CA2183522A1 - Elliptical orbit satellite coverage and deployment system - Google Patents
Elliptical orbit satellite coverage and deployment systemInfo
- Publication number
- CA2183522A1 CA2183522A1 CA002183522A CA2183522A CA2183522A1 CA 2183522 A1 CA2183522 A1 CA 2183522A1 CA 002183522 A CA002183522 A CA 002183522A CA 2183522 A CA2183522 A CA 2183522A CA 2183522 A1 CA2183522 A1 CA 2183522A1
- Authority
- CA
- Canada
- Prior art keywords
- orbit
- satellite
- satellites
- inclination
- coverage
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Classifications
-
- H—ELECTRICITY
- H04—ELECTRIC COMMUNICATION TECHNIQUE
- H04B—TRANSMISSION
- H04B7/00—Radio transmission systems, i.e. using radiation field
- H04B7/14—Relay systems
- H04B7/15—Active relay systems
- H04B7/185—Space-based or airborne stations; Stations for satellite systems
- H04B7/18576—Satellite systems for providing narrowband data service to fixed or mobile stations, e.g. using a minisatellite, a microsatellite
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
- B64G1/1007—Communications satellites
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
- B64G1/1085—Swarms and constellations
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/242—Orbits and trajectories
-
- H—ELECTRICITY
- H04—ELECTRIC COMMUNICATION TECHNIQUE
- H04B—TRANSMISSION
- H04B7/00—Radio transmission systems, i.e. using radiation field
- H04B7/14—Relay systems
- H04B7/15—Active relay systems
- H04B7/185—Space-based or airborne stations; Stations for satellite systems
- H04B7/195—Non-synchronous stations
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/002—Launch systems
Abstract
A special set of elliptical satellite orbits are described which allow preferential coverage of one parameter over another. According to a first modification, the orbits are retrograde, and preferentially cover one geographical location or time of day as compared with another. A
second modification uses prograde orbits and allows the apogee of the orbit to be offset a constant amount with respect to the sun, to thereby cover a different time of day relative to the others. According to a special preferred mode of the invention, the apogee is always pointing towards the sun.
second modification uses prograde orbits and allows the apogee of the orbit to be offset a constant amount with respect to the sun, to thereby cover a different time of day relative to the others. According to a special preferred mode of the invention, the apogee is always pointing towards the sun.
Description
W0 95/22489 r~
-: 21 83~22 Elliptical orbit satellite coverage and deployment ~ystem.
This i6 a continuation-in-part of application 5 number 07/892,239 filed June 2, 1992, pending.
FIELD OF THE INVENTION
The present invention relates to elliptical satellite orbits, constellations, methods, and communication systems.
10 BAc:r~iK~ UNI~ OF THE INVENTION
The concept of artif icial satellites circling the earth was introduced to scientific literature by Sir 5 Isaac Newton in 1686. Things have gotten considerably more complicated since that time, however. The basic concepts of an orbit are described in any orbital mechanics or a~L. udy~ lics textbook, such as "Fl~nll Lals of Astrodynamics" by Bate et al. or 10 "Orbital MPt-hAn;t-CII by Chobotov, AIAA Education Series, pl~hl ;ch~r. The following definitions of these terms will be first provided here, since they are nec~Cc;~ry for proper und~rstanding of the present invention.
The earliest satellites placed into space by 15 man were deployed into very low circular orbits. The resulting visibility footprint of one of these satellites was quite small and a single satellite had the added disadvantage of providing only a few minutes of coverage per day. In fact, it was quite common for an observer on 20 the equator to miss being in contact with such a satellite for several days. Raising the satellite to a higher orbital altitude (e.g., #600 nautical miles) helped extend both the coverage footprint, average viewing elevation, and the time in view, but for some Wo ss/22489 r~
missions El ~uuent or even continuous coverage became a requirement. This led to the deployment of early multiple satellite systems, a typical example being the Navy' 5 Transit navigation satellite system. Satellite 5 systems designers were increasingly asked to provide continuous cuv~l~ge:; first, for latitudinal zones and then, for the entire globe.
One of the first constellation designers to study zonal ~;UV~ U~ was David L~iders. The Englishman, John 10 Walker, was the first to systematize the design of multiple-ring, multiple satellites per ring, constellations and his work contributed greatly to the optimization of a number of multi-satellite systems (e.g., NAVSTAR GPS). A Russian designer, G. Mozhaev, 15 in~lDrr~nAr~ntly came up with similar arrays using a more theoretical approach based on mathematical sets and group theory. Polar constellations often employed the concept of "street-of-coverage", and further coverage uv~ Ls were made by Beste, Ballard and Rider. More 20 recently, Hanson and Linden have investigated large arrays of low earth orbit "LEO" satellites (40-200 satellites). All of these designers employed circular orbits; and even with this simplification, constellation design was cnnRirlr~red at best a difficult and time 25 conc~lm;n-J trial and error exercise.
The motion of any artif icial satellite may be described using a number of parameters. The eccentricity, e, is a measure of the amount of ellipticity. An orbit which has a greater eccentricity 30 number is more elliptical. Eccentricity e=O would describe a circle, any number between O and 1 is an ellipse, and the eccentricity number of 1 or greater would be a parabola or a hyperbola, respectively (curves which never close).
wo ss/22489 P~~ '02ll3 '''~' 2 ~ 3 3 5 22 For an elliptical orbit, the earth, or the object being orbited, is at one of the focal points of the ellipse. Therefore, the satellite is sometimes closer to the earth than at other times. The apogee is 5 defined as the point of highest altitude of a satellite, while perigee is the point of lowest altitude.
A reLL~yLade orbit is one in which the direction of revolution is opposite to that of the earth.
A posigrade or prograde orbit is an orbit in which the lO satellite revolves around the earth in the same direction as the earth.
The inclination angle i is an angle measured between the plane of the orbit, and a plane of the reference, usually the Equator. An inclination angle i 15 less than 90~ is a prograde orbit, while an inclination angle greater than 90 is a r~LLoyL~de orbit. A 90 orbit is a polar orbit.
The period, T, is a measure of how long the satellite takes to make one entire orbit. Mean anomaly M
20 is another way to describe the position in the orbit.
Mean anomaly is a f ictitious angle indicating the fraction of 360 degrees C~LL~ ing to the fraction of the period through which the satellite has passed at any point of its orbit.
25 The Right Ascension of the AcG~nAin~ Node ( "RAAN" ) i an angle between the f irst point of Aries (y), a non-rotating celestial reference, and the line of nodes, which is the line forming the intersection of a plane of the orbit and the plane of the equator. The 30 line of nodes gives a measure of the position or orientation of the orbit. The longitude of the Accpn~
node n is the angle between the i unit vector (pointing towards the Greenwich meridian) and the Aco~nrling node, in the rotating ref erence .
-: 21 83~22 Elliptical orbit satellite coverage and deployment ~ystem.
This i6 a continuation-in-part of application 5 number 07/892,239 filed June 2, 1992, pending.
FIELD OF THE INVENTION
The present invention relates to elliptical satellite orbits, constellations, methods, and communication systems.
10 BAc:r~iK~ UNI~ OF THE INVENTION
The concept of artif icial satellites circling the earth was introduced to scientific literature by Sir 5 Isaac Newton in 1686. Things have gotten considerably more complicated since that time, however. The basic concepts of an orbit are described in any orbital mechanics or a~L. udy~ lics textbook, such as "Fl~nll Lals of Astrodynamics" by Bate et al. or 10 "Orbital MPt-hAn;t-CII by Chobotov, AIAA Education Series, pl~hl ;ch~r. The following definitions of these terms will be first provided here, since they are nec~Cc;~ry for proper und~rstanding of the present invention.
The earliest satellites placed into space by 15 man were deployed into very low circular orbits. The resulting visibility footprint of one of these satellites was quite small and a single satellite had the added disadvantage of providing only a few minutes of coverage per day. In fact, it was quite common for an observer on 20 the equator to miss being in contact with such a satellite for several days. Raising the satellite to a higher orbital altitude (e.g., #600 nautical miles) helped extend both the coverage footprint, average viewing elevation, and the time in view, but for some Wo ss/22489 r~
missions El ~uuent or even continuous coverage became a requirement. This led to the deployment of early multiple satellite systems, a typical example being the Navy' 5 Transit navigation satellite system. Satellite 5 systems designers were increasingly asked to provide continuous cuv~l~ge:; first, for latitudinal zones and then, for the entire globe.
One of the first constellation designers to study zonal ~;UV~ U~ was David L~iders. The Englishman, John 10 Walker, was the first to systematize the design of multiple-ring, multiple satellites per ring, constellations and his work contributed greatly to the optimization of a number of multi-satellite systems (e.g., NAVSTAR GPS). A Russian designer, G. Mozhaev, 15 in~lDrr~nAr~ntly came up with similar arrays using a more theoretical approach based on mathematical sets and group theory. Polar constellations often employed the concept of "street-of-coverage", and further coverage uv~ Ls were made by Beste, Ballard and Rider. More 20 recently, Hanson and Linden have investigated large arrays of low earth orbit "LEO" satellites (40-200 satellites). All of these designers employed circular orbits; and even with this simplification, constellation design was cnnRirlr~red at best a difficult and time 25 conc~lm;n-J trial and error exercise.
The motion of any artif icial satellite may be described using a number of parameters. The eccentricity, e, is a measure of the amount of ellipticity. An orbit which has a greater eccentricity 30 number is more elliptical. Eccentricity e=O would describe a circle, any number between O and 1 is an ellipse, and the eccentricity number of 1 or greater would be a parabola or a hyperbola, respectively (curves which never close).
wo ss/22489 P~~ '02ll3 '''~' 2 ~ 3 3 5 22 For an elliptical orbit, the earth, or the object being orbited, is at one of the focal points of the ellipse. Therefore, the satellite is sometimes closer to the earth than at other times. The apogee is 5 defined as the point of highest altitude of a satellite, while perigee is the point of lowest altitude.
A reLL~yLade orbit is one in which the direction of revolution is opposite to that of the earth.
A posigrade or prograde orbit is an orbit in which the lO satellite revolves around the earth in the same direction as the earth.
