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Physics Bootcamp

Section 12.7 The Orbit Equation

Orbit equation gives the analytic expression of the orbit of a planet in a planet-Sun two-body system. We usullay study these oribits in the CM frame, where the orbit equation refers to the orbit of the reduced mass \(\mu\) about the origin, where we place total mass \(M\) at rest. In this approach, we obtain evolution of relative motion of the planet with respect to Sun. The equation of motion will be derived in the Calculus section below.
I will state the final answer here so that we can discuss the physical consequences of the result, which are important for understanding the motion of planets. You may not be surprised that the solution of the equation of motion, Eq. (12.58) for negative energy between minimum of the effective potential and zero effective potential, is an ellipse, which in polar coordinates with origin at one of the foci of the ellipse is
\begin{equation} r = \frac{r_0}{1+e\cos\,\theta}.\tag{12.49} \end{equation}
In terms of the masses, energy, and angular momentum, the parameters, \(r_0\) and eccentricity \(e\text{,}\) of the ellipse are
\begin{equation} r_0 = \dfrac{l^2}{G_N m_1m_2\mu}, \ \ \ e = \sqrt{1 + \dfrac{E}{|U_0|} },\tag{12.50} \end{equation}
where, as before,
\begin{equation*} M = m_1 + m_2,\ \ \mu = \dfrac{m_1m_2}{M},\ \ \ |U_0| = \dfrac{(G_N m_1 m_2)^2 \mu}{2l^2}. \end{equation*}
The semi-major \(a\text{,}\) semi-minor \(b\text{,}\) and focal distance \(c\) of the ellipse are
\begin{equation} a = \dfrac{r_0}{1-e^2},\ \ b = \dfrac{r_0}{\sqrt{1-e^2}},\ \ c = \dfrac{e\,r_0}{1-e^2}.\tag{12.51} \end{equation}
The closest approach, called perihelion for Planet/Sun system and perigee for satellite/Earth system will be
\begin{equation*} r_\text{min} = a - c = \frac{r_0}{1+e}, \end{equation*}
and the maximum distance, called aphelion for Planet/Sun system and apogee for satellite/Earth system
\begin{equation*} r_\text{max} = a + c = \frac{r_0}{1-e}. \end{equation*}
Of course, you can get them by using \(\theta=0\) and \(\theta = \pi\) in the Eq. (12.49). The min and max distances can also be used to get eccentricity of the orbit by
\begin{equation*} e = \frac{r_\text{max} - r_\text{min}}{r_\text{max} + r_\text{min}}. \end{equation*}

Example 12.36. Calculating Elliptical Orbit Properties From Aphelion and Perihelion.

The closest and farthest distance of the Earth from the Sun are \(0.98\,\text{AU}\) and \(1.02\, \text{AU}\text{,}\) where, the Astronomical Unit, \(1\, \text{AU} = 149,598,000\, \text{km}\text{.}\) Find (a) the semi-major axis, and (b) the eccentricity of Earth’s orbit.
Answer.
(a) \(1.0004\, \text{AU}\) (b) \(0.017\text{.}\)
Solution.
Given \(r_\text{max} = 1.02\) AU, \(r_\text{min} = 0.98\) AU. Therefore, the eccentricity \(e\) will be
\begin{equation*} e = \frac{r_\text{max}-r_\text{min}}{r_\text{max}+r_\text{min}} = \frac{1.02-0.98}{1.02+0.98} = 0.02. \end{equation*}
The parameter \(r_0 \text{,}\) the distance at the minimum of the potential, which is the radius of the circular orbit for the given values of the angular momentum, energy, and the masses of the Sun and the planet, for Earth will be
\begin{equation*} r_0 = (1+e) r_\text{min} = 1.02\times 0.98 = 1.00\ \textrm{AU}. \end{equation*}
Therefore, the length of the semi-major axis will be
\begin{equation*} a = \frac{r_0}{1-e^2} = 1.0004 \approx 1.00\ \textrm{AU}. \end{equation*}

Example 12.37. Energy, Angular Momentum and Elliptical Oribit of a Satellite.

