is at the perigee of its orbit, it is possible to add thrust which will increase the apogee from its present position 7 in orbit 4 to a new location 9 45 in an orbit 11. However, by increasing the apogee of the satellite's orbit, the circular orbit has been changed to the elliptical orbit 11, although the new apogee is now at the desired altitude. It is now possible to change the new elliptical orbit into 5 0 a circular obit, keeping the altitude of the apogee 9 by thrusting again when the satellite has a zero pitch angle, but this time when the satellite is at the apogee. This additional thrust will raise the perigee until the orbit is essentially a circular one 13. 55 The amount of thrust to be added first at the perigee and later at the apogee can be determined by using mathematical solutions well-known in the art of astrophysics. Referrin.@ to FIG. 5, there can be seen a perspective Go view of a satellite which is used to practice the instant invention. Disposed within the satellite frame 15, a 'plurality of despin rods 17 are grouped together and lie in a plane or planes. As shown, pairs of spaced rods lie at right angles to each other and in intersecting relation65 ship. A plurality of solar cell panels 19 are symmetrically arranged about the satellite to supply electrical energy to a plurality of batteries 21, which among other things are used to activate a plurality of electromagnets 23. The magnetic dipole axes should be parallel with the, 70 satellite axis of symmetry, to align that axis with the earth's magnetic field. The disposition of the electroma.,nets shown in FIG. 5 causes the Ion.-itudinal axis 25, to align with the earth's magnetic field. A plurality of spinup nozzles 27, two of which are shown, are sym75 metrically located upon an adjustment mod-4!q 39, aqd
[3]5 iipon activation will impart rotation about the magnetically stabilized axis 25. The spin imparted to the satellite is used to stabilize the satellite While thrusting; also for settling liquid propellants for zero-g starts, thereby eliminating the need for bladders, and for gyroscopically holding a magnetically acquired attitude for later use. The adjustment module 29 shown in a cut-away view in FIG. 6 is symmetrically constructed, thereby preventing any unwanted change in the satellite attitude once the satellite has assumed the proper attitude, and the satellite is spun-up to preserve that attitude for a later position in its orbit. The system contains a pair of hemitoroidal tanks 40 and 42. A fuel such as hydrazine may be stored in the tank 40, and an oxidizer, such as nitrogen-tetroxide, may be placed in the tank 42. Each tank contains an integral charge of pressurizing gas such as nitrogen for propellant feed, introduced prior to launching through a pressuiization connector which is not shown. A fuel line 44 conducts the fuel from the tank 40 by an inlet 45 to an injection system 46, and an,oxidizer line 48 conducts the oxidizer from the tank 42 by an inlet 49 to the injection system 46. The inlets 45 and 49 are positioned within the tanks so as to remain below the fuel and oxidizer levels represented by lines 50 and 52 respectively. The fuel and oxidizer levels assume the configuration shown in the FIG. 6, under spinup forces created by the venting of gases, which gases are stored in a toroidal pressure chamber 54, and released by means of a plurality of pressure lines 56 and nozzles 27. The chamber 54 is symmetrically disposed about the lon,@itudinal axis 25 and is attached to the tank 42 by suitable means 58. The fuel and oxidizer react in the combustion chamber 66 of the rocket motor 60 and provide thrust along the axis 25. The metering system shown in FIG. 7 has the same arrangement for both the fuel and the oxidizer input portions, although only the fuel input portion is shown. The fuel enters a normally closed explosive valve 62 through the fuel line 44. The explosive valve prevents fuel from passing into the motor prior to the first commanded thrusting period. A suitable explosive valve is that identified as type 1, Model Class A-4 maniifactured by the Conax Corp. The fuel is filtered in a filter 64 and enters the combustion chamber 66 of the rocket 60 through a solenoid valve 68. Referring again to FIGS. 6 and 7, each propellant tank 40 and 42 is initially fflled half or two-thirds full, then the ullage space is charged with all of the pressurizin@ gas that will ever be used prior to launch. When thrust is first commanded, after spin has whirled th- liquid levels out to form cylindrical surfaces 50 and 52 and immersed the propellant feed inlets 45 and 49, the sqtiib-actuated valves 62 are fired and the