The inclination angle i is an angle measured between the plane of the orbit, and a plane of the reference, usually the Equator. An inclination angle i 15 less than 90~ is a prograde orbit, while an inclination angle greater than 90 is a r~LLoyL~de orbit. A 90 orbit is a polar orbit.
The period, T, is a measure of how long the satellite takes to make one entire orbit. Mean anomaly M
20 is another way to describe the position in the orbit.
Mean anomaly is a f ictitious angle indicating the fraction of 360 degrees C~LL~ ing to the fraction of the period through which the satellite has passed at any point of its orbit.
25 The Right Ascension of the AcG~nAin~ Node ( "RAAN" ) i an angle between the f irst point of Aries (y), a non-rotating celestial reference, and the line of nodes, which is the line forming the intersection of a plane of the orbit and the plane of the equator. The 30 line of nodes gives a measure of the position or orientation of the orbit. The longitude of the Accpn~
node n is the angle between the i unit vector (pointing towards the Greenwich meridian) and the Aco~nrling node, in the rotating ref erence .
2 1 8 3 5 2~
(. ! .
-- 4 -- ~ =
The ~ILyl -- of perigee ~ is an angle measured in the plane of the orbit between the point of the ~cPn~1in~ node and the nearest point of perigee.
Most practical satellites prior to the 5 invention by the present inventors used relatively simple systems based on circular orbits. The earth was covered sy_metrically by multiple satellites, which each operate to cover a section of the earth.
Elliptical orbits have been typically avoided 10 in the art, because of their asymmetries, and the c~-nCPTlPnt problems that they might cause. However, some individual elliptical orbits and elliptical orbit constellations have been ~l~,~osed. The Russian Molniya orbit is a posigrade orbit ~Riqnod for polar and high 15 latitude coverage. Other posigrade orbits have been described by John Draim in his U.S. Patents 4,809,935 and 4,854,527.
4, 809, 935 describes a three-satellite constellation giving continuous coverage of the entire 20 Northern hPmi~rhPre, and an extension of this constellation to include an equatorial orbit resulting in a four-satellite array giving continuous global coverage of both hPm; ~rhPres. This latter four satellite array provided somewhat higher elevation coverage in the 25 Northern hemisphere than in the Southern RPmi ~rhPre.
4, 854, 527 describes a common period four-satellite array giving continuous global ~:~Jv~:Lc~g~
with satellites at a lower altitude range than in the first patent. A rli~cllC~c;on of obtaining extra Northern 30 HPm; crl.Pre coverage through use of elliptic satellite constellations may be found in ANSER Space Systems Division Note SpSDN 84-1, "Satellite Constellation Design Techniques for Future Space Systems" dated September 1984, by John Draim, and James Cooper. Another 35 application of posigrade elliptic orbits is the ACE and Wo 95t22489 1 ~ JI~1II3 2i 83522 ACE-Prime orbits developed by Mr. A. Turner of Loral Corporation .
The present invention also simplif ies the design of the solar panels by requiring no more than 1 or 5 2 degrees of freedom. In the example orbit discussed herein which is 116 r ~LLvylade, the panels need only one degree of freedom. In a similar way, a satellite usually needs to radiate its heat toward cold, empty space. In the present invention, it is much easier to face the 10 satellite in a way that always faces the heat radiators away f rom the sun .
It is also well known that the earth is not totally circular, but actually it is rather oblate. That i8, the earth is bigger at the bottom than it is at the 15 top. The J2 harmonic, due to the earth's oblateness, causes the node n and argument of perigee ~ o~ an orbit to change. The gravitational pull of the earth's equatorial bulge causes, for example, the orbital plane of an eastbound satellite to swing westward. More 20 generally, the force component is directed towards the Equator. This resultant acceleration causes any satellite to reach the Equator (node) short of the crossing point where it would have reached it on a spherical earth. For each revolution, therefore, the 25 orbit regresses a ~ amount. These effects have been the subjects of various attempts at ~ ~tion.
Sun .-y,.~llrol.v~ s circular orbits are also known. These are orbits where the rotation rate of the right ascension of the Acc~ntl;ng node is equal to and in 30 the same direction as, the right ascension rate of the mean sun.
.
ST~TARY QF TT-TT' INvT~NTIoN
The previous specification, of which this is a continuation-in-part, described the invention of non-Wo 95n2489 ~ 5~02113 uniform capacity distribution tailored by latitude andpopulation. This was done using an elliptical satellite array. The present specification adds additional information to the basic elliptical orbit. A first 5 pmhoA;~~nt of the present invention is tailored by latitude and population . A second pmhoA i- ~ is also tailored by time of day. The present invention describes a specif ic way of carrying out these options by using a I~LLo~L4de intl;nPA orbit, and/or a sun ~y~ )us 10 apogee, elliptical orbit, which is equatorial, and also describes many ways of PYpAnAin~ the design space.
The parent specification achieved the ~:uy~ ,Iphical discrimination . The present specif ication v~c s on this basic technique by using the 15 pcLLuLl,ation to achieve various effects. The present inventors started with a r:~nrln;c~l form for the equation of motion of satellites in orbit. Two cases were di~cuvt:l ~d by the present inventors: 1 ) one parameter fixed at 116, to form a ~LL.,yLc,de sun 6ynchronous 20 orbit, and 2) both parameters varied to form a sun synchronous apogee orbit where the apogee always points in a constant direction relative to the sun. This resulted ir. an expansion of the possible space which allows time of day tailoring of ~OV~L~Y~. Such was 25 completely unheard of before the present invention, except for certain conditions 8uch as circular inrl ;nPA
orbit, or the ~ce equatorial orbit. The present invention, in contrast, allows an entire domain of applicability of these two conditions in order to obtain 30 various unheard of combinations. According to the present invention, two PmhoAir- Ls are described: A
first of which includes incl ;neA sun synchronous orbit and a second which includes constant pointing apogee orbits. Both of these are based on the effects on orbits 35 from the J2 term.
~ Wo gsl22489 r~
21 ~3522 ~
We have found a way to exploit the effects of earth's oblateness, with an elliptical orbit, such that for certain combinations of orbital parameters, the secular perturbations due to the earth ' S J2 gravitational 5 term may be used to advantage. Specifically, we noticed that the effects on the ~ and n terms could be used to cate the orbit of the satellite to obtain certain controllable effects. These controllable effects are obtained taking into account that the earth revolves lOrelative to the sun by 0.9856, of orbit around the sun, per day . The present invention uses p~:L LUL l,ations arising from the J2 terms, to precess the orbits in time with respect by following the general equation d~ + ddn =o . 9 856 . .. ( 1 ) to obtain elliptical orbits which have controllable and 15 constant characteristics, and which a-~y LLically cover one parameter of co~ r ~.ye preferentially over another in a way which is constant relative to the sun, all year round. According to a first ~mho~i--rt of the invention, that parameter of coverage is geographical location, and 20 according to another ~-- ir L of the invention, that parameter of coverage is time of day.
Kepler's law of motion governs any orbit. In any orbit, including an elliptical orbit, areas swept out in equal times must be equal. Hence, a satellite in an 25 elliptical orbit spends more time towards apogee than it does towards perigee. We have exploited this effect, to bias the effective orbits to increase coverage at certain geographical locations and/or times. According to the first aspect of the present invention, a first operation 30 is carried out to set the change in ~, d/dt(~) to approach 0. The change in n is then set to 0 . 9856 /day.
C '-in;n~ this with a r~LL~/yL~de orbit produces a set o~
Wo 9~l22489 A _ I / ~,J ", _`IA 1 ~ ~
2 ~ 8 3 5 2 2 specific characteri6tics referred to herein as a design space. This design space includes L~ yLCLde orbits which f avor one geographical location on the earth over another .
5 One recognition of the present invention is that the Northern T~m; srhore includes a majority of the world's earth masses and population. By covering the Northern T~~; crhore preferentially over the Southern T~pm; crhore, the coverage can be equalized as a function 10 of population .
A second ~~'~';- ~ of the present invention uses prograde orbits, using the more general equation l.
In the second omho~ t, d/dt(~) does not approach 0.
Instead, the parameters d/dt(~) and d/dt(n) are both 15 variable, but only in a way that meets the general equation dd~t ~ ddQ = 9856 .. (1 ) We have found that this produces an orbit where the apogee always points in a specif ic direction relative to the sun. This can be used to increase the satellite 20 daytime C~VC:L~ effect, or the effect from 9AM to 5PM
for example. Of course, more people require satellite services during business hours than at any other time.
Hence, there is more of a demand for satellite servicc during the day than there is at night.
25 After identifying the advantages possible from an elliptical orbit in this way, we have identif ied a technique of choosing parameters of orbits such that different areas will always be preferentially covered during the day.
30 The present invention teaches construction of a satellite orbit, a satellite system, a method of operating a satellite system, and a method and apparatus ~ W0 9~/22489 r~ 3 ;, . . 2 ~ 83522 g ~
of deploying a satellite into a prescribed orbit, all using orbital parameter combinations with integral or non integral mean motions specified within the design space that covers a specified set of earth (or planet) coverage 5 requirements in a more optimal manner than is obtained through more conventional orbits.
The objects of the present invention include:
Col~,LLu~;~ion of a satellite orbit using ; nrl; n~cl orbital parameter combinations within a design 10 space that extends the latitude ranges in order to meet earth (or planet) coverage requirements in a more optimal manner than are obtainable through conventional circular DUI~ .~y~ lLVll~US type orbits.
Cv...,~.ucLion of a satellite with ~IL~ ~ of 15 perigee value other than 90 or 270 degrees, such that the apogee locations may be preferentially oriented in any desired direction, preferably towards the earth-sun line, giving more extensive ( in both time and earth central angle) coverage, and such that i .,~,ed coverage during 20 daylight hGurs is achieved, than during nighttime hours, for locations at all longitudes from -180 to +180 degrees (or 180W to 180E).
Provision of the required satellite elevation angles within specif ied latitude ranges, with appropriate 25 day-night biases, for the r~ ,yLc.de elliptic orbit def ined .
