Spacecraft Relative Orbit Description Through Orbit Element Differences

Home , Classical element, Mean anomaly, Orbit, Osculating orbit, True anomaly

14th U.S. National Congress of Theoretical and Applied Mechanics Blacksburg, VA, June 23–28, 2002

SPACECRAFT RELATIVE ORBIT GEOMETRY DESCRIPTION THROUGH ORBIT ELEMENT DIFFERENCES

Hanspeter Schaub∗ ORION International Technologies, Albuquerque, NM 87106

The relative orbit geometry of a spacecraft formation can be elegantly described in terms of a set of orbit element differences relative to a common chief orbit. For the non-perturbed orbit motion these orbit element differences remain constant if the anomaly difference is expressed in terms of mean anomalies. A general method is presented to estimate the linearized relative orbit motion for both circular and elliptic chief reference orbits. The relative orbit is described purely through relative orbit element differences, not through the classical method of using Cartesian initial conditions. Analytical solutions of the relative motion are provided in terms of the true anomaly angle. By sweeping this angle from 0 to 2π, it is trivial to obtain estimates of the along-track, out-of-plane and orbit radial dimensions which are dictated by a particular choice of orbit element differences. The main assumption being made in the linearization is that the relative orbit radius is small compared to the Earth relative orbit radius. The resulting linearized relative motion solution can be used for both formation flying control applications or to assist in selecting the orbit element differences that yield the desired relative orbit geometry.

Introduction ther, as the orbits decay and the orbit elements such as the The relative orbits within a spacecraft formation are typi- semi-major axis change, the orbit element differences which cally prescribed through sets of Cartesian coordinate initial determine the relative motion will only vary slowly. conditions in the rotating Hill reference frame.1–4 The six Using orbit element differences in the control of a space- Cartesian initial conditions are the invariant parameters of craft formation has advantages too. At any instant it is the relative orbit. However, to solve for the relative orbit possible to map the current inertial position and velocities motion that will result from such initial conditions, the dif- vectors into an equivalent set of orbit elements. By differ- ferential equations of motion must be solved. Except for encing the orbit elements of the deputy satellite to those of the unperturbed circular reference (or chief) orbit special the chief satellite, we are able to compare these differences case, solving the differential equations for the resulting rel- to the prescribed orbit element differences and determine ative motion analytically is a very challenging task that has any relative orbit errors. Note that no differential equations been tackled in various papers. Melton develops in Refer- have to be solved to determine any relative orbit tracking ence 5 a state transition matrix that can be used to predict errors. Several relative orbit control strategies have recently the relative motion for chief orbits with small eccentricities. been suggested that feed back orbit errors in terms of orbit Tschauner and Hempel have solved the relative equations of element differences.13–17 On the other hand, using Cartesian motion directly for the general case of having an elliptic chief coordinates to describe the desired relative orbit requires orbit.6 However, their solution is not explicit and requires solving the nonlinear differential relative equations of mo- the computation of an integral. Kechichian develops in Ref- tion of the desired relative orbit and differencing these states erence 1 the analytical solution to the relative orbit motion with the current position and velocity states, unless the con- under the influence of both the J2 and J3 zonal harmonics trol tolerances allow any of the linearized relative motion assuming that the eccentricity is a very small parameter. solutions to be used. Using the orbit element difference de- Unfortunately these methods yield relatively complex solu- scription we are not required to perform any linearization or tions and the six Cartesian relative motion initial conditions solve any nonlinear differential equations to find the current do not easily reveal the nature of the resulting relative orbit. desired deputy satellite position. It is not intuitive to the relative orbit design process how to In References 16 and 18, a linearized mapping is presented adjust the initial Cartesian conditions to obtain a relative between a particular set of orbit element differences and orbit of the desired shape and size. More recently, Broucke the Cartesian position and velocity coordinates in the ro- has presented in Reference 7 an analytical solution to the tating Hill reference frame. This work was then expanded linearized relative equations of motion for eccentric chief or- in Reference 19 to provide the state transition matrix for bits. His solution uses both time and true anomaly and finds the relative orbit motion using orbit elements. This pa- the current Cartesian coordinates of a deputy satelite given per expands on this theory and develops direct relationships the initial Cartesian coordinates. between the orbit element differences and the resulting rel- An alternate set of six invariant parameters to describe ative orbit geometry for both circular and eccentric chief the relative orbit is to use orbit element differences relative orbits. In particular, the relative orbit along-track, out-of- 8–12 to the chief orbit. In References 8 and 9 the anomaly plane and orbit radial dimensions are estimated for specific difference is prescribed in terms of the mean anomalies, sets of orbit element differences. Contrary to previous work not true or eccentric anomaly. The reason being that for in this area,11, 12 the presented results apply to both circular elliptic motions, a mean anomaly difference between two and elliptic chief orbits. But more generally, analytic solu- satellites remains constant under the assumption of classical tions are provided for the linearized relative orbit motion, Keplerian two-body motion. Prescribing the relative orbit where the linearizing assumption is that the relative orbit geometry through sets of relative orbit element differences radius is small compared to the inertial orbit radius. While has the major advantage that these relative orbit coordi- Tschauner and Hempel provide the complex solution of the nates are constants of the non-perturbed orbit motion. Even true differential equations of motion, the linearized ana- if perturbations are considered, they typically have a sim- lytical solution here is obtained from geometric arguments ilar influence on each of the satellites in the formation if without solving any differential equations. The resulting these satellites are of equal build and type. For example, linearized analytic relative orbit solution is useful when de- atmospheric drag will cause the orbits to decay. However, signing a relative orbit that must meet scientific mission all satellites in the formation will experience nearly identical requirements. If the relative orbit must have a certain along- amounts of drag. As such, the relative motion between the track behavior, then this solution directly shows how to satellites is only minorly affected by this perturbation. Fur- adjust the relative orbit element differences to achieve the desired motion. Further, the general solution for elliptic ∗Aerospace Research Engineer, AIAA and AAS Member. chief orbits is specialized for the small eccentricity and near