A satellite of mass 2000 kg is put in an elliptical orbit of eccentricity 0.5 about Mars, whose mass is approximately \(6.4 \times 10^{23}\) kg and radius \(3.4 \times 10^6\) m. The distance from the center of the planet to the closest approach of the satellite is equal to \(\frac{7}{6}\) times the radius of the planet. (a) Find the distance to the farthest point. (b) Find the energy and angular momentum of the satellite. (c) Find the speed at the closest approach. (d) Using the angular momentum conservation between the farthest and the closest approach, find the speed at the farthest point.
Answer.
(a) \(1.2 \times 10^7\) m. (b) \(E = -5.4 \times 10^9\) J, \(l = 3.18\times 10^{13}\ \textrm{kg.m/s}\text{.}\) (c) \(4,640\, \text{m/s}\text{.}\) (d) \(1550\, \text{m/s}\text{.}\)
Solution 1. a
Let \(R_M\) denote the radius of Mars. Then, we are given
\begin{equation*} r_\text{min} = \frac{7}{6} R_M;\ \ e = 0.5. \end{equation*}
The farthest distance from the planet is given by
\begin{equation*} r_\text{max} = \left(\frac{1+e}{1-e}\right) r_\text{min} = \frac{1.5}{0.5} \frac{7}{6} R_M = 3.5 R_M . \end{equation*}
Solution 2. b
The energy of a satellite of mass \(m\) about a planet \(M\) is written in terms of the semi-major axis \(a\) as
\begin{equation*} E = - G_N\frac{M m}{2 a}, \end{equation*}
with
\begin{equation*} a = \frac{r_\text{max} + r_\text{min}}{2}. \end{equation*}
For the given data we get the following upon expressing all quantities in the SI units.
\begin{equation*} E = - \frac{6.67\times 10^{-11} \times 0.64\times 10^{24} \times 2000}{2\times (7/3) \times 3.38\times 10^{6}} = -5.41 \times 10^{9}\ \textrm{J}. \end{equation*}
The angular momentum written in terms of \(r_\text{min}\) and energy take the following form.
\begin{equation*} l = r_\text{min} \sqrt{\frac{1+e}{1-e}} \sqrt{2\mu |E|}. \end{equation*}
Here \(\mu\) is the reduced mass, which is approximately the mass of the satellite. Putting in the numbers we
\begin{equation*} l = \frac{7}{6}\times 3.38\times 10^{6} \sqrt{3} \sqrt{2\times 2000 \times 5.41 \times 10^{9}} = 3.18\times 10^{13}\ \textrm{kg.m/s}. \end{equation*}
Solution 3. c
The speed at the closest approach can be obtained from the energy we have found. Energy at any distance \(r\) from the planet is given by
\begin{equation*} E = \frac{1}{2} mv^2 - G_N \frac{M m}{r}. \end{equation*}
At closes approach \(r= r_\text{min}\text{.}\) Let \(v_c\) be the speed of the satellite when it is at that point. Since the energy is conserved, we can obtain \(v_{c}\) to be
\begin{equation*} E = \frac{1}{2} mv_c^2 - G_N \frac{M m}{r_\text{min}}\ \ \Longrightarrow\ \ v_{c} = \sqrt{\frac{2}{m} \left( E + G_N \frac{M m}{r_\text{min}} \right)}. \end{equation*}
Now, using \(R_M = 3.4\times10^6\) m and other numerical values we find
\begin{equation*} v_{c} = 4640\ \textrm{m/s}. \end{equation*}
Solution 4. d
The velocities at the closes and farthest points will be perpendicular to the line from the center of mass. Therefore, the magnitude of the angular momenta at these points will be simply \(\mu \times \textrm{distance} \times \textrm{speed}\) giving the following equality with \(v_f\) the speed at the farthest point.
\begin{equation*} \mu r_\text{max} v_f = \mu r_\text{min} v_c. \end{equation*}
Therefore,
\begin{equation*} v_f = \left( \frac{r_\text{min}}{r_\text{max}} \right) v_c = \frac{1}{3} v_c =1550\ \textrm{m/s}. \end{equation*}

Example 12.38. Orbit of a Satellite about the Earth.