Provision of a satellite orbit that maintains its integrity year-in year-out through precise orbital inj ection control so that coverage characteristics are 30 maintained ~hllJU~ uL the satellite constellation lifetime. Note: minor orbital adjustments may be required to account for smaller perturbations, e.g. third order or higher and/or solar peLLuLl,ations, which are experienced by the satellite.
The novel features of this aspect include:
WO 95122489 1 . ~ 13 ~ 1 8 3 522 greater satellite Earth coverage can be provided during the daylight hours (or bl~Ci nPc5 day, when there is heavy utilization of tel~ i cations or other useful servicec), 5 116.565 degree orbit plane inclinations, as described according to the f irst pref erred Pmhofl i - L
will provide continuous coverage of the high latitude and polar regions with elliptic orbits, not obtainable from equatorial plane orbits.
10 Relatively low orbits, which can be obtained using corrP~pon~;nq smaller rocket boosters.
~TF:F DESCRIPTION OF THE DRAWINGS
These and other aspects of the invention will now be described in detail with reference to the ac -nying 15 drawings, wherein:
Fig . 1 shows a f irst des ign space f or elliptical sun synchronous It:l.LoyLade orbits according to a f irst Qmhofl; L 0~ the present invention;
Fig. 2 shows the characteristics of a special 20 orbit according to a second pmhofl;r L of the present invention in which the apogee is always pointing towards the sun;
Fig. 3 shows a design space for this second pmhofl;-- L of the present invention using prograde 25 orbits;
Fig. 4 shows a constellation of satellites, each orbiting and - ; cating with earth stations on the earth;
Fig. 5 shows a rocket and inertial guidance 30 unit used according to the present invention to propel the rocket into orbit; and W0 9s/22489 p~ 3 ` ~ ~ ` ;i; 2 ~ ~3522 .. ~
Figs 6, 7A, 7B, 8A, 8B, 9A, 9B, 9C, lOA and lOB show characteristics of preferred orbits of the present invention.
l~E~ V~lON OF ~ l) EMF~ODJM~NT
5 The present invention exploits the gravitational effects from the earth's oblateness, in combination with a preferably elliptical orbit, to allow preferential coverage of different parts of the earth as a function of parameters which are related to satellite 10 demand. This has significant advantages since it allows preferential coverage based on a chosen characteristic, here either one h~m;~rhPre over the other, or time of day .
For instance, a satellite system primarily 15 intended for use over the United States would prefer to preferentially cover the Northern hPm; ~rhPre as opposed to the Southern h~mi~rhPre. More specifically, by choosing elliptical orbits such that anything above 40 south latitude was covered, a great majority of the 20 world's land mass could be covered without wasted capacity .
This ~ l of the invention optimizes the characteristics of the elliptical satellite to have desired ~ L C~ characteristics . According to this 25 first preferred mode, structure is described for putting a satellite in a special orbit which preferentially covers part of the earth over the other part.
The f irst type of orbits, discussed according to the present invention herein, are elliptical 30 L~lL~JyLnde orbits which provide preferential coverage of one part of the earth over ~he other part through adjustment of orbital parameters.
As mentioned above, all orbits are effected by the earth's J2 gravitational term. Thi~ term effects wo 95l22489 f~r ~r~ 21 83522 the n and ~ terms of every orbit. In order to compensate the orbit, the general equation dd~lt'+ dt=0'98---- --------. . . (1) must be satisfied. This first: `~o~; L takes a special case of the equation (1).
5 The signif icance of the constant on the right hand of the equality sign in Equation (1~ lies in its synchronism with the Earth's yearly motion about the Sun.
In order to preserve the orientation of the orbital plane with respect to the earth-sun line, it is nPc Pcs~ry to 10 advance the plane of the orbit by 360 degrees/365.25 days or 0 . 9 8 5 6 deg/ day .
Specifically, the effect of J2 term on n and can be ex~ressed as follows:
dtn,, =-1. SnJ2 (Rv/a) Z (cosi) (1-e2) ~'.. (2) ~-2 . 06474xlO1~a~7/2 (cosi) (i-ea) ~' dt~* 0~75nJ2(R~/a)2(4-5sin2i) (1_e2)~ (3) ~1. 03237xlO1~a~~/Z (4-5sin2i) (l-e2) ~
, where n is the mean motion in degrees per 15 day, Re is the earth's equatorial radius, a is the semi major axis in kilometers, e is the eccentricity, i is the inclination and the change in n and ~ are bo~h in degrees per day.
According to this first Pmhor7;r~nt, we want 20 to make the d~/dt term approach zero. Luckily, this can WO9S/22489 r~ l~u~ ll3 2 ~ ~ 3 5 2 2 be easily done by a.lju~; - i of the sine term in equation 3 to zero. Therefore, we set 5sin2i = 4, requiring that sin2i=4/5 or i = arc sin {square root (4/5~}; 50 i = 63.435 or its 5 complement 116 . 5 65 .
This: '- 'ir- -t preferably uses an elliptical orbit of 116 . 565 degrees. The prior art has used circular sun synchronous orbits. All so-called circular orbits may have some slight degree of ellipticity. For 10 purposes of this specification, an elliptical orbit is def ined as an orbit whose ellipticity is greater than o. 002 . This effectively excludes circular orbits which are slightly elliptical due to imperfections in the orbits. These elliptical orbits, with 15 e sY o. 001 are sometimes called frozen orbits.
Therefore, we set ~ ) ) to zero, and we dt set dt (Q) to ~0.9856, the amount per day by which the earth revolves around the sun. By substituting this into equation (3), a set of combinations of apogee, perigee 20 and inclination are found which satisfy the attached formula which are shown in the attached Figure 1.
For an elliptical sun syn~ o~ us orbit, only a very small circumscribed part of this design space can be used. First, this satellite should have no apsidal 25 rotation, to keep the apogee in one hpm;crhpre.
Accordingly, the inclination must be 116.565. A certain amount of leeway is possible, however, and practically 8rPAk;n~ the orbit can be ;nr~ ;nPd anywhere between 115 and 118 and still obtain sufficiently stable 30 characteristics, although some minor orbit corrections may be nPcPCcAry from time to time.
W0 9s~22~89 P~ .,,5~1~7~
3 ~ 2 2 Along this line, only a certain class of orbits are usable. Circular orbits are known in the prior art, and do not have the ability to produce the preferential coverage characteristics in the way done 5 according to the present invention. Therefore, a leftmost limit on the design space 6hown by point 102 in Fig. 1 rel,Les.~ b the limit to require an elliptical orbit. The rightmost limit is set by the minimum satellite height at perigee. A satellite orbit should 10 be, practically sr~k;ng, greater than, for example, 100 nautical miles. Preferably, the lowest limit is 250.
The point 104 L~L-~s~al~l~s the position where perigee will fall below 100 nautical miles. Therefore, the design space extends between the points 102 and 104. Within 15 this design space, the inclination varie6 between 115 and 119. The usable design area is therefore shown in the box in Figure 1.
Within that box, period varie6 from 2 . 6 to 3.1 hours, ~pogee varies from 100 to 4600 nautical miles, 20 and perigee varies from 100 to 2200 nautical miles.
These orbits allow the coverage to be adjusted, or biased, to favor the Northern h~mi ~rh~-re over the Southern h~ rh~re.
More specifically, the allowable range of 25 orbital parameters includes orbital periods between 2. 68 and 3.1 hours, and orbital eccentricities between 0.002 and o . 3 8 .
The postulated orbit preferably has an orbit or orbits with the integral period value of 3 . 0 hours.
30 This 3-hour orbit with cuLLF~ n~ mean motion of an even 8 revolutions per day will result in a repeating ground track. The use of other, non-integral values for orbital period(s) still results in the satellite's ground track crossing the Equator on the ~F:r~n~;n~ and 35 t~ r.~ntl;n~ nodes at given values o~ local time, but the WO 95l22489 P~~ 2113 2 ~ 33522 points of such crossings will not now occur at fixed longitudinal points. Any point along the design space horizontal line (116.565 degrees) may be selected to provide a base line set of orbital parameters upon which 5 such an orbit or constellation may be configured.
A~pl ications This invention may be used for - irations~ earth sensing, survp;ll;,nre, weather, or any other satellite function found useful for satisfying 10 mission requirements. The invention can be used in a single satellite mode, and will provide better coverage during daylight hours than during nighttime hours.
Effectively, a~v~ is "stolen" from nighttime coverage ~md diverted to daytime coverage. The most probable 15 future application of the invention in this case will be found in the construction and use of ordered arrays (or constellations) of such satellites.
In order to show how this system would be used, a few examples from the design space in Figure 1 20will be diF:rllc~led herein. These examples are analyzed using a computer program such as Orbital Workbench, or OSAC written by the Naval Research Lab, or Graftrak, available from Silicon Solutions, Inc; Houston Texas.
This program is run with the inclination, apogee and 25 other information from the chart in Figure l. The characteristics of that orbit are obtained. Then, the desired characteristics are used to modify the orbit until the proper places from the design space are identified. Some preferred orbits according to the 30 present invention will be described herein.
The second embodiment of the present invention is one which produces a special kind of elliptical orbit. This special orbit has a constant-pointing apogee, which faces in a constant direction Wo 95/22489 PCr/USs5l02113 2 t ~ 3 5 2 2 relative to the sun all year round. This iB obtained by a posigrade orbit in which the equation dd~t ~ ddQt = 9 8 - - - - - .. ... ( 1 ) is satisf ied.
Figure 2 shows a resulting sun ~y~ r~ ous 5 orbit with apogee pointing towards the sun. This preferred ~lnho~ nt of the present invention comprises a satellite in an orbit which has a sun ~yl~ ,us apogee which assumes an orbit around the earth such that the apogee of the satellite is always facing towards the sun.
10 The satellite 100 i8 shown with its orbit 102, orbiting the eartll 104. Different seasons find the earth at different portions around the sun, and these portions are shown as positions 110, 120, 130 and 140. The apogee point, sllown as element 142, is always facing the sun.
