orbit - Determining distance from semi-major axis and eccentricity

Supposing that the mass of the object is negligible compared with the mass of the Earth, you can derive the orbital period $T$ from the 3rd Keplero's law:

$frac{T^2}{a^3} = frac{4pi^2}{G(m_E + m_b)} approx frac{4pi^2}{Gm_E},$

where $a$ is the semi-major. With $T$, for each time istant you also know the mean anomaly $M$, given by (suppose $t = 0$ at perigee):

$M(t) = frac{2pi}{T}t$.

Solving numerically the Keplero's equation for the eccentric anomaly $E$ (where $e$ is the eccentricity)

$M = E - esin E$

and then use the following equation to derive the true anonaly $nu$, which is the angle between the direction of periapsis and the current position of the body, as seen from the Earth:

$ cos nu = frac{cos E - e}{1 - ecos E}$
and
$ sin nu = frac{sqrt{1-e^2}sin E}{1 - ecos E}$.

The distance from the Earth is just given by the orbit equation

$ r = frac{a(1-e^2)}{1 + ecos nu}$.

If I'm not wrong with calculation, it should be:

$ T = 9364 s = 2.6 h.$

$ M(90min) = 207.60°$

$ E(90min) = 203.11°$

$ nu(90min) = 198.95°$

$ r = 11'366 km$

To get the covered distance you should calculate the line integral of the orbit equation.