Orbital lifetime estimates

Perini, L.L.

doi:10.2172/4269348

Cited by 1 publication

(1 citation statement)

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“…Cowell's form of the system of differential equation subject to the central body gravity and other external forces. The acceleration experienced by the satellite due to atmospheric drag [7] is computed in NPOE using the following vector expression: Hence, it was neglected as being an uncertain parameter and BC = 70 kg/m 2 was taken for the prediction of re-entry time of OSIRIS 3U. The observed semi-major axis and the osculating eccentricity from the TLEs are given in Table 1.…”

Section: Working With Npoe and Matlabmentioning

confidence: 99%

Re-Entry of Space Objects from Low Eccentricity Orbits

Lawrence¹,

Sharma²

2019

IJAA

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This paper deals with the re-entry predictions of the space objects from the low eccentric orbit. Any re-entering object re-enters the Earth's atmosphere with a high orbital velocity. Due to the aerodynamic heating the object tends to break into multiple fragments which later pose a great risk hazard to the population. Here a satellite is considered as the space object for which the re-entry prediction is made. This prediction is made with a package where the trajectory path, the time of re-entry and the survival rate of the fragments is done. The prediction is done using DRAMA 2.0-ESA's Debris Risk Assessment and Mitigation Analysis Tool suite, MATLAB and Numerical Prediction of Orbital Events software. The predicted re-entry time of OSIRIS 3U was found to be on 7th March 2019, 7:25 (UTC), whereas the actual re-entry time was on 7th March 2019, 7:03 (UTC). The trajectory path found was 51.5699 deg. (Lat), −86.5738 deg. (Long.) with an altitude of 168.643 km. But the actual trajectory was 51.76 deg. (Lat), −89.01deg. (Long.) with an altitude of 143.5 km.

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Section: Working With Npoe and Matlabmentioning

confidence: 99%