1 2 SCHAUB: RELATIVE ORBIT GEOMETRY circular orbit case. With the small eccentricity case, linear Thus, given both the relative position vector ρ and the chief terms in the eccentricity e are retained, while higher order position vector r, we are able to determine the inertial mo- terms of e are dropped. The near-circular analysis drops any tion of a deputy satellite. terms containing e. The linearized relationship between the Instead of using Cartesian coordinates to describe the rel- orbit element difference description of the relative orbit and ative position of a deputy to the chief, we can also use orbit the classical Cartesian coordinate description is usefull in element differences.8, 9 Let the vector e be defined through both designing and controlling relative orbits in terms of the orbit elements orbit element differences. T e = (a, θ, i, q1, q2, Ω) (5) Relative Orbit Definitions where a is the semi-major axis, θ = ω+f is the true latitude To following nomenclature is adopted to describe the angle, i is the orbit inclination angle, Ω is the argument of satellites within a formation. The satellite about which the ascending node and q are defined as all other satellites are orbiting is referred to as the chief i satellite. The remaining satellites, referred to as the deputy q = e cos ω (6) satellites, are to fly in formation with the chief. Note that 1 it is not necessary that the chief position actually be occu- q2 = e sin ω (7) pied by a physical satellite. Sometimes this chief position is simply used as an orbiting reference point about which The parameter e is the eccentricity, ω is the argument of the deputy satellites orbit. To express how the relative perigee, and f be the true anomaly. Let the relative or- orbit geometry is seen by the chief, we introduce the Hill bit be described through the orbit element difference vector coordinate frame O.20 Its origin is at the osculating chief δe. Whereas all six elements of the relative orbit state X satellite position and its orientation is given by the vec- vector are time varying, all the orbit element differences, tor triad {oˆ , oˆ , oˆ } shown in Figures 1. The unit vector except for the true anomaly difference, are constant for a r θ h non-perturbed Keplerian orbit. This has many advantages oˆr is in the orbit radius direction, while oˆh is parallel to the orbit momentum vector in the orbit normal direction. when measuring the relative orbit error motion and applying it to a control law. In References 18 and 16 a convenient The unit vector oˆθ then completes the right-handed coordinate system. Let r be the chief orbit radius and h be the direct mapping is presented which translates between the chief angular momentum vector. Unless noted otherwise, Cartesian states X and the orbit element differences δe. any non-differenced states or orbit elements are assumed to In deriving this mapping, it is assumed that the relative be those of the chief. Differenced states are assumed to be orbit radius ρ is small in comparison to the inertial chief differences between the deputy and chief satellite. Mathe- orbit radius r. While Reference 16 shows both the forward matically, these O frame orientation vectors are expressed and backward mapping between these relative orbit coor- as dinates, only the mapping from orbit element differences to Cartesian Hill frame coordinates is used in the following r oˆr = (1a) development. The relative position vector components are r given in terms of orbit elements through: oˆθ = oˆh × oˆr (1b) r Vr r h x ≈ δa + r δθ − (2aq1 + r cos θ)δq1 oˆh = (1c) a V p h t (8a) r − (2aq + r sin θ)δq with h = r × r˙. Note that if the inertial chief orbit is p 2 2 ˆ circular, then oθ is parallel to the satellite velocity vector. y ≈ r(δθ + cos i δΩ) (8b) z ≈ r(sin θ δi − cos θ sin i δΩ) (8c)

t i 2 2 rb The parameter p = a(1−e ) = aη is the semi-latus rectum, O √ e 2 v with η = 1 − e being another convenient measure of the ti oˆh Deputy la e Satellite orbit eccentricity. The chief radial and transverse velocity R oˆθ ρ z components Vr and Vt are deﬁned as x h rc y V =r ˙ = (q sin θ − q cos θ) (9) r p 1 2 Chief Satellite oˆr h V = rθ˙ = (1 + q cos θ + q sin θ) (10) t p 1 2 Deputy The orbit radius r is deﬁned in terms of the orbit elements Inertial Orbit used in Eq. (5) as Chief Inertial Orbit a(1 − q2 − q2) aη2 Fig. 1 Illustration of a General Type of Spacecraft For- r = 1 2 = (11) mation with Out-Of-Plane Relative Motion 1 + q1 cos θ + q2 sin θ 1 + e cos f

The relative orbit position vector ρ and velocity vector Note that the ratio Vr/Vt in Eq. (8a) can be rewritten as ρ˙ of a deputy satellite relative to the chief is expressed in Cartesian O frame components as Vr q1 sin θ − q2 cos θ e sin f = = (12) Vt 1 + q1 cos θ + q2 sin θ 1 + e cos f ρ = (x, y, z)T (2) T Alternate mappings between orbit element diﬀerences and ρ˙ = (x, ˙ y,˙ z˙) (3) Cartesian relative orbit coordinates are found in Refer- The relative position and velocity vectors are compactly ences 10 and 21. written as General Elliptic Orbits ρ X = (4) Note that Eq. 8 provides us a direct linear mapping ρ˙ between orbit element diﬀerences δe and the Hill frame SCHAUB: RELATIVE ORBIT GEOMETRY 3