15 To obtain the preferred operating range for this equation, the equations d Q,, =-1. 5n'J2 ~RE/a) 2 (cosi) (l -e2) - .. (2) ~-2.06474xl0l4a7/2(cosi) (i-e2)~' dt ~J. ~75nJ2 (RF/a) 2 (4-5sin2i) ~'l _e2) -' ~3 ~1 . 03237xl ol~a7~2 ~4-58in2i) (1 -e2) ~' are '-in~cl with equation ~1) to plot the characteristics shown in Figure 3. Figure 3 shows the apogee altitude, perigee altitude, and inclination 20 forming the design space. As in the first ~ nt, only certain parts of this design space can be used. For W0 9~/22489 2 1 ~3522 example, the practical limit on the altitude of a satellite is greater than 100 nautical miles. All other parts of this design space can be used.
These orbits have characteristics which are 5 synchronous with respect to the time of year. By specifying any initial RAhN and epoch, therefore, the Right Ascension of the apogee of this orbit will stay constant over time with respect to the sun. For one special class of orbits, the apogee will always be 10 pointing towards the sun as shown in Figure 3. For another special class of orbits, the apogee will be pointing for eYample at 2 degrees relative to the sun.
In any of these orbits, theref ore, the apogee is controlled to be ~ D-arlL.
15 For this ' ~'i --'1~, the apogee is always at a constant right ascension angle from the right a6cension of the earth-sun line: usable inclinations range from 0 to 43 degrees, usable periods from 1.7 to 5.0 hours (again, preferab~y 3 hours to obtain a repeating ground 20 track), and usable eccentricities from 0 . 0002 to 0 . 56 .
A few examples of how these orbits would be chosen and the characteristics thereof are explained herein .
According to another pref erred mode of the 25 invention, the first and /or second embodiments are further modified to include multiple satellite configurations. This modification comprises a constellation of satellites which preferentially cover the Northern h~mi srhPre~ as compared with the Southern 30 h-~mi crh~re or vice versa.
The constellation of satellites orbiting the earth 400 is shown in Figure 4. Of course, it should be under6tood that while Figure 4 shows only two satellites, 402 and 404, in reality there would be many more. These 35 two satellites are located and operate to preferentially Wo 95/22489 P~ O2II3 'S -18- ~183522 cover one portion of the earth over another (fir6t ~ :';r- lt) and/or one time of day (second Prhs~lir--lt) over another.
Each of the satellites communicates with a 5 earth-based earth station, shown schematically as station 406, in a conventional way to exchange information therewitll. Accordingly, the present invention also contemplates use of an earth station with such satellites, this earth station having characteristics to 10 track satellites having the characteristics ~ cllcced above, and to communicate therewith. There are a plurality of earth stations, each positioned on the earth, and each including tracking eql~; L to track a motion of at least one of said satellites. Each earth 15 station, and each satellite also includes communication equipment to communicate between the earth station and the at least one satellite.
The satellites according to the present invention are initially boosted into their orbits by 20 special rockets of the type intended to deliver satellites. One such rocket, 500, with the satellite 502 therein is shown in Figure 5. The rocket include6 a first stage engine 504, of any known solid or liquid fuel type, and a second stage engine 506. Rocket engines are 25 well known in the art, and it will be assumed that the second stage engine is a liquid type rocket fuel engine.
This engine combines a liquid fuel with an oxygenator at point 508, which ignites the fuel. The ignition accelerates the speed of the fuel through a constriction 30 510, causing a sonic shock wave shown as 512 which travels out the output nozzle 514. (It must be understood that the f ixture in Figure 5 shows this stage rocket with the first stage still attached. ) WO 95/22489 l ~ ,.,' C~113 2 ~ 8 3 5 2 2 The rocket is controllable both in direction and in thrust. ~ore generally, the vector control of the rocket is controllable.
The rocket is controlled by an onboard 5 navigation computer 516. The basic characteristics of a booster rocket and guidance unit are shown, for example, in U.S. Patent 4,964,340, the disclosure of which is herewith incoLyulated by reference.
According to a fourth embodiment of the 10 rocket of the present invention, the inertial guidance unit is controlled to boost the rocket into an elliptical ~LoyLade orbit selected from the design space box around line 100 shown in Figure 1. The satellite is then delivered into that orbit, to maintain that or~it.
15 According to a fifth ~ L of the present invention, the rocket of Figure 5 has an internal guidance unit which is ~I c,yL - ' to boost the rocket into a posigrade orbit of an elliptical type, 5PlectP~l from the design space shown in Figure 3. At that time, 20 the satellite is released into the orbit, to thereby maintain thereafter the appropriate orbit.
The third, fourth and fifth Pmhorlir-nts are usable in combination with either of the f irst or second ts described above.
25 Some examples of the preferred orbits used according to the present invention will now be described.
First ~referred orbit confiquration The first preferred orbit is a four satellite minimum array ring which covers any northern hPm; RrhP~e 30 region north of 20 north latitude during daylight hours, with a minimum 15 degree elevation angle a. The satellites have an optimized afternoon ~RcPn~lin~ node, a three hour period and an argument of perigee ~ other ~han 270. The ellipse actually therefore tilts towards the WO 95/22489 1 ~ ,95, 1 ~ ~
t ~ 2 ~ 8 ~ 5 2 2 `,, .. ~! ' sun and provides a ring of orbits which are both sun S.y~ lL~ US and always have their apogee pointing towards the sun.
The characteristics of these orbits are such 5 that the satellites appear to be moving backwards from west to east since they are in ~ LoyL~.de orbit.
Using the basic satellites ~ cc~ above, sPl~c~in~ of the main orbital parameters were adjusted through trial iterations be~inn;n~ around the beginning 10 values oE a = 270 and RAAN = F (YY, MM, YY, HH, M~, and SS). The resulting graph track view 6how visibility circles and lines which reach down to a certain latitude.
This system is very unique, since with only four LEO-MEO satellites, all regions north of 20 15 latitude can be covered with visibility angles of 15.
It would take three to four times as many circular satellites to do the same thing.
Second ~Ieferred orbit confiquration The second pref erred orbit covers everything 20 in the northern h~mi~rh~re above 20 north latitude both day and night. One ring of satellites has noon ~aC-~n~i n~
nodes and the other has midnight i~c~n~l; n7 nodes. This has the significant advantage of simplifying the design of the solar array of the satellite.
25 Most satellites have solar arrays, which need to face the sun in order to power the satellite. If we u6e an orbit like the present example, then this solar array needs only one degree of freedom to follow the sun.
This simplifies the satellite design. This requirement 30 is satisfied by placing one ring with noon i~l cPn~l;n~
nodes and another ring with RAANs displaced 180 from the first ring and having midnight ~r/~nrl;n~ nodes.
Figures 6, 7A and 7B show this basic orbit.
Figure 6 shows the noon orbit, and the four satellites ~ WO gSI22489 r~
therein, l~hpllpd 01, 02, 03, and 04. Figure 7A shows the midnight ring, with the satellitefi labelled 05, 06, 07, and 08. Figure 7B shows the noon plus midnight rings . The : `- i nPcl view of Figure 7B shows that most of 5 the coverage is in the northern hPmi ~rhPre. There is only spotty coverage in the southern hPmi crhpre~ but the clustering is in the north.
rrhird Dreferred orbit confiquration A third example is a six satellite 10 equatorial, prograde, apogee pointing towards the sun orbit. This third example uses terms of the formula for advance of the line of nodes at 0.9856 per day and provides an extra degree of rPlll-n~n~ y and higher elevation angles in the tropical and equatorial zones.
15 F~ rth Preferred orbit confiquration The fourth example is another equatorial prograde orbit with apogee pointing towards the sun with only four satellites. This array emphasizes continuous equatorial region daytime C~V~LC~Y~:: with visibility angle 20 of 10. Figure 8A shows 1100 GMT which is daylight over Europe, and shows that most of Europe is well covered.
However, Europe is less well covered at 2300 GMT shown in Figure 8B.
F;fth ~referred orbit confiq~ration 25 The fifth preferred orbit constrains the visibility angle to 0 and obtains continuous equatorial region daytime coverage with only three satellites.
Agaim, there are gaps at nighttime, but none in the daylight hours. Figures 9A, 9B, and gC show the various 30 daylight hour coverages. Figure 9A shows coverage at 1535 Glrr, Figure 9B shows coverage at 740 GMT, and Figure 9C shows coverage at 1500 GMT.
WO 95~22-189 PCTIUS95/OZ113 'Q.............. ~ 2183522 Sixth Preferred orbit conf iquration Finally, the sixth preferred mode i5 shown in Figures lOA and lOB. This four satellite array combines classic sun synchronism condition f dt = 9856 with 5 the apogee on the sunward side of the earth. ~ = 262.
Apogee always occurs close to the meridian of the Earth at local apparent noon. This four satellite array provides continuous coverage of day lit areas north of 20O north latitude all year round in all countries. This 10 sixth example has an afternoon ~RcPn~lin~ node, apogee at noon, a forced inclin-~l plane, and a three hour period with the apogee equals about 4000 nautical miles.
Although only a few ~mho~;- L6 have been described in detail above, those having ordinary skill in 15 the art will certainly understand that many modifications are possible in the preferred: ~:';r ~ without departing from the teachings thereof.
All such modif ications are intended to be en ~Pcl within the following claims.
(. ! .
-- 4 -- ~ =
The ~ILyl -- of perigee ~ is an angle measured in the plane of the orbit between the point of the ~cPn~1in~ node and the nearest point of perigee.
Most practical satellites prior to the 5 invention by the present inventors used relatively simple systems based on circular orbits. The earth was covered sy_metrically by multiple satellites, which each operate to cover a section of the earth.
Elliptical orbits have been typically avoided 10 in the art, because of their asymmetries, and the c~-nCPTlPnt problems that they might cause. However, some individual elliptical orbits and elliptical orbit constellations have been ~l~,~osed. The Russian Molniya orbit is a posigrade orbit ~Riqnod for polar and high 15 latitude coverage. Other posigrade orbits have been described by John Draim in his U.S. Patents 4,809,935 and 4,854,527.
4, 809, 935 describes a three-satellite constellation giving continuous coverage of the entire 20 Northern hPmi~rhPre, and an extension of this constellation to include an equatorial orbit resulting in a four-satellite array giving continuous global coverage of both hPm; ~rhPres. This latter four satellite array provided somewhat higher elevation coverage in the 25 Northern hemisphere than in the Southern RPmi ~rhPre.
