• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.7
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• pp.616-621
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• 2014

Consumption of the fuel on the satellite operating in low earth orbit, is increased due to the air resistance and the amount of increase makes the satellite lifetime decrease or the satellite mass risen. Therefore the prediction of drag force of the satellite is important. In the paper, drag force and drag coefficient analysis of the parabolic antenna satellite in low earth orbit using direct simulation monte carlo method (DSMC) is conducted according to the mission altitude and angle of attack. To verify the DSMC simulated rarefied air movement, Starshine satellite drag coefficient according to the altitude and gas-surface interaction are compared with the flight data. Finally, from the analysis results, it leads to appropriate satellite drag coefficient for orbit lifetime calculation.

• International Journal of Aeronautical and Space Sciences
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• v.17 no.4
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• pp.579-592
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• 2016

CNUSAIL-1, to be launched into low-earth orbit, is a cubesat-class satellite equipped with a $2m{\times}2m$ solar sail. One of CNUSAIL's missions is to deploy its solar sail system, thereby deorbiting the satellite, at the end of the satellite's life. This paper presents the design results of the attitude control system for CNUSAIL-1, which maintains the normal vector of the sail by a 3-axis active attitude stabilization approach. The normal vector can be aligned in two orientations: i) along the anti-nadir direction, which minimizes the aerodynamic drag during the nadir-pointing mode, or ii) along the satellite velocity vector, which maximizes the drag during the deorbiting mode. The attitude control system also includes a B-dot controller for detumbling and an eigen-axis maneuver algorithm. The actuators for the attitude control are magnetic torquers and reaction wheels. The feasibility and performance of the design are verified in high-fidelity nonlinear simulations.

• Journal of Aerospace System Engineering
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• v.9 no.2
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• pp.57-62
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• 2015

The satellite orbit is continuously changing due to space environment. Especially for low earth orbit, atmospheric drag plays an important role in the orbit altitude decay. Recently, solar activities are expected to be high, and relevant events are occurring frequently. In this paper, analysis on the impact of geomagnetic storm on LEO satellite orbit is presented. For this, real flight data of KOMPSAT-2, KOMPSAT-3, and KOMPSAT-5 are analyzed by using the daily decay rate of mean altitude is calculated from the orbit determination. In addition, the relationship between the solar flux and geomagnetic index, which are the metrics for solar activities, is statistically analyzed with respect to the altitude decay. The accuracy of orbit prediction with both the fixed drag coefficient and estimated one is examined with the precise orbit data as a reference. The main results shows that the improved accuracy can be achieved in case of using estimated drag coefficient.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.205-212
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• 2006

Atmospheric drag is the most significant factor effecting the low Earth satellites under the altitude of 800 km Although the atmospheric density of the low Earth orbit is very low compared to that of the sea level, the accumulated effect of the atmospheric drag slowly lowers the satellite velocity at the perigee. Decrease in velocity at perigee directly causes decrease in altitude at apogee which changes the eccentricity of the orbit. The orbit finally reaches a circular orbit before reentering the Earth. This paper states the methods of calculating the atmospheric drag and the lifetime of the satellite. The lifetime of the kick motor and the satellites which will be used on KSLV-L are calculated by Satellite Tool Kit.

• Advances in aircraft and spacecraft science
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• v.7 no.4
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• pp.371-385
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• 2020

A forward-facing aerospike attached to a payload fairing of a satellite launch vehicle significantly alters its flowfield and decreases the aerodynamic drag in transonic and low supersonic speeds. The present payload fairing is an axisymmetric configuration and consists of a blunt-nosed body along with a conical section, payload shroud, boat tail and followed by a booster. The main purpose of the present numerical simulations is to evaluate flowfield and assess the performance of aerodynamic drag coefficient with and without aerospike attached to a payload fairing of a typical satellite launch vehicle in freestream Mach number range 0.8 ≤ M_∞ ≤ 3.0 and freestream Reynolds number range 33.35 × 10⁶/m ≤ Re_∞ ≤ 46.75 × 10⁶/m whichincludes the maximum aerodynamic drag and maximum dynamic conditions during ascent flight trajectory of the satellite launch vehicle. A numerical simulation has been carried out to solve time-dependent compressible turbulent axisymmetric Reynolds-averaged Navier-Stokes equations. The closure of the system of equations is achieved using the Baldwin-Lomax turbulence model. The aerodynamic drag reduction mechanism is analysed employing numerical results such as velocity vector plots, density and Mach contours in conjunction with the experimental flow visualization pictures. The variations of wall pressure coefficient over the payload fairing with and without aerospike are exhibiting different kind of flowfield characteristics in the transonic and low supersonic speeds. The numerically computed results are compared with schlieren pictures, oil flow patterns and measured wall pressure distributions and exhibit good agreement between them.