Cartesian coordinates ρ. The only linearizing assumption express the relative position coordinates (x, y, z) in terms of that was made is that the relative orbit radius ρ is small the orbit element differences in Eq. (20) through compared to the inertial chief orbit radius r. However, when describing a relative orbit through orbit element differ- r ae sin f x(f) ≈ δa + δM − a cos fδe (22a) ences, it is not convenient to describe the anomaly difference a η through δθ or δf. For elliptic chief orbits, the difference in r 2 true anomaly between two orbits will vary with time. To y(f) ≈ (1 + e cos f) δM + rδω η3 avoid this issue, the desired anomaly difference between two (22b) r sin f orbits is expressed in terms of a mean anomaly difference + (2 + e cos f)δe + r cos iδΩ δM. This anomaly difference will remain constant, assum- η2 ing unperturbed Keplerian motion, even if the chief orbit is z(f) ≈ r(sin θδi − cos θ sin iδΩ) (22c) elliptic. To express the mean anomaly differences in terms of other anomaly differences, we make use of the mean anomaly Note that with this linearized mapping the difference in the definition argument of perigee δω does not appear in the x(f) expression. Further, these equations are valid for both circular M = E − e sin E (13) and elliptic chief orbits. Only the δM and δe terms contribute periodic terms to the radial x solution. Due to the where E is the eccentric anomaly. Taking its first variation dependence of r on the true anomaly f, all orbit element dif- we express differences in mean anomaly in terms of differ- ference terms in the along-track y motion contribute both ences in eccentric anomaly and differences in eccentricity. static offsets as well as periodic terms. For the out-of-plane z motion both the δi and δΩ terms control the out-of-plane ∂M ∂M δM = δE + δe oscillations. By dividing the dimensional (x, y, z) expres- ∂E ∂e (14) sions in Eq. (22) by the chief orbit radius r and making use = (1 − e cos E)δE − sin Eδe of Eq. (11), we obtain the non-dimensional relative orbit coordinates (u, v, w). Using the mapping between eccentric anomaly E and true anomaly f δa e sin f u(f) ≈ + (1 + e cos f) δM a η3 r (23a) f 1 + e E (1 + e cos f) tan = tan (15) − cos fδe 2 1 − e 2 η2 δM and taken its first variation, differences in E are then ex- v(f) ≈ (1 + e cos f)2 + δω η3 pressed as differences in f and e through (23b) sin f + (2 + e cos f)δe + cos iδΩ η sin f δe η2 δE = δf − (16) 1 + e cos f 1 + e cos f η w(f) ≈ sin θδi − cos θ sin iδΩ (23c)

Substituting Eq. (16) into Eq. (14) and making use of the Since (y, z) r, the non-dimensional coordinates (v, w) orbit identities are the angular deputy satellite relative orbit position with respect to the chief orbit radius axis. η2 (1 − e cos E) = (17) However, the present form of Eq. (23) is not convenient (1 + e cos f) to determine the overall non-dimensional shape of the rela- η sin f tive orbit. Reason is that there are several sin() and cos() sin E = (18) (1 + e cos f) functions being added here. Using the identities

p A the desired relationship between diﬀerences in true and A sin t+B cos t= A2 +B2 cos t − tan−1 mean anomalies is found. B (24) 2 p B (1 + e cos f) sin f = − A2 +B2 sin t − tan−1 δf = δM + (2 + e cos f)δe (19) −A η3 η2

Let us redefine the orbit element difference vector δe to as well as standard trigonometric identities, we are able to consist of rewrite the linearized non-dimensional relative orbit motion as T δe = (δa,δM,δi,δω,δe,δΩ) (20) s δa 1 e2δM 2 u(f) ≈ + + δe2 cos(f − f ) Note that all these orbit element differences are constants a η2 η2 u for Keplerian two-body motion. Further, while using q1 and s (25a) 2 2 q2 instead of e and ω allows us to avoid singularity issues eδe e e δM − + + δe2 cos(2f − f ) for near-circular orbits, for the following relative orbit ge- 2η2 2η2 η2 u ometry discussion such singularities do not appear. In fact, describing the relative orbit path using δe and δω instead of e2 δM v(f) ≈ 1 + + δω + cos iδΩ δq1 and δq2 yields a simpler and more elegant result. Using 2 η3 Eqs. (6) and (7), the differences in the qi parameters are s expressed as 2 e2δM 2 + + δe2 cos(f − f ) η2 η2 v (25b) δq1 = cos ωδe − e sin ωδω (21a) s e e2δM 2 δq2 = sin ωδe + e cos ωδω (21b) + + δe2 cos(2f − f ) 2η2 η2 v After substituting Eqs. (19) and (21) into the linear map- p 2 2 2 ping in Eq. (8) and simplifying the result, we are able to w(f) ≈ δi + sin iδΩ cos (θ − θw) (25c) 4 SCHAUB: RELATIVE ORBIT GEOMETRY with the phase angles fu, fv and θw being defined as Deputy Orbit hc Plane eδM f = tan−1 (26a) u −ηδe Chief Orbit Plane hd −1 ηδe π δ f = tan = f − (26b) w v eδM u 2 −1 δi θw = tan (26c) − sin i δΩ δw At these phase angles, the trigonometric terms will reach either their minimum or maximum value. Note that 180 de- Ω i+δi grees can be added or subtracted from these angles to yield i the second extrema point of the trigonometric functions. To further reduce the expression in Eq. (25), let us introduce the small states δu and δw: δΩ s e2δM 2 δ = + δe2 (27a) u η2 p δ = δi2 + sin2 iδΩ2 (27b) Fig. 2 Illustration of Orbit Plane Orientation Difference w between Chief and Deputy Satellites