solenoid valves 68 are opened. The pressurized propellants are then forced through the metering system 46 to the combustion chamber 66, where they ignite hypergolically. The axial thrust then produces an acceleration which re-orients the liquid level as indicated by the dotted lines 69 and 70, in FIG. 6, with the propellant feed inlets 45 and 49 still immersed. Thrust continues, at a gradually reducing level as the gas pres,sure drops with increasing ullage volume, until the solenoid valves are signalled to close. This thrusting period produces a certain total impulse @vhich corresponds to the desired velocity change. Through knowledge of the system characteristics and a record of cumulative burn time, each velocity change can thusly be controlled by the preselected duration of valve opening, or burn time. ,It is to be noted that this system is not Emited to instances requiring substantial velocity changes requiring a large rocket. On the contrary, orbit alterations may be required which need only modest propulsion forces. The choice of the type of propulsion employed is heavily dependent upon the magnitude of the orbit alteration de3,215,365 sired, the accuracy desired, the necessity for repetitive operation, either for vernier corrections in order to obtain high accuracy or to obtain more than one major alteration of the orbit, and the maintainability of the propulsion system in the space environment. Thus, applicable propulsion techniques may include bi@propellant liquid rockets, monopropellant liquid rockets, solid propellant rockets, cold gas jets, and small explosions or slug ejections. 10 The electronics required in the instant invention are shown in block diagram in FIG. 8, and include a very stable oscillator 71. A suitable oscillator for the purpose is that described by J. B. Oakes et al. in their article entitled "D@-sign of an Ultrastable Oscillator for Satellites," 15 in the September-October 1962 issue of the Technical Digest published by the Applied Physics Laboratory of The Johns Hopkins University. The oscillator 71 has three output signals; one is applied to a Doppler transmitter 72 which furnishes a signal to be transmitted by 20 an antenna 74 to provide a means for Dcyppler tracking of the satellite. The second output of the oscillator 71 is applied to a suitable memory 76 and controls the filling of the memory and the reading out of stored information for the delayed execution of command sig25 nals. The third si,-nal fro-m the oscillator is applied to a clock 78 which is used in conjunction with the memory 76 and connected thereto, to determine when the stored command signals will be executed. Command signals are received by an antenna 80 and 30 separated from their carrier in a -command re--eiver 82. The:command receiver 82 applies its output to a demodulator 84 which recognizes the incoming information signals and changes their form for application to the memory 76. The memory 76 applies its output signals to a regulat35 ing circuit 86 which may activate the electromagnets 23 the spin-up jets 27, or the thrust rocket 60. The ma-netically oriented satellite 3 is shown tumbling in the earth's magnetic field in FIG. 9, having a polar orbit $8. By using the Vestine model of the magnetic field, it 40 is determined that the satellite will have the proper pitch in -position 89 which must be coupled with the perigee at position 90 in order to provide an in-plane component of thrust to chan.-e the present orbit to a new orbit 91. As the satellite 3 reaches position 92, an earthbound 45 injection station 93 fills its memory system 76, as seen in FIG. 8, with command signals for determining the sequence of the following events. The satellite continues to turnble in the earth's magnetic field thereby continuously changing its pitch angle. Computations involving the ro satellite's orbital parameters and the Vestine model of the earth's magnetic field show that the rocket should be spun-up when it reaches position 89 where it possesses the desired magnetic pitch orientation. The memory system signals the spin-up jets, deployed about the rocket, 5,5 to ignite and to impart rotation about the satellite's thrust axis thereby gyroscopically maintaining the desired pitch attitude of the satellite. The attitude is frozen as the satellite continues in its orbit to position 90 at which time the memory signals the thrust rocket to operate and 60 add the thrust required to alter its present orbit to a new orbit 91. Obviously, many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within 65 the scope o'L the appended claims the invention.may be practiced otherwise than as specifically described. What is