• Journal of Astronomy and Space Sciences
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• v.20 no.1
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• pp.11-20
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• 2003

A method for estimating the NORAD SGP4 atmospheric drag term from minimum osculating orbit states, i.e., two osculating orbits, is developed. The first osculating orbit state is converted into the NORAD TLE-type mean orbit state by iterative procedure. Then the converted TLE is propagated to the second orbit state using the SGP4 model with the incremental SGP4 drag term. The iterative orbit propagation procedure is finished when the difference of the two osculating semi-major axes between the propagated orbit and the given second orbit is minimized. In order to minimize the effect of the short-term variations of the osculating semi-major axis, the osculating argument of latitude of the second orbit is propagated to the same argument of latitude of the first orbit. The method is applied to the estimation of the NORAD-type TLE for the KOMPSAT-1 spacecraft. The SGP4 drag terms are estimated from both NORAD SGP4 orbit propagation and the numerical orbit propagation results. Variations of the estimated drag terms are analyzed for the KOMPSAT-1 satellite orbit determination results.

• Journal of Astronomy and Space Sciences
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• v.24 no.2
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• pp.125-134
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• 2007

In this paper, we analyze the orbital variation of the Korea Multi-Purpose SATellite-1(KOMPSAT-1) in a strong space environment due to satellite drag by solar and geomagnetic activities. The satellite drag usually occurs slowly, but becomes serious satellite drag when the space environment suddenly changes via strong solar activity like a big flare eruption or coronal mass ejections(CMEs). Especially, KOMPSAT-1 as a low earth orbit satellite has a distinct increase of the drag acceleration by the variations of atmospheric friction. We consider factors of solar activity to have serious effects on the satellite drag from two points of view. One is an effect of high energy radiation when the flare occurs in the Sun. This radiation heats and expands the upper atmosphere of the Earth as the number of neutral particles is suddenly increased. The other is an effect of Joule and precipitating particle heating caused by current of plasma and precipitation of particles during geomagnetic storms by CMEs. It also affects the density of neutral particles by heating the upper atmo-sphere. We investigate the satellite drag acceleration associated with the two factors for five events selected based on solar and geomagnetic data from 2001 to 2002. The major results can be summarized as follows. First, the drag acceleration started to increase with solar EUV radiation with the best cross-correlation (r = 0.92) for 1 day delayed F10.7. Second, the drag acceleration and Dst index have similar patterns when the geomagnetic storm is dominant and the drag acceleration abruptly increases during the strong geomagnetic storm. Third, the background variation of the drag accelerations is governed by the solar radiation, while their short term (less than a day) variations is governed by geomagnetic storms.

• Bulletin of the Korean Space Science Society
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• 2004.10b
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• pp.278-280
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• 2004

GPS Navigation Solutions are used for operational orbit determination for the KOMPSAT-1 spacecraft. GPS point position data are definitely affected by systematic errors as well as noise. Indeed, the systematic error effects tend to be longer term since the GPS spacecrafts have periods of 12 hours. And then, the overlap method of determining orbit accuracy is always optimistic because of the presence of systematic errors with longer term effects. In this paper, we investigated the measurement noise and the system error for the KOMPSAT-l GPS Navigation Solutions. To assess orbit accuracy with this type of data, we use longer data arcs such as 5-7 days instead of 30 hour data arc. For this assessment, we should require much more attention to drag and solar radiation drag parameters or even general acceleration parameters in order to assess orbit accuracy with longer data arcs. Thus, the effects of the consideration of the drag, solar radiation drag, and general acceleration parameters were also investigated.

• 제어로봇시스템학회:학술대회논문집
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• 1991.10a
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• pp.425-430
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• 1991

For satellite orbit determination, a satellite (K-3H) which is affected by the earth's gravitational field and the earth's atmospheric drag, the sun, and the moon is chosen as a dynamic model. The state vector include orbit parameters, uncertain parameters associated with perturbations and tracking stations. These perturbations include gravitational constant, atmospheric drag, and jonal harmonics due to the earth nonsphericity. Early orbit was obtained with given the predicted orbital parameter of the satellite. And orbit determination, which is applied to Extended Kalman Filter(EKF) for real time implementation , use the observation data which is given by satellite tracking radar system and then orbit estimation is accomplished. As a result, extended sequential estimation algorithm has a fast convergence and also indicate effectiveness for real time operation.

• Journal of Institute of Control, Robotics and Systems
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• v.13 no.10
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• pp.962-969
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• 2007

In this paper, an approach is developed for relative state estimation of satellite formation flying. To estimate relative states of two satellites, the Extended Kalman Filter Algorithm is adopted with the relative distance and speed between two satellites and attitude of satellite for measurements. Numerical simulations are conducted under two circumstances. The first one presents both chief and deputy satellites are orbiting a circular reference orbit around a perfectly spherical Earth model with no disturbing acceleration, in which the elementary relative orbital motion is taken into account. In reality, however, the Earth is not a perfect sphere, but rather an oblate spheroid, and both satellites are under the effect of $J_2$ geopotential disturbance, which causes the relative distance between two satellites to be on the gradual increase. A near-Earth orbit decays as a result of atmospheric drag. In order to remove the modeling error, the second scenario incorporates the effect of the $J_2$ geopotential force, and the atmospheric drag, and the eccentricity in satellite orbit are also considered.