Using these δu and δw deﬁnitions as well as Eq. (26b), the linearized relative orbit motion is described through of the chief orbit and the angular momentum vector of the deputy orbit. To prove that δw is indeed this angle, let us δa eδe u(f) ≈ − make use of the spherical law of cosines for angles. Using a 2η2 the spherical trigonometric law of cosines, we are able to (28a) 22 δ e relate the angles δΩ, i, δi and δw through: + u cos(f − f ) + cos(2f − f ) η2 u 2 u 2 e δM cos δw = cos i cos(i+δi) + sin i sin(i+δi) cos δΩ (29) v(f) ≈ 1 + + δω + cos iδΩ 2 η3 (28b) δu e Assuming that δΩ, δi and δw are small angles, we approxi- − 2 sin(f − fu) + sin(2f − fu) 2 η2 2 mate sin x ≈ x and cos x ≈ 1 − x /2 to solve for δw.

w(f) ≈ δw cos (θ − θw) (28c) p δ = δi2 + sin2 iδΩ2 (30) Note that the cos(2f) and sin(2f) terms are multiplied by w the eccentricity e. Only if the chief orbit is very eccentric will these terms have a signiﬁcant contribution to the over- Using the angle δw, the out-of-plane motion w(f) in all relative orbit dimension. For the more typical case of Eq. (28c) is written in the compact form shown.12, 23 having a chief orbit with a small eccentricity e, these terms only provide small perturbations to the dominant sin(f) and cos(f) terms. Using Eq. (28), it is trivial to determine the maximum radial, along-track and out-of-plane dimension of Chief Orbits with Small Eccentricity a relative orbit provided that the relative orbit geometry is prescribed through the set of orbit element diﬀerences In this section we assume that the chief orbit eccentricity {δa, δe, δi, δΩ, δω, δM}. Note that this linearized relative e is a small quantity. In particular, we assume that e is orbit motion is valid for both circular and elliptic chief ref- small but greater than ρ/r, while powers of e are smaller erence orbits. The only linearizing assumption made so far is than ρ/r. In this case we only retain terms which are linear that the relative orbit radius is small compared to the planet in e and drop higher order terms of e. The orbit radius r is centric inertial orbit radius. However, note that we are only now approximated as estimating the non-dimensional relative orbit shape. To obtain the true radial, along-track and out-of-plane motions, aη2 we need to mulitpy (u, v, w) by the chief orbit radius r. r = ≈ a(1 − e cos f) (31) Since r is time dependent for an elliptic chief orbit, the 1 + e cos f points of maximum angular separation between deputy and chief satellites may not correspond to the point of maximum while η2 ≈ 1. The linearized dimensional relative orbit mo- physical distance. To plot the dimensional linearized relation in Eq. (22) is written for the small eccentricity case tive orbit motion, we use Eq. (22) instead. However, due to as: the ratio’s of sin() and cos() terms, it is not trivial to obtain the maximum physical dimensions of the relative orbit. Let us take a closer look at the out-of-plane motion. The true latitude angle θ , at which the maximum angular ae sin f w x(f) ≈ (1 − e cos f)δa + δM out-of-plane motion will occur, is given by Eq. (26c). As η (32a) expected, if only a δΩ is prescribed, then the maximum w − a cos fδe motion occurs during the equator crossing at θ = 0 or 180 a degrees. If only a δi is prescribed, then the maximum w y(f) ≈ (1 + e cos f)δM + a(1 − e cos f)δω motion occurs at θ = ±90 degrees. η The maximum angular out-of-plane motion is given by the + a sin f(2 − e cos f)δe (32b) angle δw as shown in Figure 2. This angle δw is the tilt angle + a(1 − e cos f) cos iδΩ of the deputy orbit plane relative to the chief orbit plane. As such, it is the angle between the angular momentum vector z(f) ≈ a(1 − e cos f)(sin θδi − cos θ sin iδΩ) (32c) SCHAUB: RELATIVE ORBIT GEOMETRY 5