A satellite of mass 3000 kg is put in an elliptic orbit about the Earth with the semi-major and the semi-minor axes of the orbit being 10,000 km and 8,000 km respectively. (a) What are the energy and angular momentum of the satellite? (b) How far away is the satellite from the center of the Earth at its nearest approach, perigee, and its farthest point, apogee? (c) What are the speed of the satellite at the perigee and the apogee? Ignore the effect of the Sun.
Solution 1. a
Here, we have the situation \(m \ll M\text{,}\) where \(m\) is the mass of the satellite and \(M\) the mass of Earth. Therefore, we will substitute \(m\) for \(\mu\text{.}\)
We have \(G_NMm = 1.2\times 10^{18}\ \textrm{N.m}^2\text{,}\) \(a = 1.0\times 10^7\ \textrm{m}\text{,}\) and \(b=8.0\times10^6\ m\text{.}\) Therefore, the energy of the satellite is
\begin{equation*} E = -\frac{G_NMm}{2a}= -6.0\times 10^{10}\ \textrm{J}, \end{equation*}
and the angular momentum has the following magnitude.
\begin{equation*} l= b\sqrt{-2\mu E} = 1.5\times 10^{14}\ \textrm{kg.m}^2/\textrm{s}. \end{equation*}
The direction of the angular momentum is pointed perpendicular to the plane of the orbit.
Solution 2. b
We will use the orbit formula to answer these questions. The eccentricity is obtained from \(a\) and \(b\text{.}\)
\begin{equation*} e = \sqrt{1-\frac{b^2}{a^2}} = 0.6. \end{equation*}
The circular orbit radius is
\begin{equation*} r_0 = a(1-e^2) = 6.4\times 10^6\ \textrm{m}. \end{equation*}
The perigee \(r_{\textrm{min}}\) occurs at \(\theta = 180^{\circ}\) and apogee \(r_{\textrm{max}}\) at \(\theta = 0^{\circ}\) in the orbit equation \(r = r_0/(1-e\cos\theta)\text{.}\)
\begin{align*} \amp r_{\textrm{max}} = 1.6\times 10^{7}\ \textrm{m}\\ \amp r_{\textrm{min}} = 2a - r_\text{max}=4.0\times 10^{6}\ \textrm{m} \end{align*}
Solution 3. c
Equating the energy \(E\) of the satellite to the sum of the kinetic and potential energies, we find the speed of the satellite.
\begin{align*} \amp E = \frac{1}{2}\mu v^2 + U \approx \frac{1}{2}mv^2 - \frac{G_N M m }{r} \\ \amp v = \sqrt{\frac{2}{m}\left( E+ \frac{G_N M m }{r} \right)}. \end{align*}
We use the value of energy from part (a) that is same for all \(r\) and obtain the value of speed at the perigee by setting \(r = r_{\textrm{min}}\) and at apogee by setting \(r=r_{\textrm{max}}\) in this equation. The values obtained are \(v_{\textrm{perigee}} = 12,700\ \textrm{m/s}\) and \(v_{\textrm{apogee}} = 3,200\ \textrm{m/s}\text{.}\)

