STK Training Session 8: Transferring a Satellite between Circular Orbits Rev 1.1 Sept 23, 2003
In this session you are tasked with transferring a satellite from an inner circular orbit to an outer circular
orbit. This requires two firings of a rocket engine aboard the satellite. This type of transfer is called a
Hohmann transfer, as it uses the least amount of energy to transfer the satellite. We will follow example 61 from Understanding Space, in which NASA wants to transfer a communications satellite from LEO to
GEO. The questions to answer are: what is the total delta vee for this transfer and how long will it take?
Transfer orbit
Inner
orbit
orbit
orbit
v1
v2
Outer orbit
 for the Earth: 3.986005 x 1014 m3/s2.
Radius of the Earth: 6378.137 km
The radius of LEO for this example is taken to be 6570 km. The radius of GEO is 42,160 km.
Following the procedure for calculating Hohmann transfer orbit parameters, calculate:
1) The specific mechanical energy of the inner orbit:
 = ________________
2) The velocity of the satellite in the inner orbit:
v1 = _________________
3) The semi-major axis of the transfer orbit:
at = __________________
4) The specific mechanical energy of the transfer orbit:
t = __________________
5) The eccentricity of the transfer orbit:
et = __________________
6) The specific angular momentum of the transfer orbit:
ht = __________________
7) The velocity of perigee of the transfer orbit:
vtp = _________________
8) The delta vee to go from inner orbit to the transfer orbit:
v1 = ________________
9) The time of flight along the transfer orbit:
Pt = __________________
10) The velocity of apogee of the transfer orbit:
vta = __________________
11) The specific mechanical energy of the outer orbit:
2 = __________________
12) The velocity of the satellite in the outer orbit:
v2 = ___________________
13) The delta vee to go from apogee of the transfer orbit to the outer orbit: v2 = _______________
STK Session 8
page 2 Rev 1.1
Sept 23, 2003
Launch STK.
Create a new scenario named Hohmann.
Select scenario, basic properties, and change the time period from one day to three days.
Create a new satellite.
Highlight the satellite, select basic properties, and change the propagator to Astrogator.
Segment 1: inner orbit
Right click on Initial State. Select Properties. Change the name from Initial State to inner orbit. Change
the element type to Modified Keplerian. Set the radius of periapsis to 6570 km. Set the Eccentricity to
0.000. Set the Inclination to 28.5 degrees. Click on Satellite properties, and set the fuel mass to 5000 kg.
Segment 2: propagate
Right click on inner orbit, and Insert After: Propagate. With this segment highlighted, change the
Propagator to Earth Point Mass. Set the trip duration to 7200 sec (which is two hours).
Segment 3: delta vee 1
Right click on the propagate segment, and Insert After: Impulsive Maneuver. Right click on this new
segment, select properties, and change the name to delta vee 1. (That represents the first change in velocity
maneuver.) Select Thrust Vector for Attitude Control. Select Cartesian for Vector Type. Select VNC for
Thrust Axes. Set the X(Velocity) component to your answer to calculation #8. Turn ON the Decrement
Mass Based on Fuel Usage option.
Segment 4: transfer ellipse
Right click on delta vee 1, and Insert After it a propagate segment. Right click on this new segment, select
Properties, and rename it Transfer Ellipse. Change its color to red. Select Earth Point Mass as the
propagator. Select the value of Duration to match your answer to calculation #9.
Segment 5: delta vee 2
Right click on the Transfer Ellipse segment and Insert After: Impulsive Maneuver. Right click on this new
segment, select Properties, and rename it delta vee 2. Select Thrust Vector for Attitude Control. Select
Cartesian as the Vector Type. Select VNC as the Thrust Axes. Set X (Velocity) to your answer to
calculation #13. Turn ON the Decrement Mass Based on Fuel Usage option.
Segment 6: propagate
Right click on delta vee 2, and Insert After: propagate. Right click on this new segment, and rename it as
Outer Orbit. Set its color to white. Select Earth Point Mass as the propagator. Set the Duration to at least
86400 sec.
At last, you are ready to animate your satellite. Click the green Run button labeled, “Go.” Animate the
scenario. Rotate your view of the satellite to look down from above the North Pole region of the Earth.
Zoom out to see the full trajectory.
Determine/verify the radius of your satellite’s outer orbit. Do this by highlighting the satellite in the main
browser window, then select Tools: report. Can you finish from here yet? Good. (Just in case, from the
dialog window, select ECF LLR Position. Then hit the Create Report button. Scroll down the report to
view the report at the end of the trajectory.)
Satellite outer orbit radius: _______________________