Making use of the trigonometric identity in Eq. (24), the then a bounded relative motion will occur. Assuming this (x, y, z) motion is written as constraint is satisfied, then the differential CW equations can now be solved for an analytical solution of the relative x(f) ≈ δa + aδx cos(f − fx) (33a) orbit motion. δM y(f) ≈ a + δω + cos iδΩ x(t) = A0 cos(nt + α) (38a) η (33b) ae y(t) = −2A0 sin(nt + α) + yoff (38b) − aδy sin(f − fy) − sin(2f)δe 2 z(t) = B0 cos(nt + β) (38c) ae z(f) ≈ aδ cos(θ − θ ) − δ cos(2f − f ) z z z z The integration constants A , B , α, β and y are deter- 2 (33c) 0 0 off ae mined through the relative orbit initial conditions. These − (sin ωδi − cos ω sin iδΩ) 2 equations have been extensively used to generate relative orbits if the chief orbit is circular. Let us now compare the with the small states δx, δy and δz defined as predicted (x, y, z) motion in terms of the true anomaly in s Eq. (33) to the CW solution in Eq. (38) if the chief orbit e2δM 2 δa 2 is assumed to be near-circular (i.e. e < ρ/r). In this case δ = + δe + (34a) x η2 a terms containing the eccentricity e are dropped, as compared to the small eccentricity case studied earlier where s δM 2 only higher order terms of e were dropped. Assuming that δ = 4δe2 + e2 − δω − cos iδΩ (34b) y η all δe components are small (i.e. the relative orbit radius is assumed to be small compared to the inertial orbit radius), p 2 2 2 δz = δi + sin iδΩ (34c) and letting e → 0, we find that r → a and η → 1. Fur- ther, note that fx and fy approach 0. Using Eq. (33) the and the phase angles fx, fy, θz and fz defined as relative orbit motion (x(f), y(f), z(f)) is expressed for the near-circular chief orbit special case as ! −1 eδM fx = tan (35a) x(f) ≈ δa − a cos fδe (39a) −η δe + δa a y(f) ≈ a(δω + δM + cos iδΩ) + 2a sin fδe (39b)  δM  e η − δω − cos iδΩ p 2 2 2 f = tan−1 (35b) z(f) ≈ a δi + sin iδΩ cos (θ − θz) (39c) y  −2δe  Note that the maximum width of the oscillatory along-track −1 δi motion y is given by 2aδe. This result has been previously θ = tan (35c) z − sin iδΩ presented in References 11 and 12. Comparing Eqs. (38) and (39) and noting that nt = f for this case, we are able −1 cos ωδi + sin ω sin iδΩ fz = tan (35d) to establish a direct relationship between the CW constants sin ωδi − cos ω sin iδΩ and the orbit element differences. Note that the orbital radial motion x(f) for the small eccen- A = −aδe (40a) tricity case is identical to the general orbit radial coordinate 0 p 2 2 2 in Eq. (22a) if δa is zero. The semi-major axis difference B0 = a δi + sin iδΩ (40b) must be zero for bounded relative motion if no perturba- α = 0 (40c) tions are present. With perturbations present, δa may be non-zero and the orbit radial coordinate will then be differ- β = ω − θz (40d) ent between the linearizing approximations. The estimated yoff = a(δω + δM + cos iδΩ) (40e) along-track motion y(f) and out-of-plane motion z(f) will always be numerically different between the generally ellip- Recall that Eqs. (38) require that the bounded relative mo- tic case and the small eccentricity case. tion constraint is satisfied. Thus the δa term is set to zero The dimensional form of the relative orbit motion in when comparing the two forms of the relation orbit motion Eq. (32) is convenient to determine the amplitudes of the expression. sinusoidal motion in either the along-track, orbit radial or out-of-plane motion. Note that since e is considered small, the double-orbit frequency terms sin(2f) are only a minor Prescribing Bounded Relative Motion perturbutation to the dominant orbit frequency sinusoidal Conditions terms. Eqs. (22), (28), (33) or (39) show how the orbit element differences and the Hill frame Cartesian coordinates must Near-Circular Chief Orbit be related for any chief true anomaly angle f. Since this If the chief orbit is circular or near-circular, then the lin- mapping must hold at any instance of time, however, these earized relative equations of motion are given through the linearized equations also approximate a solution for the rel- famous Clohessy-Wiltshire or CW equations.24 These are ative orbit motion ρ(t). To map between time and the sometimes also referred to as Hill’s equations.20 true anomaly we must solve Kepler’s equation. However, to be able to describe the relative orbit geometry in terms x¨ − 2ny˙ − 3n2x = 0 (36a) of the Hill frame Cartesian coordinates, the forms presented in Eqs. (22), (28), (33) or (39) in terms of the true anomaly y¨ + 2nx˙ = 0 (36b) f are preferred. The reason for this is that by sweeping z¨ + n2z = 0 (36c) f through a complete revolution, the (x, y, z) coordinates found through these equations will yield the linearized rel- Further, these differential equations are only valid if the ative orbit approximation that results due to a prescribed relative orbit radius is small compared to the planet centric set of constant orbit element differences. Note that no dif- orbit radius. If the relative orbit initial conditions satisfy ferential equations are solved here to determine the relative the constraint orbit motion, and that the dominant relative orbit radial (x-direction), along-track (y-direction) and out-of-plane mo- y˙0 + 2nx0 = 0 (37) tion (z-direction) can be trivially extracted. 6 SCHAUB: RELATIVE ORBIT GEOMETRY

