Eccentricity of Elliptical Orbit given Apogee and Perigee Calculator

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The Two Body Problem

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Elliptical Orbit Parameters

Orbital Position as Function of Time

✖Apogee Radius in Elliptic Orbit represents the maximum distance between an orbiting body and the object it orbits.ⓘ Apogee Radius in Elliptic Orbit [r_e,apogee]			+10% -10%
✖Perigee Radius in Elliptic Orbit refers to the distance between the center of the Earth and the point in a satellite's orbit that is closest to the Earth's surface.ⓘ Perigee Radius in Elliptic Orbit [r_e,perigee]			+10% -10%

✖Eccentricity of Elliptical Orbit is a measure of how stretched or elongated the orbit's shape is.ⓘ Eccentricity of Elliptical Orbit given Apogee and Perigee [e_e]

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Formula

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Eccentricity of Elliptical Orbit given Apogee and Perigee

Formula

`"e"_{"e"} = ("r"_{"e,apogee"}-"r"_{"e,perigee"})/("r"_{"e,apogee"}+"r"_{"e,perigee"})`

Example

`"0.599976"=("27110km"-"6778km")/("27110km"+"6778km")`

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Eccentricity of Elliptical Orbit given Apogee and Perigee Solution

STEP 0: Pre-Calculation Summary

Formula Used

Eccentricity of Elliptical Orbit = (Apogee Radius in Elliptic Orbit-Perigee Radius in Elliptic Orbit)/(Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit)
e_e = (r_e,apogee-r_e,perigee)/(r_e,apogee+r_e,perigee)
This formula uses 3 Variables

Variables Used

Eccentricity of Elliptical Orbit - Eccentricity of Elliptical Orbit is a measure of how stretched or elongated the orbit's shape is.
Apogee Radius in Elliptic Orbit - (Measured in Meter) - Apogee Radius in Elliptic Orbit represents the maximum distance between an orbiting body and the object it orbits.
Perigee Radius in Elliptic Orbit - (Measured in Meter) - Perigee Radius in Elliptic Orbit refers to the distance between the center of the Earth and the point in a satellite's orbit that is closest to the Earth's surface.

STEP 1: Convert Input(s) to Base Unit

Apogee Radius in Elliptic Orbit: 27110 Kilometer --> 27110000 Meter (Check conversion here)
Perigee Radius in Elliptic Orbit: 6778 Kilometer --> 6778000 Meter (Check conversion here)

STEP 2: Evaluate Formula

Substituting Input Values in Formula

e_e = (r_e,apogee-r_e,perigee)/(r_e,apogee+r_e,perigee) --> (27110000-6778000)/(27110000+6778000)

Evaluating ... ...

e_e = 0.599976392823418

STEP 3: Convert Result to Output's Unit

0.599976392823418 --> No Conversion Required

FINAL ANSWER

0.599976392823418 ≈ 0.599976 <-- Eccentricity of Elliptical Orbit

(Calculation completed in 00.004 seconds)

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< 17 Elliptical Orbit Parameters Calculators

True Anomaly in Elliptic Orbit Given Radial Position, Eccentricity, and Angular Momentum

Go True Anomaly in Elliptical Orbit = acos((Angular Momentum of Elliptic Orbit^2/([GM.Earth]*Radial Position in Elliptical Orbit)-1)/Eccentricity of Elliptical Orbit)

Time Period of Elliptical Orbit given Semi-Major Axis

Go Time Period of Elliptic Orbit = 2*pi*Semi Major Axis of Elliptic Orbit^2*sqrt(1-Eccentricity of Elliptical Orbit^2)/Angular Momentum of Elliptic Orbit

Radial Velocity in Elliptic Orbit given True Anomaly, Eccentricity, and Angular Momentum

Go Radial Velocity of Satellite = [GM.Earth]*Eccentricity of Elliptical Orbit*sin(True Anomaly in Elliptical Orbit)/Angular Momentum of Elliptic Orbit

Eccentricity of Elliptical Orbit given Apogee and Perigee

Go Eccentricity of Elliptical Orbit = (Apogee Radius in Elliptic Orbit-Perigee Radius in Elliptic Orbit)/(Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit)

Time Period for One Complete Revolution given Angular Momentum

Go Time Period of Elliptic Orbit = (2*pi*Semi Major Axis of Elliptic Orbit*Semi Minor Axis of Elliptic Orbit)/Angular Momentum of Elliptic Orbit

Elliptical Orbit Time Period given Angular Momentum and Eccentricity

Go Time Period of Elliptic Orbit = (2*pi)/[GM.Earth]^2*(Angular Momentum of Elliptic Orbit/sqrt(1-Eccentricity of Elliptical Orbit^2))^3

Time Period of Elliptical Orbit given Angular Momentum

Go Time Period of Elliptic Orbit = (2*pi)/[GM.Earth]^2*(Angular Momentum of Elliptic Orbit/sqrt(1-Eccentricity of Elliptical Orbit^2))^3

Apogee Radius of Elliptic Orbit Given Angular Momentum and Eccentricity

Go Apogee Radius in Elliptic Orbit = Angular Momentum of Elliptic Orbit^2/([GM.Earth]*(1-Eccentricity of Elliptical Orbit))

Specific Energy of Elliptic Orbit given Angular Momentum

Go Specific Energy of Elliptical Orbit = -1/2*[GM.Earth]^2/Angular Momentum of Elliptic Orbit^2*(1-Eccentricity of Elliptical Orbit^2)