4, 854, 527 describes a common period four-satellite array giving continuous global ~:~Jv~:Lc~g~
with satellites at a lower altitude range than in the first patent. A rli~cllC~c;on of obtaining extra Northern 30 HPm; crl.Pre coverage through use of elliptic satellite constellations may be found in ANSER Space Systems Division Note SpSDN 84-1, "Satellite Constellation Design Techniques for Future Space Systems" dated September 1984, by John Draim, and James Cooper. Another 35 application of posigrade elliptic orbits is the ACE and Wo 95t22489 1 ~ JI~1II3 2i 83522 ACE-Prime orbits developed by Mr. A. Turner of Loral Corporation .
The present invention also simplif ies the design of the solar panels by requiring no more than 1 or 5 2 degrees of freedom. In the example orbit discussed herein which is 116 r ~LLvylade, the panels need only one degree of freedom. In a similar way, a satellite usually needs to radiate its heat toward cold, empty space. In the present invention, it is much easier to face the 10 satellite in a way that always faces the heat radiators away f rom the sun .
It is also well known that the earth is not totally circular, but actually it is rather oblate. That i8, the earth is bigger at the bottom than it is at the 15 top. The J2 harmonic, due to the earth's oblateness, causes the node n and argument of perigee ~ o~ an orbit to change. The gravitational pull of the earth's equatorial bulge causes, for example, the orbital plane of an eastbound satellite to swing westward. More 20 generally, the force component is directed towards the Equator. This resultant acceleration causes any satellite to reach the Equator (node) short of the crossing point where it would have reached it on a spherical earth. For each revolution, therefore, the 25 orbit regresses a ~ amount. These effects have been the subjects of various attempts at ~ ~tion.
Sun .-y,.~llrol.v~ s circular orbits are also known. These are orbits where the rotation rate of the right ascension of the Acc~ntl;ng node is equal to and in 30 the same direction as, the right ascension rate of the mean sun.
.
ST~TARY QF TT-TT' INvT~NTIoN
The previous specification, of which this is a continuation-in-part, described the invention of non-Wo 95n2489 ~ 5~02113 uniform capacity distribution tailored by latitude andpopulation. This was done using an elliptical satellite array. The present specification adds additional information to the basic elliptical orbit. A first 5 pmhoA;~~nt of the present invention is tailored by latitude and population . A second pmhoA i- ~ is also tailored by time of day. The present invention describes a specif ic way of carrying out these options by using a I~LLo~L4de intl;nPA orbit, and/or a sun ~y~ )us 10 apogee, elliptical orbit, which is equatorial, and also describes many ways of PYpAnAin~ the design space.
The parent specification achieved the ~:uy~ ,Iphical discrimination . The present specif ication v~c s on this basic technique by using the 15 pcLLuLl,ation to achieve various effects. The present inventors started with a r:~nrln;c~l form for the equation of motion of satellites in orbit. Two cases were di~cuvt:l ~d by the present inventors: 1 ) one parameter fixed at 116, to form a ~LL.,yLc,de sun 6ynchronous 20 orbit, and 2) both parameters varied to form a sun synchronous apogee orbit where the apogee always points in a constant direction relative to the sun. This resulted ir. an expansion of the possible space which allows time of day tailoring of ~OV~L~Y~. Such was 25 completely unheard of before the present invention, except for certain conditions 8uch as circular inrl ;nPA
orbit, or the ~ce equatorial orbit. The present invention, in contrast, allows an entire domain of applicability of these two conditions in order to obtain 30 various unheard of combinations. According to the present invention, two PmhoAir- Ls are described: A
first of which includes incl ;neA sun synchronous orbit and a second which includes constant pointing apogee orbits. Both of these are based on the effects on orbits 35 from the J2 term.
~ Wo gsl22489 r~
21 ~3522 ~
We have found a way to exploit the effects of earth's oblateness, with an elliptical orbit, such that for certain combinations of orbital parameters, the secular perturbations due to the earth ' S J2 gravitational 5 term may be used to advantage. Specifically, we noticed that the effects on the ~ and n terms could be used to cate the orbit of the satellite to obtain certain controllable effects. These controllable effects are obtained taking into account that the earth revolves lOrelative to the sun by 0.9856, of orbit around the sun, per day . The present invention uses p~:L LUL l,ations arising from the J2 terms, to precess the orbits in time with respect by following the general equation d~ + ddn =o . 9 856 . .. ( 1 ) to obtain elliptical orbits which have controllable and 15 constant characteristics, and which a-~y LLically cover one parameter of co~ r ~.ye preferentially over another in a way which is constant relative to the sun, all year round. According to a first ~mho~i--rt of the invention, that parameter of coverage is geographical location, and 20 according to another ~-- ir L of the invention, that parameter of coverage is time of day.
Kepler's law of motion governs any orbit. In any orbit, including an elliptical orbit, areas swept out in equal times must be equal. Hence, a satellite in an 25 elliptical orbit spends more time towards apogee than it does towards perigee. We have exploited this effect, to bias the effective orbits to increase coverage at certain geographical locations and/or times. According to the first aspect of the present invention, a first operation 30 is carried out to set the change in ~, d/dt(~) to approach 0. The change in n is then set to 0 . 9856 /day.
C '-in;n~ this with a r~LL~/yL~de orbit produces a set o~
Wo 9~l22489 A _ I / ~,J ", _`IA 1 ~ ~
2 ~ 8 3 5 2 2 specific characteri6tics referred to herein as a design space. This design space includes L~ yLCLde orbits which f avor one geographical location on the earth over another .
5 One recognition of the present invention is that the Northern T~m; srhore includes a majority of the world's earth masses and population. By covering the Northern T~~; crhore preferentially over the Southern T~pm; crhore, the coverage can be equalized as a function 10 of population .
A second ~~'~';- ~ of the present invention uses prograde orbits, using the more general equation l.
In the second omho~ t, d/dt(~) does not approach 0.
Instead, the parameters d/dt(~) and d/dt(n) are both 15 variable, but only in a way that meets the general equation dd~t ~ ddQ = 9856 .. (1 ) We have found that this produces an orbit where the apogee always points in a specif ic direction relative to the sun. This can be used to increase the satellite 20 daytime C~VC:L~ effect, or the effect from 9AM to 5PM
for example. Of course, more people require satellite services during business hours than at any other time.
Hence, there is more of a demand for satellite servicc during the day than there is at night.
25 After identifying the advantages possible from an elliptical orbit in this way, we have identif ied a technique of choosing parameters of orbits such that different areas will always be preferentially covered during the day.
30 The present invention teaches construction of a satellite orbit, a satellite system, a method of operating a satellite system, and a method and apparatus ~ W0 9~/22489 r~ 3 ;, . . 2 ~ 83522 g ~
of deploying a satellite into a prescribed orbit, all using orbital parameter combinations with integral or non integral mean motions specified within the design space that covers a specified set of earth (or planet) coverage 5 requirements in a more optimal manner than is obtained through more conventional orbits.
The objects of the present invention include:
Col~,LLu~;~ion of a satellite orbit using ; nrl; n~cl orbital parameter combinations within a design 10 space that extends the latitude ranges in order to meet earth (or planet) coverage requirements in a more optimal manner than are obtainable through conventional circular DUI~ .~y~ lLVll~US type orbits.
Cv...,~.ucLion of a satellite with ~IL~ ~ of 15 perigee value other than 90 or 270 degrees, such that the apogee locations may be preferentially oriented in any desired direction, preferably towards the earth-sun line, giving more extensive ( in both time and earth central angle) coverage, and such that i .,~,ed coverage during 20 daylight hGurs is achieved, than during nighttime hours, for locations at all longitudes from -180 to +180 degrees (or 180W to 180E).
Provision of the required satellite elevation angles within specif ied latitude ranges, with appropriate 25 day-night biases, for the r~ ,yLc.de elliptic orbit def ined .
Provision of a satellite orbit that maintains its integrity year-in year-out through precise orbital inj ection control so that coverage characteristics are 30 maintained ~hllJU~ uL the satellite constellation lifetime. Note: minor orbital adjustments may be required to account for smaller perturbations, e.g. third order or higher and/or solar peLLuLl,ations, which are experienced by the satellite.
The novel features of this aspect include:
WO 95122489 1 . ~ 13 ~ 1 8 3 522 greater satellite Earth coverage can be provided during the daylight hours (or bl~Ci nPc5 day, when there is heavy utilization of tel~ i cations or other useful servicec), 5 116.565 degree orbit plane inclinations, as described according to the f irst pref erred Pmhofl i - L
will provide continuous coverage of the high latitude and polar regions with elliptic orbits, not obtainable from equatorial plane orbits.
10 Relatively low orbits, which can be obtained using corrP~pon~;nq smaller rocket boosters.
~TF:F DESCRIPTION OF THE DRAWINGS
These and other aspects of the invention will now be described in detail with reference to the ac -nying 15 drawings, wherein:
Fig . 1 shows a f irst des ign space f or elliptical sun synchronous It:l.LoyLade orbits according to a f irst Qmhofl; L 0~ the present invention;
Fig. 2 shows the characteristics of a special 20 orbit according to a second pmhofl;r L of the present invention in which the apogee is always pointing towards the sun;
Fig. 3 shows a design space for this second pmhofl;-- L of the present invention using prograde 25 orbits;
Fig. 4 shows a constellation of satellites, each orbiting and - ; cating with earth stations on the earth;
Fig. 5 shows a rocket and inertial guidance 30 unit used according to the present invention to propel the rocket into orbit; and W0 9s/22489 p~ 3 ` ~ ~ ` ;i; 2 ~ ~3522 .. ~
Figs 6, 7A, 7B, 8A, 8B, 9A, 9B, 9C, lOA and lOB show characteristics of preferred orbits of the present invention.
l~E~ V~lON OF ~ l) EMF~ODJM~NT
5 The present invention exploits the gravitational effects from the earth's oblateness, in combination with a preferably elliptical orbit, to allow preferential coverage of different parts of the earth as a function of parameters which are related to satellite 10 demand. This has significant advantages since it allows preferential coverage based on a chosen characteristic, here either one h~m;~rhPre over the other, or time of day .