Subsection 12.7.1 Four Types of Orbits

Depending upon energy, we have seen four types of possible orbits. From Eq. (12.50) eccenetricity of orbit depends on energy. The part of energy that is other than the kinetic energy of radial motion, i.e., the term containing \((dr/dt)^2\text{,}\) is called effective potential energy, \(U_\text{eff}\text{.}\) You can see that \(U_\text{eff}\) is sum of regular potential energy and the angular motion part.
\begin{equation} U_\text{eff} = \frac{1}{2}\frac{l^2}{\mu r^2} - \frac{G_N m_1 m_2}{r}.\tag{12.52} \end{equation}
When you plot this with respect to \(r\text{,}\) as shown in Figure 12.39, you find that, depending on energy \(E\text{,}\) there are cases of one turning point and two turning points in the motion. In the case of two turning points, they corresponding to the minimum \(r\) and maximum \(r\text{.}\) These are perihelion and aphelion in the case of Sun-Planet system.
Figure 12.39. The effective potential energy diagram shows four different types of motion depending upon the energy \(E\) of the system. (1) Circular motion when \(E = U_0\text{,}\) the minimum of \(U_\text{eff}\text{,}\) (2) Elliptical motion when \(U_0 \lt E \lt 0 \text{,}\) (3) Parabolic motion when \(E=0\text{,}\) and (4) Elliptical motion when \(E \gt 0\text{.}\)
To discuss the type of orbits that result for various values of energy as compared to the minimum effective potential energy, it is best to rewrite the parameters \(r_0\) and \(e\) in terms of the minimum value of the effective potential energy \(U^{\text{min}}_{\text{eff}}\text{,}\) which we will denote by symbol \(U_0\text{.}\) By an elementary calculation, you can show that the minimum value of the effective potential energy is at
\begin{equation} U_0 = -\dfrac{(G_N m_1 m_2)^2 \mu}{2l^2}\tag{12.53} \end{equation}
Therefore, the parameters \(r_0\) and \(e\) in the orbit equation are
\begin{equation} r_0 = \frac{l}{\sqrt{-2\mu U_0}},\ \ \ \ e = \sqrt{1- \frac{E}{U_0}}.\tag{12.54} \end{equation}
Four Cases:
  1. \(E = 0.\)
    If \(E = 0\text{,}\) then \(e=1\text{.}\) The trajectory is not an ellipse, but a parabola.
  2. \(E \gt 0.\)
    If \(E \gt 0\text{,}\) then \(e \gt 1 \) since \(U_0 \lt 0\text{.}\) This corresponds to a hyperbolic trajectory. This type of object will get closest to the Sun only once. In the case of satellite/planet system, this type of orbit is very uselful in guiding satellites using the influence of gravity of planets. See Figure 12.40 for an example.
  3. \(U_0 \lt E \lt 0. \)
    This case corresponds to \(0 \lt e \lt 1 \text{.}\) Hence, the orbit is an ellipse. This is the case for the bound planetary orbits about the Sun. See Figure 12.41 for an example.
  4. \(E=U_0.\)
    This makes \(e = 0\text{.}\) With \(e = 0\text{,}\) \(r = r_0\text{.}\) Therefore, the trajectory is a circle of radius \(r_0\text{.}\) The orbit of Earth is very close to this case.
Figure 12.40. Hyperbolic orbit for \(r_0 = 1\) unit and \(e = 1.5\) for an attractive potential. The other branch is for a repulsive force since acceleration will be pointed away from the CM.
Figure 12.41. Elliptical orbits for three eccentricities \(0.2, 0.4, 0.6\text{.}\) The thick solid line is a circular orbit which corresponds to \(e = 0\text{.}\) The higher the eccentricity more elongated the ellipse. The value of \(r_0\) was set to 1 for the plots.
Both cases 3 and 4 correspond to bound orbits. Hence planets (as well as comets) have either elliptical or circular orbits around the Sun confirming Kepler’s first law. Checkout eccentricities and other parameters of planets at the NASA website
 4 
nssdc.gsfc.nasa.gov/planetary/factsheet/
.

Remark 12.42. Eccentricities of Planetary Orbits.

Planets (as well as comets) have either elliptical or circular orbits around the Sun. The orbits of different planets differ in their eccentricities. The orbit of Venus is almost a circle while the orbit of Pluto is quite elongated. Comets have much higher eccentricities.
For instance, the orbit of Halley’s comet has an eccentricity of 0.967. Recall that if \(e \gt 1\) the orbit will not be bounded, but hyperbolic. The table summarizes eccentricities of planets. Except for Mercury and Pluto, all other planets have minor eccentricities.
Object Eccentricity (\(e\))
Mercury 0.206
Venus 0.007
Earth 0.017
Mars 0.093
Jupiter 0.048
Saturn 0.055
Uranus 0.051
Neptune 0.007
Pluto 0.252
Sedna 0.8