However, note that Eqs. (22), (28), (33) or (39) do not the constraint is written in its final form as25 explicitly contain any secular terms. For the classical two- body orbital motion, the only condition on two inertial or- y˙0 −n(2 + e) = p (52) bits to have a closed relative orbit is that their orbit energies x0 (1 + e)(1 − e)3 must be equal. In terms of orbit elements this corresponds to demanding that Let us linearize this constraint about a small eccentricity. In this case terms which are linear in e are retained and higher δa = 0 (41) order terms in e are dropped. The bounded relative motion This constraint is valid for both circular and elliptical chief constraint on the initial Cartesian coordinates is then given orbits. Also, note that this constraint is the precise require- by ment of the Keplerian motion for bounded relative orbit y˙ + (2 + 3e)nx = 0 (53) paths; no linearizations have been made here. For Keple- 0 0 rian two-body motion, all the orbit element differences will The find the initial Cartesian coordinates constraint for naturally remain constant except for the mean anomaly dif- bounded relative motion at the chief orbit apoapses, we set ference. If δa is not zero between two orbits, then these r(t0) = ra = a(1 + e) and follow the same steps. The re- orbits will drift apart due to having different orbit periods. sulting constraint for chief orbits with a general eccentricity In this case δM will not remain a constant, but grow larger is with time. The linearizations in Eqs. (22), (28), (33) or (39) can still be used to predict the relative orbit motion, y˙0 −n(2 − e) but only until the relative orbit radius ρ is no longer small = p (54) x0 (1 − e)(1 + e)3 compared to the inerital chief orbit radius r. Next we would like to see what conditions the δa = 0 while the constraint for chief orbits with a small eccentricty constraint imposes on the initial Hill frame Cartesian coordi- is given by nates X(t0) = X0. Let us use the inverse mapping between orbit element differences δe and the Hill frame Cartesian y˙0 + (2 − 3e)nx0 = 0 (55) coordinates X found in Reference 16 to express δa as Note that if the chief orbit is circular and e = 0, then this δa = 0 = 2α(2 + 3κ1 + 2κ2)x(t) + 2αν(1 − 2κ1 + κ2)y(t) bounded relative motion constraint reduces to the familiar 2α2νp 2a form of + x˙(t) + (1 + 2κ1 + κ2)y ˙(t) (42) Vt Vt y˙0 + 2nx0 = 0 (56) where the parameters which is classically obtained by solving the Clohessy- α = a/r (43) Wiltshire differential equations. The more general bounded ν = Vr/Vt (44) relative orbit constraint in Eq. (52) is valid for eccentric p chief orbits. However, its form requires that t0 be defined κ1 = α − 1 (45) to be at the orbit perigee point. r If we include the J2 perturbation, then the various orbit 2 p κ2 = αν (46) elements will experience short period, long period and sec- r ular drift.26, 27 Mapping the instanteneous osculating orbit are used to simplify the expression. This general constraint elements to mean elements, only the ascending node, argu- can be further simplified by expressing it at the initial time, ment of perigee and mean anomaly will experience secular where t0 is defined as the time where the true anomaly f drifts. It was shown in References 8 and 9 that the following is equal to zero and the satellite is at the orbit periapses. two mean orbit element constraints will make all orbits drift Note that the orbit radius is now given by on average at the same angular rates θ˙ and Ω:˙ r(t ) = r = a(1 − e) (47) η 0 p δη = − tan i δi (57a) 4 Further, the radial velocity V is given by r J δa = 2 (4 + 3η) 1 + 5 cos2 i r δη (57b) h 5 eq V (t ) =r ˙(t ) = (q sin ω − q cos ω) = 0 (48) 2η r 0 0 p 1 2 The scalar parameter req is the equatorial Earth radius. Thus, using Eqs. (44) and (46) we find that ν = 0 and Note that all orbit elements and orbit element differences κ2 = 0. The bounded relative orbit constraint equation is in Eq. (57) are assumed to be mean states, not osculat- now written specifically at the initial time as ing states.26 Whereas a Keplerian motion requires δa to be zero to obtain a repeating, closed relative orbit, if the a a p J2 induced orbit drift is also considered than δa must be 0 = 2 2 + 3 − 1 x0 rp rp rp non-zero if either a change in eccentricity or change in incli- a a p nation is required. However, note that with this constraint + 2 1 + 2 − 1 y˙ (49) ˙ V (t ) r r 0 the mean δθ will be zero, but the mean δω and δM will t 0 p p have secular drifts. The linearized relative motion predic- tions in Eqs. (22), (28), (33) or (39) should be treated as Since Vt(t0) = rpθ˙p and making use of Eq. (47), this constraint is further reduced to the simpler form the mean relative motion where the short and long period components have been removed. Over one orbit the affect 1 of the J2 induced drift will be very small. As such, we can (2 + e)x0 + (1 + e)y ˙0 = 0 (50) θ˙p ignore the small drifts in the orbit element differences and simply hold them constant to evaluate the shape and geome- Let n = pµ/a3 be the chief mean angular motion. Express- try of relative orbit. To see the long term effect of the orbit ing the true latitude rate θ˙ at perigee as elements drift (i.e. study multiple orbit revolutions), the orbit element differences δω and δM must be considered to √ s be slowly time varying.9 Other orbit parameter constraints ˙ h µp 1 + e θp = = = n (51) could be imposed as well, as long as these constraint will r2 a2(1 − e)2 (1 − e)3 p yield a bounded relative orbit. SCHAUB: RELATIVE ORBIT GEOMETRY 7