For instance, a satellite system primarily 15 intended for use over the United States would prefer to preferentially cover the Northern hPm; ~rhPre as opposed to the Southern h~mi~rhPre. More specifically, by choosing elliptical orbits such that anything above 40 south latitude was covered, a great majority of the 20 world's land mass could be covered without wasted capacity .
This ~ l of the invention optimizes the characteristics of the elliptical satellite to have desired ~ L C~ characteristics . According to this 25 first preferred mode, structure is described for putting a satellite in a special orbit which preferentially covers part of the earth over the other part.
The f irst type of orbits, discussed according to the present invention herein, are elliptical 30 L~lL~JyLnde orbits which provide preferential coverage of one part of the earth over ~he other part through adjustment of orbital parameters.
As mentioned above, all orbits are effected by the earth's J2 gravitational term. Thi~ term effects wo 95l22489 f~r ~r~ 21 83522 the n and ~ terms of every orbit. In order to compensate the orbit, the general equation dd~lt'+ dt=0'98---- --------. . . (1) must be satisfied. This first: `~o~; L takes a special case of the equation (1).
5 The signif icance of the constant on the right hand of the equality sign in Equation (1~ lies in its synchronism with the Earth's yearly motion about the Sun.
In order to preserve the orientation of the orbital plane with respect to the earth-sun line, it is nPc Pcs~ry to 10 advance the plane of the orbit by 360 degrees/365.25 days or 0 . 9 8 5 6 deg/ day .
Specifically, the effect of J2 term on n and can be ex~ressed as follows:
dtn,, =-1. SnJ2 (Rv/a) Z (cosi) (1-e2) ~'.. (2) ~-2 . 06474xlO1~a~7/2 (cosi) (i-ea) ~' dt~* 0~75nJ2(R~/a)2(4-5sin2i) (1_e2)~ (3) ~1. 03237xlO1~a~~/Z (4-5sin2i) (l-e2) ~
, where n is the mean motion in degrees per 15 day, Re is the earth's equatorial radius, a is the semi major axis in kilometers, e is the eccentricity, i is the inclination and the change in n and ~ are bo~h in degrees per day.
According to this first Pmhor7;r~nt, we want 20 to make the d~/dt term approach zero. Luckily, this can WO9S/22489 r~ l~u~ ll3 2 ~ ~ 3 5 2 2 be easily done by a.lju~; - i of the sine term in equation 3 to zero. Therefore, we set 5sin2i = 4, requiring that sin2i=4/5 or i = arc sin {square root (4/5~}; 50 i = 63.435 or its 5 complement 116 . 5 65 .
This: '- 'ir- -t preferably uses an elliptical orbit of 116 . 565 degrees. The prior art has used circular sun synchronous orbits. All so-called circular orbits may have some slight degree of ellipticity. For 10 purposes of this specification, an elliptical orbit is def ined as an orbit whose ellipticity is greater than o. 002 . This effectively excludes circular orbits which are slightly elliptical due to imperfections in the orbits. These elliptical orbits, with 15 e sY o. 001 are sometimes called frozen orbits.
Therefore, we set ~ ) ) to zero, and we dt set dt (Q) to ~0.9856, the amount per day by which the earth revolves around the sun. By substituting this into equation (3), a set of combinations of apogee, perigee 20 and inclination are found which satisfy the attached formula which are shown in the attached Figure 1.
For an elliptical sun syn~ o~ us orbit, only a very small circumscribed part of this design space can be used. First, this satellite should have no apsidal 25 rotation, to keep the apogee in one hpm;crhpre.
Accordingly, the inclination must be 116.565. A certain amount of leeway is possible, however, and practically 8rPAk;n~ the orbit can be ;nr~ ;nPd anywhere between 115 and 118 and still obtain sufficiently stable 30 characteristics, although some minor orbit corrections may be nPcPCcAry from time to time.
W0 9s~22~89 P~ .,,5~1~7~
3 ~ 2 2 Along this line, only a certain class of orbits are usable. Circular orbits are known in the prior art, and do not have the ability to produce the preferential coverage characteristics in the way done 5 according to the present invention. Therefore, a leftmost limit on the design space 6hown by point 102 in Fig. 1 rel,Les.~ b the limit to require an elliptical orbit. The rightmost limit is set by the minimum satellite height at perigee. A satellite orbit should 10 be, practically sr~k;ng, greater than, for example, 100 nautical miles. Preferably, the lowest limit is 250.
The point 104 L~L-~s~al~l~s the position where perigee will fall below 100 nautical miles. Therefore, the design space extends between the points 102 and 104. Within 15 this design space, the inclination varie6 between 115 and 119. The usable design area is therefore shown in the box in Figure 1.
Within that box, period varie6 from 2 . 6 to 3.1 hours, ~pogee varies from 100 to 4600 nautical miles, 20 and perigee varies from 100 to 2200 nautical miles.
These orbits allow the coverage to be adjusted, or biased, to favor the Northern h~mi ~rh~-re over the Southern h~ rh~re.
More specifically, the allowable range of 25 orbital parameters includes orbital periods between 2. 68 and 3.1 hours, and orbital eccentricities between 0.002 and o . 3 8 .
The postulated orbit preferably has an orbit or orbits with the integral period value of 3 . 0 hours.
30 This 3-hour orbit with cuLLF~ n~ mean motion of an even 8 revolutions per day will result in a repeating ground track. The use of other, non-integral values for orbital period(s) still results in the satellite's ground track crossing the Equator on the ~F:r~n~;n~ and 35 t~ r.~ntl;n~ nodes at given values o~ local time, but the WO 95l22489 P~~ 2113 2 ~ 33522 points of such crossings will not now occur at fixed longitudinal points. Any point along the design space horizontal line (116.565 degrees) may be selected to provide a base line set of orbital parameters upon which 5 such an orbit or constellation may be configured.
A~pl ications This invention may be used for - irations~ earth sensing, survp;ll;,nre, weather, or any other satellite function found useful for satisfying 10 mission requirements. The invention can be used in a single satellite mode, and will provide better coverage during daylight hours than during nighttime hours.
Effectively, a~v~ is "stolen" from nighttime coverage ~md diverted to daytime coverage. The most probable 15 future application of the invention in this case will be found in the construction and use of ordered arrays (or constellations) of such satellites.
In order to show how this system would be used, a few examples from the design space in Figure 1 20will be diF:rllc~led herein. These examples are analyzed using a computer program such as Orbital Workbench, or OSAC written by the Naval Research Lab, or Graftrak, available from Silicon Solutions, Inc; Houston Texas.
This program is run with the inclination, apogee and 25 other information from the chart in Figure l. The characteristics of that orbit are obtained. Then, the desired characteristics are used to modify the orbit until the proper places from the design space are identified. Some preferred orbits according to the 30 present invention will be described herein.
The second embodiment of the present invention is one which produces a special kind of elliptical orbit. This special orbit has a constant-pointing apogee, which faces in a constant direction Wo 95/22489 PCr/USs5l02113 2 t ~ 3 5 2 2 relative to the sun all year round. This iB obtained by a posigrade orbit in which the equation dd~t ~ ddQt = 9 8 - - - - - .. ... ( 1 ) is satisf ied.
Figure 2 shows a resulting sun ~y~ r~ ous 5 orbit with apogee pointing towards the sun. This preferred ~lnho~ nt of the present invention comprises a satellite in an orbit which has a sun ~yl~ ,us apogee which assumes an orbit around the earth such that the apogee of the satellite is always facing towards the sun.
10 The satellite 100 i8 shown with its orbit 102, orbiting the eartll 104. Different seasons find the earth at different portions around the sun, and these portions are shown as positions 110, 120, 130 and 140. The apogee point, sllown as element 142, is always facing the sun.
15 To obtain the preferred operating range for this equation, the equations d Q,, =-1. 5n'J2 ~RE/a) 2 (cosi) (l -e2) - .. (2) ~-2.06474xl0l4a7/2(cosi) (i-e2)~' dt ~J. ~75nJ2 (RF/a) 2 (4-5sin2i) ~'l _e2) -' ~3 ~1 . 03237xl ol~a7~2 ~4-58in2i) (1 -e2) ~' are '-in~cl with equation ~1) to plot the characteristics shown in Figure 3. Figure 3 shows the apogee altitude, perigee altitude, and inclination 20 forming the design space. As in the first ~ nt, only certain parts of this design space can be used. For W0 9~/22489 2 1 ~3522 example, the practical limit on the altitude of a satellite is greater than 100 nautical miles. All other parts of this design space can be used.
These orbits have characteristics which are 5 synchronous with respect to the time of year. By specifying any initial RAhN and epoch, therefore, the Right Ascension of the apogee of this orbit will stay constant over time with respect to the sun. For one special class of orbits, the apogee will always be 10 pointing towards the sun as shown in Figure 3. For another special class of orbits, the apogee will be pointing for eYample at 2 degrees relative to the sun.
In any of these orbits, theref ore, the apogee is controlled to be ~ D-arlL.
15 For this ' ~'i --'1~, the apogee is always at a constant right ascension angle from the right a6cension of the earth-sun line: usable inclinations range from 0 to 43 degrees, usable periods from 1.7 to 5.0 hours (again, preferab~y 3 hours to obtain a repeating ground 20 track), and usable eccentricities from 0 . 0002 to 0 . 56 .
A few examples of how these orbits would be chosen and the characteristics thereof are explained herein .
According to another pref erred mode of the 25 invention, the first and /or second embodiments are further modified to include multiple satellite configurations. This modification comprises a constellation of satellites which preferentially cover the Northern h~mi srhPre~ as compared with the Southern 30 h-~mi crh~re or vice versa.
The constellation of satellites orbiting the earth 400 is shown in Figure 4. Of course, it should be under6tood that while Figure 4 shows only two satellites, 402 and 404, in reality there would be many more. These 35 two satellites are located and operate to preferentially Wo 95/22489 P~ O2II3 'S -18- ~183522 cover one portion of the earth over another (fir6t ~ :';r- lt) and/or one time of day (second Prhs~lir--lt) over another.