Subsection 12.7.2 (Calculus) The Equation of Motion for Orbit Equation and Kepler’s First Law

We have seen in previous sections that mechanics of two-body system interacting via a gravitational force is best done in the CM frame, where it suffices to work out the equivalent problem of the orbit of a fictitious particle with mass equal to the reduced mass and center of mass kept at rest at the origin. Once, we solve this equivalent problem of the orbit of the fictitious particle, we can immediately work out the orbits of two masses by utilizing the relation between the coordinates of the reduced mass and the coordinates of the individual masses as we have already discussed.
Furthermore, a two-body system interacting via a gravitational force, it is sufficient to examine the conservation of angular momentum and the conservation of energy in the system. As above, let us denote the conserved angular momentum by \(l\) and conserved energy by \(E\) and work with polar coordinates \(r\) and \(\theta\) in the plane perpendicular to the direction of the constant angular momentum vector.
\begin{align} \amp \mu r^2 \dfrac{d\theta}{dt} = l,\tag{12.55}\\ \amp \dfrac{1}{2}\mu\left( \dfrac{dr}{dt} \right)^2 + U_\text{eff}(r) = E.\tag{12.56}\\ \amp U_\text{eff}(r) = \frac{1}{2}\frac{l^2}{\mu r^2} - \frac{G_N m_1 m_2}{r}\tag{12.57} \end{align}
To get trajectory of \(\mu\) in the two-dimensional plane from these equations, we can deduce the equation obeyed by \(dr/d\theta\) and try to integrate that.
\begin{equation} \frac{dr}{d\theta} = \pm \frac{\sqrt{2\mu}}{l}\,r^2\,\sqrt{E - U_\text{eff}(r)}\tag{12.58} \end{equation}
This task is simplified if we redefine radial and energy in terms of natural scales in this problem, which has the added benefit of absorbing the constants into one overall symbol. We use the minimum of the effective potential energy for scaling purposes. That minimum provides us with a length scale and an energy scale. From \(dU_\text{eff} = 0\text{,}\) we get
\begin{align*} \amp r_0 = \frac{2\alpha}{\beta} = \frac{l^2}{G_N m_1 m_2 \mu}\\ \amp U_0 \equiv U_\text{eff,min} = -\frac{-\beta^2}{4\alpha} = \frac{G_N^2 m_1^2 m_2^2 \mu}{2 l^2}. \end{align*}
Now, we intruduce the following scaled quantities
\begin{align} \amp \rho = \frac{r}{r_0}\tag{12.59}\\ \amp \epsilon = \frac{E}{|U_0|} \tag{12.60} \end{align}
Try to set \(r = r_0\,\rho\) and \(E = |U_0|\,\epsilon\) in (12.58) and perform all the algebra to arrive at
\begin{equation} \frac{d\rho}{d\theta} = \pm \rho^2\,\sqrt{\epsilon^2 - \frac{1}{\rho^2} + \frac{2}{\rho} }. \tag{12.61} \end{equation}
This can be integrated to give
\begin{equation*} \int_{\theta_\text{ref}}^\theta\,d\theta = \pm \int_{\rho_\text{ref}}^{\rho} \frac{d\rho}{\rho^2\sqrt{\epsilon - (1/\rho^2) + (2/\rho) }}, \end{equation*}
where (\(\rho_\text{ref}, \theta_\text{ref}\)) are arbitrary point in the two-dimensional plane. You can look up an integral table to do the integral on the right side. Or, do a couple of substitutions: (1) \(u=1/\rho\text{,}\) (2) \(v-u-1\text{,}\) (3) \(w = v/\sqrt{\epsilon + 1}\text{,}\) (4) \(w = \cos\phi\text{.}\) And, then substitute back. Answer will be
\begin{equation*} \theta - \theta_\text{ref} = \cos^{-1}\left( \frac{(1/\rho) - 1}{ \sqrt{\epsilon + 1}} \right) - \cos^{-1}\left( \frac{(1/\rho_\text{ref}) - 1}{ \sqrt{\epsilon + 1}} \right) \end{equation*}
Now, since \(\theta_\text{ref}\) was arbitrary to begin with, we can choose a value that simplifies the formula. Let’s choose
\begin{equation*} \theta_\text{ref} = \cos^{-1}\left( \frac{(1/\rho_\text{ref}) - 1}{ \sqrt{\epsilon + 1}} \right). \end{equation*}
Then, the relation between scaled radial coordinate \(\rho\) and \(\theta\) will be
\begin{equation*} \cos\theta = \frac{(1/\rho) - 1}{ \sqrt{\epsilon + 1}}. \end{equation*}
Rearranging and setting \(e= \sqrt{\epsilon + 1}\) we see that it the the orbit equation presented above.
\begin{equation*} \rho = \frac{1}{1 + e \cos\theta}, \end{equation*}
Now, we replace the scaled radial variable \(\rho\) by \(r/r_0\text{,}\) where \(r_0\) is the minimum of the effective potential to get the orbit equation give above.
\begin{equation*} r = \frac{r_0}{1 + e \cos\theta}, \end{equation*}