Numerical Simulations (dropping higher order terms in e) does show some visible The following numerical simulations verify that the rel- departure from the true relative motion. As expected, the ative motion approximation in Eqs. (22), (33) and (39) do circular chief orbit assumption (case 4) yields a very poor indeed predict the spacecraft formation geometry. These prediction of the relative orbit motion. simulations also illustrate the accuracy at which these sim- In Figures 3(c) and 3(d) the RMS relative orbit errors are plified linearized solutions are valid. Let the chief orbit be shown in polar plots versus the chief orbit true anomaly. For given by the orbit elements shown in Table 1. the e = 0.03 simulations, the relative orbit errors for case 2 lie between 20 and 40 meters. Since the relative orbit radius Table 1 Chief Orbit Elements is roughly 10 kilometers, this corresponds to a 0.2–0.4 percent relative motion error. The RMS relative motion error Orbit Elements Value Units for case three is only marginally worse. As was discussed a 7555 km earlier, dropping the higher order e terms should begin to e 0.03 or 0.13 have a noticably affect on the relative motion errors. For i 48.0 deg the e = 0.13 simulations, the relative motion errors for case Ω 20.0 deg 2 lie between 50 and 100 meters (roughly 0.5–1.0 percent ω 10.0 deg errors). However, dropping the higher order e terms in case M0 0.0 deg 3 has a very noticable effect with the relative motion errors growing as large as 500 meters (about 5.0 percent error). The relative orbits are studied for two different chief eccentricities. For the relative orbits studied, the ratio ρ/r is Conclusion about 0.003. The smaller of the two eccentricities considered is already an order of magnitude larger than this, while Analytical linearized relative orbit descriptions are pro- the second eccentricities is even larger again. The numeri- vided for several types of chief orbit eccentricities where cal simulations show that the small eccentricity assumption orbit element differences are chosen to define the relative (i.e. retaining terms in e but dropping higher order terms orbit geometry. The relative orbit motion between a deputy in e) will still yield a reasonable relative orbit prediction for and chief satellite is given in terms of the true anomaly dif- e = 0.03, even though it is larger than the small term ρ/r. ference. With these linearized relative motion solutions, it The orbit element differences which define the relative or- is trivial to estimate what the effect of changing a particular bit are given in Table 2. Since these simulations assume a orbit element difference will be. Orbit element differences two-body Keplerian motion of the satellites, the semi-major have the advantage that they are constants of the Keple- axis difference δa must be zero to achieve a bounded relative rian two-body solution if the mean anomaly difference is motion. selected as the relative anomaly measure. Assuming that a bounded relative orbit constraint is satisfied, then all six or- Table 2 Orbit Element Differences Defining the Space- bit element differences will remain constant. If the bounded craft Formation Geometry relative orbit constraint is not satisfied, or there are other perturbations present such as the J2 gravitational pertur- Orbit Elements Value Units bations, then some or all of the orbit element differences δa 0 km will vary slowly with time. The presented analytical solu- δe 0.00095316 tions in terms of the orbit element differences are still valid. δi 0.0060 deg However, care must be taken to treat the appropriate orbit δΩ 0.100 deg element differences as time varying. The resuting relative δω 0.100 deg orbit motion descriptions can be used both in spacecraft formation flying control developments, as well as in the rel- δM0 -0.100 deg ative orbit design phase. The relative orbit solutions are The following figures compare the relative orbit motion for written such that their secular offset and sinusoidal motions four different cases. Case 1 is the relative motion that will are clearly separated. As such, it is easy to see what the result using the true nonlinear equations of motion. Case 2 offsets and sinusoidal amplitudes will be for a given set of uses the dimensional linearized analytical relative orbit so- orbit element differences in the orbit radial, along-track and lution in Eq. (22). The only assumption that has been made out-of-plane motion. here is that the ratio between the relative orbit radius ρ and the inertial chief orbit radius r is small and terms involving Acknowledgment ρ/r have been dropped. Case 3 assumes that the chief or- I would like to thank Prof. K. T. Alfriend for the fruitful bit eccentricity is small, but not near zero. As such, higher discussions that led to an enhanced version of this paper. order terms in e are dropped, while terms which depend lin- early on e are kept. The relative orbit motion is described through Eq. (33). Case 4 assumes that the chief orbit is References 1Kechichian, J. and Kelly, T., “Analytical Solution of Perturbed near-circular and that e is very close to zero. Any terms Motion in Near-Circular Orbit Due to J2, J3 Earth Zonal Harmonics involving the eccentricity e are dropped here to yield the in Rotating and Inertial Cartesian Reference Frames,” AIAA 27th classical CW equations in Eq. (39). Case 4 is not included Aerospace Sciences Meeting, Reno, Nevada, January 1989, Paper No. 89-0352. here to suggest that a circular orbit assumption should be 2Kechichian, J. A., “The Analysis of the Relative Motion in made when the chief orbit is clearly eccentric. The circular General Elliptic Orbit with Respect to a Dragging and Precessing chief orbit assumption case is included to provide a relative Coordinate Frame,” AAS/AIAA Astrodynamics Specialist Confer- comparison illustrating the extent of the eccentricity effect. ence, Sun Valley, Idaho, August 1997, Paper No. 97-733. 3Queiroz, M. S. D., Kapila, V., and Yan, Q., “Nonlinear Con- The resulting relative orbit motion is illustrated in Fig- trol of Multiple Spacecraft Formation Flying,” Proceedings of AIAA ure 3. Figures 3(a) and 3(b) show the three-dimensional Guidance, Navigation, and Control Conference, Portland, OR, Aug. relative orbits for cases 1 through 4 as seen by the rotating 1999, Paper No. AIAA 99–4270. Hill reference frame. The relative orbit radii vary betwen 10 4Sedwick, R., Miller, D., and Kong, E., “Mitigation of Dif- ferential Perturbations in Clusters of Formation Flying Satellites,” and 20 kilometers. When e = 0.03, note that the relative AAS/AIAA Space Flight Mechanics Meeting, February 1999, Paper orbits for cases 1–3 are virtually indistinguishable. Only the No. AAS 99-124. relative orbit prediction assuming a circular orbit (case 4) 5Melton, R. G., “Time-Explicit Representation of Relative Mo- has a clearly distinct motion. Studying Figure 3(b) with e tion Between Elliptical Orbits,” Journal of Guidance, Control, and = 0.13, the case 2 relative orbit is still indistiniguishable on Dynamics, Vol. 23, No. 4, 2000, pp. 604–610. 6Tschauner, J. and Hempel, P., “Rendezvous zu einem in Ellip- this scale from the true relative motion in case 1. With this tischer Bahn Umlaufenden Ziel,” Astronautica Acta, Vol. 11, 1965, larger eccentricity the relative motion predicted in case 3 pp. 104–109. 8 SCHAUB: RELATIVE ORBIT GEOMETRY

10 case 4 case 3

5 0 rack [km] Along-T 10 case 1,2,3 20 case 1,2 case 4 10 0

Out-of-Plane [km]

-5 Out-of-Plane [km] Out-of-Plane -5 Along-T 5 rack [km] -10 7.5 5 10 Orbit Radial0 [km] -10 12.5 Orbit Radial [km] -5 -2 2 0 a) Relative Orbits in Hill Frame for e = 0.03 b) Relative Orbits in Hill Frame for e = 0.13

case 3

case 2 case 3

f f

o o o case 2 180 0 180 0o 0.02 0.04 0.06km 0.2 0.4 0.6km

c) RMS Relative Orbit Error in kilometers vs. Chief True d) RMS Relative Orbit Error in kilometers vs. Chief True Anomaly Angle f for e = 0.03 Anomaly Angle f for e = 0.13

Fig. 3 Comparison of the Linearized Relative Orbit Solutions for Cases 1–4 with e = 0.03 and e = 0.13.