Each of the satellites communicates with a 5 earth-based earth station, shown schematically as station 406, in a conventional way to exchange information therewitll. Accordingly, the present invention also contemplates use of an earth station with such satellites, this earth station having characteristics to 10 track satellites having the characteristics ~ cllcced above, and to communicate therewith. There are a plurality of earth stations, each positioned on the earth, and each including tracking eql~; L to track a motion of at least one of said satellites. Each earth 15 station, and each satellite also includes communication equipment to communicate between the earth station and the at least one satellite.
The satellites according to the present invention are initially boosted into their orbits by 20 special rockets of the type intended to deliver satellites. One such rocket, 500, with the satellite 502 therein is shown in Figure 5. The rocket include6 a first stage engine 504, of any known solid or liquid fuel type, and a second stage engine 506. Rocket engines are 25 well known in the art, and it will be assumed that the second stage engine is a liquid type rocket fuel engine.
This engine combines a liquid fuel with an oxygenator at point 508, which ignites the fuel. The ignition accelerates the speed of the fuel through a constriction 30 510, causing a sonic shock wave shown as 512 which travels out the output nozzle 514. (It must be understood that the f ixture in Figure 5 shows this stage rocket with the first stage still attached. ) WO 95/22489 l ~ ,.,' C~113 2 ~ 8 3 5 2 2 The rocket is controllable both in direction and in thrust. ~ore generally, the vector control of the rocket is controllable.
The rocket is controlled by an onboard 5 navigation computer 516. The basic characteristics of a booster rocket and guidance unit are shown, for example, in U.S. Patent 4,964,340, the disclosure of which is herewith incoLyulated by reference.
According to a fourth embodiment of the 10 rocket of the present invention, the inertial guidance unit is controlled to boost the rocket into an elliptical ~LoyLade orbit selected from the design space box around line 100 shown in Figure 1. The satellite is then delivered into that orbit, to maintain that or~it.
15 According to a fifth ~ L of the present invention, the rocket of Figure 5 has an internal guidance unit which is ~I c,yL - ' to boost the rocket into a posigrade orbit of an elliptical type, 5PlectP~l from the design space shown in Figure 3. At that time, 20 the satellite is released into the orbit, to thereby maintain thereafter the appropriate orbit.
The third, fourth and fifth Pmhorlir-nts are usable in combination with either of the f irst or second ts described above.
25 Some examples of the preferred orbits used according to the present invention will now be described.
First ~referred orbit confiquration The first preferred orbit is a four satellite minimum array ring which covers any northern hPm; RrhP~e 30 region north of 20 north latitude during daylight hours, with a minimum 15 degree elevation angle a. The satellites have an optimized afternoon ~RcPn~lin~ node, a three hour period and an argument of perigee ~ other ~han 270. The ellipse actually therefore tilts towards the WO 95/22489 1 ~ ,95, 1 ~ ~
t ~ 2 ~ 8 ~ 5 2 2 `,, .. ~! ' sun and provides a ring of orbits which are both sun S.y~ lL~ US and always have their apogee pointing towards the sun.
The characteristics of these orbits are such 5 that the satellites appear to be moving backwards from west to east since they are in ~ LoyL~.de orbit.
Using the basic satellites ~ cc~ above, sPl~c~in~ of the main orbital parameters were adjusted through trial iterations be~inn;n~ around the beginning 10 values oE a = 270 and RAAN = F (YY, MM, YY, HH, M~, and SS). The resulting graph track view 6how visibility circles and lines which reach down to a certain latitude.
This system is very unique, since with only four LEO-MEO satellites, all regions north of 20 15 latitude can be covered with visibility angles of 15.
It would take three to four times as many circular satellites to do the same thing.
Second ~Ieferred orbit confiquration The second pref erred orbit covers everything 20 in the northern h~mi~rh~re above 20 north latitude both day and night. One ring of satellites has noon ~aC-~n~i n~
nodes and the other has midnight i~c~n~l; n7 nodes. This has the significant advantage of simplifying the design of the solar array of the satellite.
25 Most satellites have solar arrays, which need to face the sun in order to power the satellite. If we u6e an orbit like the present example, then this solar array needs only one degree of freedom to follow the sun.
This simplifies the satellite design. This requirement 30 is satisfied by placing one ring with noon i~l cPn~l;n~
nodes and another ring with RAANs displaced 180 from the first ring and having midnight ~r/~nrl;n~ nodes.
Figures 6, 7A and 7B show this basic orbit.
Figure 6 shows the noon orbit, and the four satellites ~ WO gSI22489 r~
therein, l~hpllpd 01, 02, 03, and 04. Figure 7A shows the midnight ring, with the satellitefi labelled 05, 06, 07, and 08. Figure 7B shows the noon plus midnight rings . The : `- i nPcl view of Figure 7B shows that most of 5 the coverage is in the northern hPmi ~rhPre. There is only spotty coverage in the southern hPmi crhpre~ but the clustering is in the north.
rrhird Dreferred orbit confiquration A third example is a six satellite 10 equatorial, prograde, apogee pointing towards the sun orbit. This third example uses terms of the formula for advance of the line of nodes at 0.9856 per day and provides an extra degree of rPlll-n~n~ y and higher elevation angles in the tropical and equatorial zones.
15 F~ rth Preferred orbit confiquration The fourth example is another equatorial prograde orbit with apogee pointing towards the sun with only four satellites. This array emphasizes continuous equatorial region daytime C~V~LC~Y~:: with visibility angle 20 of 10. Figure 8A shows 1100 GMT which is daylight over Europe, and shows that most of Europe is well covered.
However, Europe is less well covered at 2300 GMT shown in Figure 8B.
F;fth ~referred orbit confiq~ration 25 The fifth preferred orbit constrains the visibility angle to 0 and obtains continuous equatorial region daytime coverage with only three satellites.
Agaim, there are gaps at nighttime, but none in the daylight hours. Figures 9A, 9B, and gC show the various 30 daylight hour coverages. Figure 9A shows coverage at 1535 Glrr, Figure 9B shows coverage at 740 GMT, and Figure 9C shows coverage at 1500 GMT.
WO 95~22-189 PCTIUS95/OZ113 'Q.............. ~ 2183522 Sixth Preferred orbit conf iquration Finally, the sixth preferred mode i5 shown in Figures lOA and lOB. This four satellite array combines classic sun synchronism condition f dt = 9856 with 5 the apogee on the sunward side of the earth. ~ = 262.
Apogee always occurs close to the meridian of the Earth at local apparent noon. This four satellite array provides continuous coverage of day lit areas north of 20O north latitude all year round in all countries. This 10 sixth example has an afternoon ~RcPn~lin~ node, apogee at noon, a forced inclin-~l plane, and a three hour period with the apogee equals about 4000 nautical miles.
Although only a few ~mho~;- L6 have been described in detail above, those having ordinary skill in 15 the art will certainly understand that many modifications are possible in the preferred: ~:';r ~ without departing from the teachings thereof.
All such modif ications are intended to be en ~Pcl within the following claims.
Claims (79)
1. A satellite in orbit around the earth, defining an elliptical orbit which has orbital parameters to satisfy the equation =0.9856.......... (1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said satellite in orbit asymmetrically covers one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round.
2. A satellite as in claim 1, wherein said parameter of coverage is geographical location.
3. A satellite as in claim 1, wherein said parameter of coverage is time of day.
4. A satellite as in claim 2, wherein said elliptical orbit is retrograde.
5. A satellite as in claim 3, wherein said elliptical orbit is prograde.
6. A satellite as in claim 4, wherein said orbit is chosen such that approaches zero.
7. A satellite as in claim 6, wherein an inclination of said orbit is set to substantially 116°.
8. A satellite as in claim 6, wherein an inclination of said orbit is set to between substantially 115 and 118°
9. A satellite as in claim 1, having a period of three hours.
10. A satellite as in claim 8, wherein period varies from 2.6 to 3.0 hours, apogee varies from 200 to 400 nautical miles, and perigee varies from 2200 to 400 nautical miles.
11. A satellite as in claim 8, having orbital periods between 2.68 and 3.05 hours, and orbital eccentricities between 0.002 and 0.371.
12. A satellite as in claim 5, wherein the apogee is always at a constant angle from the earth-sun line: an inclination is between 0 and 43 degrees, period is between 1.7 to 5.0 hours, and eccentricity is between 0.0002 to 0.56.
13. A satellite as in claim 5, wherein inclination is greater than 0.5°, but less than 43°.
14. A satellite as in claim 5, wherein inclination is greater than 10°, but less than 43°.
15. A satellite in orbit as in claim 1, wherein said orbit is inclined.
16. A satellite in orbit around the earth, defining an elliptical, retrograde orbit in which =
0 and = 0.9856, where .OMEGA. is the right ascension of the ascending node and .omega. is the argument of perigee, and wherein an inclination of the orbit is set in a range between 115° and 118°.
0 and = 0.9856, where .OMEGA. is the right ascension of the ascending node and .omega. is the argument of perigee, and wherein an inclination of the orbit is set in a range between 115° and 118°.
17. A satellite in orbit around the earth, defining an elliptical orbit which has orbital parameters to satisfy the equation =0.9856...........(1) wherein neither nor equals 0, where .OMEGA.
equals the right ascension of the ascending node and .omega. is the argument of perigee and wherein said orbit is prograde and said satellite in orbit asymmetrically covers one parameter of coverage, preferentially over another in a way which is constant relative to the sun, all year round.
equals the right ascension of the ascending node and .omega. is the argument of perigee and wherein said orbit is prograde and said satellite in orbit asymmetrically covers one parameter of coverage, preferentially over another in a way which is constant relative to the sun, all year round.
18. A satellite in orbit as in claim 17, wherein said orbit is inclined.
19. A satellite in orbit as in claim 18, wherein said inclination is 10° or greater.
20. A satellite in orbit as in claim 17, wherein a period of the orbit is three hours.
21. A satellite in orbit as in claim 17, wherein an apogee of said orbit is always pointing towards the sun, all year round.
22. A satellite in orbit as in claim 1, wherein an ascending node of the satellite is either at noon or at midnight.
23. A method of orbiting a satellite around the earth, comprising the steps of obtaining a satellite, and propelling said satellite into an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said orbit has characteristics to asymmetrically cover one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round.