Exercises 12.7.3 Exercises

1. Properties of the Orbit of Mars from its Perihelion and Aphelion.

The closest and farthest distance of Mars from the Sun are 1.38 AU and 1.67 AU. Find (a) the semi-major axis, (b) the eccentricity , and (c) the period of Mars’s orbit.
Solution 1. a,b
Given \(r_\text{max} = 1.67\) AU, \(r_\text{min} = 1.38\) AU. Therefore, the eccentricity \(e\) will be
\begin{equation*} e = \frac{r_\text{max}-r_\text{min}}{r_\text{max}+r_\text{min}} = \frac{1.67-1.38}{1.67+1.38} = 0.095. \end{equation*}
The parameter \(r_0\text{,}\) the distance at the minimum of the potential, which is the radius of the circular orbit for the given values of the angular momentum, energy, and the masses of the Sun and the planet, for Earth will be
\begin{equation*} r_0 = (1+e) r_\text{min} = 1.095\times 1.38 = 1.51\ \textrm{AU}. \end{equation*}
Therefore, the length of the semi-major axis will be
\begin{equation*} a = \frac{r_0}{1-e^2} = \frac{1.51\ \textrm{AU}}{1-0.095^2} = 1.52 \ \textrm{AU}. \end{equation*}
Solution 2. c
Using Kepler’s third law we can determine the period of one planet to the period of another by using the \(a\) of the corresponding orbits. I will use the values of \(a\) and period of Earth to get the period for Mars.
\begin{align*} T_{\textrm{Mars}} \amp = \left(\frac{a_{\textrm{Mars}}}{a_{\textrm{Earth}}} \right)^{3/2} T_{\textrm{Earth}} \\ \amp = \left( \frac{1.52\ \textrm{AU}}{1.00\ \textrm{AU}} \right)^{3/2} \times 1\ \textrm{yr}= 1.87 \ \textrm{yr}. \end{align*}

2. Properties of the Orbit of Mercury from its Perihelion and Aphelion.

The closest and farthest distance of Mercury from the Sun are 0.31 AU and 0.47 AU. Find (a) the semi-major axis, (b) the eccentricity , and (c) the period of Mercury’s orbit.
Answer. b
\(0.206\text{.}\)

3. Closest and Farthest Distances of Halley’s Comet from its Period and Eccentricity.

The orbit of Halley’s comet is approximately elliptical with \(e=0.967\text{.}\) Halley’s comet comes around every 76 years. Find (a) the distance of the closest approach to the Sun and (b) the farthest distance the comet goes from the Sun.
Solution 1. a
We use Kepler’s third law to find \(a\text{,}\) and then from \(a\) and the eccentricity \(e\) we will calculate \(r_\text{min}\text{.}\)
\begin{align*} a_{\textrm{H}} \amp = \left(\frac{T_{\textrm{H}}}{T_{\textrm{E}}} \right)^{2/3} a_{\textrm{E}} \\ \amp = \left( \frac{76\ \textrm{yr}}{1 \ \textrm{yr}} \right)^{2/3} \times 1.02\ \textrm{AU}= 18.3 \ \textrm{AU}. \end{align*}
Therefore, the distance for the closest approach is
\begin{equation*} r_\text{min} = (1-e) a = (1-0.967)\times 18.3 \ \textrm{AU} = 0.604\ \textrm{AU}. \end{equation*}
Solution 2. b
The farthest distance is calculated from \(a\) and \(e\) from
\begin{equation*} r_\text{max} = (1+e) a = (1+0.967)\times 18.3 \ \textrm{AU} = 36.0\ \textrm{AU}. \end{equation*}