7Broucke, R. A., “A Solution of the Elliptic Rendezvous Prob- mation Flying Control Using Mean Orbit Elements,” Journal of the lem with the Time as Independent Variable,” AAS/AIAA Space Astronautical Sciences, Vol. 48, No. 1, 2000, pp. 69–87. Flight Mechanics Meeting, San Antonio, TX, January 2002, Paper 14Tan, Z., Bainum, P. M., and Strong, A., “The Implementation of No. AAS 02-144. Maintaining Constant Distance Between Satellites in Elliptic Orbits,” 8 Schaub, H. and Alfriend, K. T., “J2 Invariant Reference Orbits AAS Spaceflight Mechanics Meeting, Clearwater, Florida, Jan. 2000, for Spacecraft Formations,” Celestial Mechanics and Dynamical As- Paper No. 00-141. tronomy, Vol. 79, 2001, pp. 77–95. 15Schaub, H. and Alfriend, K. T., “Impulsive Feedback Control to 9Alfriend, K. T. and Schaub, H., “Dynamics and Control of Establish Specific Mean Orbit Elements of Spacecraft Formations,” Spacecraft Formations: Challenges and Some Solutions,” Journal of Journal of Guidance, Control and Dynamics, Vol. 24, No. 4, July– the Astronautical Sciences, Vol. 48, No. 2 and 3, April–Sept. 2000, Aug. 2001, pp. 739–745. pp. 249–267. 16Schaub, H. and Alfriend, K. T., “Hybrid Cartesian and Orbit 10Garrison, J. L., Gardner, T. G., and Axelrad, P., “Relative Mo- Element Feedback Law for Formation Flying Spacecraft,” Journal of tion in Highly Elliptic Orbits,” AAS/AIAA Space Flight Mechanics Guidance, Control and Dynamics, Vol. 25, No. 2, March–April 2002. Meeting, Albuquerque, NM, Feb. 1995, Paper No. 95-194. 17Naasz, B. J., Classical Element Feedback Control for Spacecraft Orbital Maneuvers, Master’s thesis, Virginia Polytechnic Institute 11Chichka, D. F., “Satellite Cluster with Constant Apparent Dis- and State University, Blacksburg, VA, May 2002. tribution,” Journal of Guidance, Control and Dynamics, Vol. 24, No. 1, Jan.–Feb. 2001, pp. 117–122. 18Alfriend, K. T., Schaub, H., and Gim, D.-W., “Gravitational 12 Perturbations, Nonlinearity and Circular Orbit Assumption Effect on Hughes, S. P. and Hall, C. D., “Optimal Configurations of Rotat- Formation Flying Control Strategies,” AAS Guidance and Control ing Spacecraft Formations,” Journal of the Astronautical Sciences, Conference, Breckenridge, CO, February 2000, Paper No. AAS 00- Vol. 48, No. 2 and 3, April–Sept. 2000, pp. 225–247. 012. 13Schaub, H., Vadali, S. R., and Alfriend, K. T., “Spacecraft For- 19Gim, D.-W. and Alfriend, K. T., “The State Transition Matrix SCHAUB: RELATIVE ORBIT GEOMETRY 9 of Relative Motion for the Perturbed Non-Circular Reference Orbit,” AAS/AIAA Space Flight Mechanics Meeting, Santa Barbara, CA, Feb. 2001, Paper No. 01-222. 20Hill, G., “Researches in the Lunar Theory,” American Journal of Mathematics, Vol. 1, 1878, pp. 5–26. 21DeVries, J. P., “Elliptic Elements in Terms of Small Increments of Position and Velocity Components,” AIAA Journal, Vol. 1, No. 9, Nov. 1963, pp. 2626–2629. 22Beyer, W. H., Standard Mathematical Tables, CRC Press, Inc., West Palm Beach, Fl., 1974. 23Hughes, S. P. and Mailhe, L. M., “A Preliminary Formation Fly- ing Orbit Dynamics Analysis for Leonardo-BRDF,” IEEE Aerospace Conference, Big Sky, Montana, March 11–17 2001. 24Clohessy, W. H. and Wiltshire, R. S., “Terminal Guidance Sys- tem for Satellite Rendezvous,” Journal of the Aerospace Sciences, Vol. 27, No. 9, Sept. 1960, pp. 653–658. 25Inalhan, G. and How, J. P., “Relative Dynamics & Control of Spacecraft Formations in Eccentric Orbits,” Journal of Guidance, Control and Dynamics, Vol. 25, No. 1, Jan.–Feb. 2002, pp. 48–59. 26Brouwer, D., “Solution of the Problem of Artificial Satellite The- ory Without Drag,” The Astronomical Journal, Vol. 64, No. 1274, 1959, pp. 378–397. 27Battin, R. H., An Introduction to the Mathematics and Meth- ods of Astrodynamics, AIAA Education Series, New York, 1987.