24. A method as in claim 23, wherein said propelling step includes propelling in a retrograde direction, said one parameter includes at least geographical location, and orbit is defined such that approaches zero.
25. A method as in claim 24, wherein said propelling step includes propelling said satellite into an orbit which has an inclination substantially between 115 and 118°.
26. A method as in claim 23, wherein said propelling step includes propelling in a prograde direction, said one parameter includes at least time of day, and orbit is defined such that is not zero.
27. A method as in claim 26, wherein said propelling step includes propelling said satellite into an orbit whose inclination is greater than 0.5°, but less than 43°.
28. A method as in claim 26, wherein said propelling step includes propelling said satellite into an orbit whose inclination is greater than 10°, but less than 43°.
29. A constellation of satellites in orbit around the earth, each defining an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein each of said satellites in orbit asymmetrically covers one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round.
30. A constellation of satellites as in claim 29, wherein said parameter of coverage is geographical location.
31. A constellation of satellites as in claim 30, wherein said elliptical orbit is retrograde.
32. A constellation of satellites as in claim 31, wherein said orbit is chosen such that approaches zero.
33. A constellation of satellites as in claim 32, wherein an inclination of said orbit is set to substantially 116°.
34. A constellation of satellites as in claim 32, wherein an inclination of said orbit is set to substantially 115 and 118°
35. A constellation of satellites as in claim 29, wherein said parameter of coverage is time of day.
36. A constellation of satellites as in claim 35, wherein each said elliptical orbit is prograde.
37. A constellation of satellites as in claim 29, wherein each satellite has a period of three hours.
38. A constellation of satellites as in claim 36, wherein the apogee is always at a constant angle from the earth-sun line: an inclination is between 0 and 43 degrees, period is between 1.7 to 5.0 hours, and eccentricity is between 0.0002 to 0.56.
39. A constellation of satellites as in claim 35, wherein inclination is greater than 0.5°.
40. A constellation of satellites as in claim 35, wherein inclination is greater than 10°.
41. A constellation of satellites as in claim 36, wherein said orbit is inclined.
42. A constellation of satellites as in claim 29, wherein an ascending node of each of the satellites is either at noon or at midnight.
43. A constellation of satellites as in claim 29, wherein said constellation includes a first ring of satellites, each of which have noon ascending nodes, and a second ring of satellites, each of which have midnight ascending nodes.
44. A communication system, comprising: a constellation of satellites in orbit around the earth, each defining an elliptical orbit which has orbital parameters to satisfy the equation -.09856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein each of said satellites in orbit asymmetrically covers one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round; and a plurality of earth stations, each positioned on the earth, and each including tracking equipment to track a motion of at least one of said satellites, and communication equipment to communicate with said at least one of said satellites.
45. A system as in claim 44, wherein said parameter of coverage is geographical location.
46. A system as in claim 45, wherein said elliptical orbit is retrograde.
47. A system as in claim 46, wherein said orbit is chosen such that approaches zero.
48. A system as in claim 47, wherein an inclination of said orbit is set to substantially 116°.
49. A system as in claim 47, wherein an inclination of said orbit is set to substantially 115 and 118°
50. A system as in claim 44, wherein said parameter of coverage is time of day.
51. A system as in claim 50, wherein said elliptical orbit is prograde.
52. A system as in claim 44, having a period of three hours.
53. A system as in claim 51, wherein the apogee is always at a constant angle from the earth-sun line: an inclination is between 0 and 43 degrees, period is between 1.7 to 5.0 hours, and eccentricity is between 0.0002 to 0.56.
54. A system as in claim 51, wherein inclination is greater than 0.5°.
55. A system as in claim 51, wherein inclination is greater than 10°.
56. A system as in claim 51, wherein said orbit is i inclined.
57. A system as in claim 44, wherein an ascending node of each of the satellites is either at noon or at midnight.
58. A system as in claim 44, wherein said constellation includes a first ring of satellites, each of which have noon ascending nodes, and a second ring of satellites, each of which have midnight ascending nodes.
59. A method of communicating with a satellite, comprising the steps of :
propelling said satellite into an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said orbit has characteristics to asymmetrically cover one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round;
providing a ground station which tracks where in the orbit the satellite will be at any given time; and communicating between the ground station and the satellite.
propelling said satellite into an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said orbit has characteristics to asymmetrically cover one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round;
providing a ground station which tracks where in the orbit the satellite will be at any given time; and communicating between the ground station and the satellite.
60. A method as in claim 59, wherein said parameter of coverage is geographical location.
61. A method as in claim 59, wherein said propelling step propels into an elliptical retrograde orbit.
62. A method as in claim 61, wherein said orbit is chosen such that approaches zero.
63. A method as in claim 62, wherein said propelling step includes propelling said satellite into an orbit whose inclination of said orbit is set to substantially 116°.
64. A method as in claim 62, wherein said propelling step includes propelling said satellite into an orbit whose inclination of said orbit is set to substantially 115 and 118°
65. A method as in claim 59, wherein said parameter of coverage is time of day.
66. A method as in claim 65, wherein said elliptical orbit is prograde.
67. A method as in claim 59, having a period of three hours.
68. A method as in claim 66, wherein the apogee is always at a constant angle from the earth-sun line: an inclination is between 0 and 43 degrees, period is between 1.7 to 5.0 hours, and eccentricity is between 0.0002 to 0.56.
69. A method as in claim 66, wherein said propelling step includes propelling said satellite into an orbit whose inclination is greater than 0.5°.
70. A method as in claim 66, wherein said propelling step includes propelling said satellite into an orbit whose inclination is greater than 10°.
71. A method as in claim 59, wherein said propelling step includes propelling said satellite into an orbit whose said orbit is inclined.
72. A method as in claim 59, wherein an ascending node of each of the satellites is either at noon or at midnight.
73. A method as in claim 59, wherein there are a constellation of satellites which include a first ring of satellites, each of which have noon ascending nodes, and a second ring of satellites, each of which have midnight ascending nodes.
74. A rocket and satellite combination comprising:
a first part, including a rocket which includes boosters for boosting a satellite into orbit, and a inertial guidance unit, said inertial guidance unit including means for propelling said satellite into an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said orbit has characteristics to asymmetrically cover one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round and a satellite in said rocket, said rocket also including means for releasing said satellite into said orbit.
a first part, including a rocket which includes boosters for boosting a satellite into orbit, and a inertial guidance unit, said inertial guidance unit including means for propelling said satellite into an elliptical orbit which has orbital parameters to satisfy the equation =0.9856..........(1) , where .OMEGA. is the right ascension of the ascending node, and .omega. is the argument of perigee, and wherein said orbit has characteristics to asymmetrically cover one parameter of coverage preferentially over another in a way which is constant relative to the sun, all year round and a satellite in said rocket, said rocket also including means for releasing said satellite into said orbit.
75. A combination as in claim 74, wherein said parameter of coverage is geographical location.
76. A combination as in claim 75, wherein said elliptical orbit is retrograde.
77. A combination as in claim 76, wherein said orbit is chosen such that approaches zero.
78. A combination as in claim 74, wherein said parameter of coverage is time of day.
79. A combination as in claim 78, wherein said elliptical orbit is prograde.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US08/197,260 | 1994-02-16 | ||
| US08/197,260 US5582367A (en) | 1992-06-02 | 1994-02-16 | Elliptical orbit satellite, system, and deployment with controllable coverage characteristics |
Publications (1)
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|---|---|
| CA2183522A1 true CA2183522A1 (en) | 1995-08-24 |
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ID=22728674
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002183522A Abandoned CA2183522A1 (en) | 1994-02-16 | 1995-02-14 | Elliptical orbit satellite coverage and deployment system |
Country Status (5)
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| US (2) | US5582367A (en) |
| EP (1) | EP0760778A4 (en) |
| AU (1) | AU1879495A (en) |
| CA (1) | CA2183522A1 (en) |
| WO (1) | WO1995022489A1 (en) |
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| US3349398A (en) * | 1964-08-27 | 1967-10-24 | Itt | Satellite communication system |
| US3340531A (en) * | 1964-10-05 | 1967-09-05 | Martin Marietta Corp | Satellite communication system |
| US3384891A (en) * | 1965-02-11 | 1968-05-21 | Gen Electric | Method and system for long distance navigation and communication |
| US3497807A (en) * | 1966-08-31 | 1970-02-24 | Us Navy | Multipurpose satellite system |
| US4854527A (en) * | 1985-07-19 | 1989-08-08 | Draim John E | Tetrahedral multi-satellite continuous-coverage constellation |
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| US4872015A (en) * | 1986-12-01 | 1989-10-03 | Hughes Aircraft Company | Satellite communications system for mobile users |
| FR2628274B1 (en) * | 1988-03-02 | 1990-08-10 | Centre Nat Etd Spatiales | COMMUNICATIONS SYSTEM WITH MOBILES USING SATELLITES |
| IL91529A0 (en) * | 1988-10-28 | 1990-04-29 | Motorola Inc | Satellite cellular telephone and data communication system |
| GB8905610D0 (en) * | 1989-03-11 | 1989-04-26 | British Aerospace | Geostationary satellite system |
| US5186419A (en) * | 1990-01-30 | 1993-02-16 | Scott David R | Space transfer vehicle and integrated guidance launch system |
| EP0510789B1 (en) * | 1991-04-22 | 1997-03-12 | Trw Inc. | Cellular telephone satellite system |
| US5433726A (en) * | 1991-04-22 | 1995-07-18 | Trw Inc. | Medium-earth-altitude satellite-based cellular telecommunications system |
| US5322248A (en) * | 1992-03-26 | 1994-06-21 | General Dynamics Corporation Space Systems Division | Methods and arrangements tailoring aerodynamic forces afforded by a payload to reduce flight loads and to assist flight control for the coupled system |
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| US5669585A (en) | 1997-09-23 |
| WO1995022489A1 (en) | 1995-08-24 |
| AU1879495A (en) | 1995-09-04 |
| EP0760778A4 (en) | 1997-05-21 |
| EP0760778A1 (en) | 1997-03-12 |
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