4. Circular Orbit of a Satellite around the Earth.

What is the radius of the circular orbit for a satellite of mass 2500 kg about the Earth if it has an energy of $-2\times10^9\ J$?
Answer.
$2.49\times 10^8\ m$.
Solution.
By using the equation of motion $GMm/r^2 = mv^2/r$ and energy $E = mv^2/2 - GMm/r$ we can show that the energy of the satellite in a circular orbit will be
\begin{equation*} E = -\frac{1}{2} \frac{GMm}{r}. \end{equation*}
Therefore,
\begin{equation*} r = -\frac{1}{2}\times \frac{6.67\times 10^{-11}\times5.97\times 10^{24}\times 2500}{-2\times 10^9} = 2.49\times 10^8\ \textrm{m} \end{equation*}

5. Orbit Equation of an Elliptical Orbit in Cartesian Coordinates.

The orbit equation in polar coordinates \((r, \theta)\) is
\begin{equation*} r = \frac{r_0}{ 1 + e\,\cos\theta}. \end{equation*}
For \(0 \lt e \lt 1\text{,}\) this is an equation of an ellipse. Write this in Cartesian coordinate so that it takes the following form.
\begin{equation*} \frac{x^2}{a^2} + \frac{y^2}{b^2} = 1. \end{equation*}
Find the expressions of the semimajor and semiminor axes and the location of the foci.
Hint.
Use \(r\cos\theta = x\) and \(r = \sqrt{x^2 + y^2}\text{.}\) You will also need to displace the origin to the center of the ellipse.
Solution.
Let’s start with the polar form of the orbit equation.
\begin{equation*} r + e r\cos\theta = r_0. \end{equation*}
This is in a coordinate system where origin is at one of the foci, so, we expect to get center at non-zero place. Let’s carry out the substitution of \(r = \sqrt{x^2 + y^2}\) and \(x = r\cos\theta\text{.}\)
\begin{equation*} \sqrt{x^2 + y^2} + e x = r_0. \end{equation*}
Now, move the \(x\) term to the right side, thus isolating the square-root term on one side. Then, we square and rearrange terms to
\begin{equation*} (1-e^2) x^2 + 2 e r_0 x + y^2 = r_0^2. \end{equation*}
We now complete the square on \(x\text{.}\)
\begin{equation*} \left( \sqrt{1-e^2}\, x - \frac{e r_0}{\sqrt{1-e^2}} \right)^2 + y^2 = r_0^2 + \frac{e^2 r_0^2}{1-e^2} = \frac{r_0^2}{1-e^2}. \end{equation*}
Pull the multiplier of \(x\) out and then divide every term by \((1-e^2)\text{.}\)
\begin{equation*} \left( x - \frac{e r_0}{1-e^2} \right)^2 + \frac{y^2}{1- e^2} = \left( \frac{r_0}{1-e^2} \right)^2. \end{equation*}
Now, divide every term by the term on the right side so that we get \(1\) on the right side. We will get
\begin{equation*} \frac{(x - c)^2}{a^2} + \frac{y^2}{b^2} = 1, \end{equation*}
where we get
\begin{equation*} a = \frac{r_0}{1-e^2},\ \ b=\frac{r_0}{\sqrt{1-e^2}},\ \ c = ea. \end{equation*}
Now, we need to change coordinates to the center of ellipse by
\begin{equation*} x' = x - c, \ \ y' = y. \end{equation*}
Then
\begin{equation*} \frac{x'^2}{a^2} + \frac{y'^2}{b^2} = 1. \end{equation*}