Azimuth-Averaged Radius Given Apogee and Perigee Radii

Go Azimuth Averaged Radius = sqrt(Apogee Radius in Elliptic Orbit*Perigee Radius in Elliptic Orbit)

Semimajor Axis of Elliptic Orbit given Apogee and Perigee Radii

Go Semi Major Axis of Elliptic Orbit = (Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit)/2

Angular Momentum in Elliptic Orbit Given Perigee Radius and Perigee Velocity

Go Angular Momentum of Elliptic Orbit = Perigee Radius in Elliptic Orbit*Velocity of Satellite at Perigee

Radial Velocity in Elliptic Orbit given Radial Position and Angular Momentum

Go Radial Velocity of Satellite = Angular Momentum of Elliptic Orbit/Radial Position in Elliptical Orbit

Angular Momentum in Elliptic Orbit Given Apogee Radius and Apogee Velocity

Go Angular Momentum of Elliptic Orbit = Apogee Radius in Elliptic Orbit*Velocity of Satellite at Apogee

Apogee Velocity in Elliptic Orbit Given Angular Momentum and Apogee Radius

Go Velocity of Satellite at Apogee = Angular Momentum of Elliptic Orbit/Apogee Radius in Elliptic Orbit

Eccentricity of Orbit

Go Eccentricity of Elliptical Orbit = Distance Between Two Foci/(2*Semi Major Axis of Elliptic Orbit)

Specific Energy of Elliptic Orbit given Semi Major Axis

Go Specific Energy of Elliptical Orbit = -[GM.Earth]/(2*Semi Major Axis of Elliptic Orbit)

Eccentricity of Elliptical Orbit given Apogee and Perigee Formula

How does a rocket make enough thrust to escape earths gravity?

It uses a chemical reaction to turn its fuel into very large amounts of gas. The gas exhaust is channeled from the combustion chamber through a hole to a nozzle that directs the gas downward, away from the rocket. Through Newton's Third Law, the force of the escaping gas, carefully oriented, pushes the rocket in the opposite direction.

How to Calculate Eccentricity of Elliptical Orbit given Apogee and Perigee?

Eccentricity of Elliptical Orbit given Apogee and Perigee calculator uses Eccentricity of Elliptical Orbit = (Apogee Radius in Elliptic Orbit-Perigee Radius in Elliptic Orbit)/(Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit) to calculate the Eccentricity of Elliptical Orbit, The Eccentricity of Elliptical Orbit given Apogee and Perigee formula is defined as mathematical expression used to determine the degree of elongation or deviation from a perfect circle in the path of a celestial body or satellite. This formula calculates eccentricity by taking into account the distances of the apogee and the perigee from the center of the central body, such as Earth. Eccentricity of Elliptical Orbit is denoted by e_e symbol.

How to calculate Eccentricity of Elliptical Orbit given Apogee and Perigee using this online calculator? To use this online calculator for Eccentricity of Elliptical Orbit given Apogee and Perigee, enter Apogee Radius in Elliptic Orbit (r_e,apogee) & Perigee Radius in Elliptic Orbit (r_e,perigee) and hit the calculate button. Here is how the Eccentricity of Elliptical Orbit given Apogee and Perigee calculation can be explained with given input values -> 0.599976 = (27110000-6778000)/(27110000+6778000).

FAQ

What is Eccentricity of Elliptical Orbit given Apogee and Perigee?

The Eccentricity of Elliptical Orbit given Apogee and Perigee formula is defined as mathematical expression used to determine the degree of elongation or deviation from a perfect circle in the path of a celestial body or satellite. This formula calculates eccentricity by taking into account the distances of the apogee and the perigee from the center of the central body, such as Earth and is represented as e_e = (r_e,apogee-r_e,perigee)/(r_e,apogee+r_e,perigee) or Eccentricity of Elliptical Orbit = (Apogee Radius in Elliptic Orbit-Perigee Radius in Elliptic Orbit)/(Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit). Apogee Radius in Elliptic Orbit represents the maximum distance between an orbiting body and the object it orbits & Perigee Radius in Elliptic Orbit refers to the distance between the center of the Earth and the point in a satellite's orbit that is closest to the Earth's surface.

How to calculate Eccentricity of Elliptical Orbit given Apogee and Perigee?

The Eccentricity of Elliptical Orbit given Apogee and Perigee formula is defined as mathematical expression used to determine the degree of elongation or deviation from a perfect circle in the path of a celestial body or satellite. This formula calculates eccentricity by taking into account the distances of the apogee and the perigee from the center of the central body, such as Earth is calculated using Eccentricity of Elliptical Orbit = (Apogee Radius in Elliptic Orbit-Perigee Radius in Elliptic Orbit)/(Apogee Radius in Elliptic Orbit+Perigee Radius in Elliptic Orbit). To calculate Eccentricity of Elliptical Orbit given Apogee and Perigee, you need Apogee Radius in Elliptic Orbit (r_e,apogee) & Perigee Radius in Elliptic Orbit (r_e,perigee). With our tool, you need to enter the respective value for Apogee Radius in Elliptic Orbit & Perigee Radius in Elliptic Orbit and hit the calculate button. You can also select the units (if any) for Input(s) and the Output as well.

How many ways are there to calculate Eccentricity of Elliptical Orbit?

In this formula, Eccentricity of Elliptical Orbit uses Apogee Radius in Elliptic Orbit & Perigee Radius in Elliptic Orbit. We can use 1 other way(s) to calculate the same, which is/are as follows -

Eccentricity of Elliptical Orbit = Distance Between Two Foci/(2*Semi Major Axis of Elliptic Orbit)

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