Hi guys and gals(adsbygoogle = window.adsbygoogle || []).push({});

I have been given this question and have spent so far around 4 hours pondering calculating researching and reading and have got no where except into a large state of confusion!

Its probably really really simple but i just cant seem to get logical answers and now i have done so many methods i dont know where i am.

Do you mind giving me some guidance? The question is below:

A spacecraft of mass m = 80000 kg is initially in a circular meridianal orbit of altitude H0 = 300 km above the earth’s surface. When the spacecraft is in the position A above the North Pole its two orbital-manoeuvring-system (OMs) engines, each of which has a trust of F = 25 kN, are fired during the time period dt = 35.5 seconds to increase the velocity of the spacecraft and thus to transfer it to the new elliptical orbit. Calculate the altitude H1 of the spacecraft at the apogee point B of the new elliptical orbit. Ignore the change in position of the spacecraft while the engines are on.

Use the following values for the other required parameters:

Radius of the earth: R = 6371 km

Gravity acceleration: g = 9.825 m/s2

the nearest i have is ~410km but got that using hohmann transfer and data that is not given in this question so doesnt really count!

Any help greatly appreciated

Hope this is the right forum to post it :S its my first post! :)

Many Thanks in Advance

Dan

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# Orbital Motion Help

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