TLE two line elements

The main aim of this project is to write a plug and play dynamic MATLAB code that would read any given of two line element (TLE) set data (in a text file format) and first extract and display the identification and status parameters of the given satellite data and, next extract and generate the Time (t)  variation of orbital elements like Inclination(i), Right Ascension of The Ascending Node (RAAN)(Ω), Argument of the Perigee(ω), Semi-Major Axis (a) and Eccentricity(e). Finally, interpretations and inference based on the analysis report is provided to give further context to the work done. For this project, the TLE code is tested on sat23710 with following results like Satellite Identification and Status, Inclination(i) vs Time (t) Variation Graph with trend explanation, Right Ascension of The Ascending Node (RAAN)(Ω)  vs Time (t)  Variation Graph with trend explanation, Argument of the Perigee(ω) vs Time (t)  Variation Graph with trend explanation, Semi-Major Axis (a) vs Time (t)  Variation Graph with trend explanation, Eccentricity(e) vs Time (t)  Variation Graph with trend explanation and final Inferences and Conclusions are to be expected.

Her geçen gün bir çok yeni ülkenin uzaya uydu gönderme becerisi kazanmasıyla beraber özellikle de alçak irtifa uyduların yörüngelerindeki çarpışma riski artmaktadır. Uzay'da aktif uydular, roket parçaları ve pasif uydular, görevine devam edemediği için kontrol edilmeyen uydular, bulunmaktadır. Amerika Birleşik Devletleri Hava Kuvvetleri tüm bu cisimleri tanımlar, takip eder ve kataloglar. Bu işlemleri gerçekleştiriken SGP4 isimli analitik yörünge tahmin metodu ve dünya üzerinde farklı konumlarda bulunan optik teleskop ve radarlar aracılığıyla yapar. Aynı zamanda uzay ortamı farkındalığı için tüm aktif uydular için çarpışma analizi yapar ve daha önce tanımladığı kriterlere göre ilgili uydu sahiplerini çarpışma anıyla ilgili detayları göndermek suretiyle ikaz eder. Fakat, manevra kararı ve sorumluluğu tamamen uydu sahibindedir. Bu çalışma, yapılan ikaz metninde yer alan bilgileri kullanarak optimum manevra parametrelerini hesaplamayı amaçlar. Böylelikle verilen manevra kararı uyduyu muhtemel çarpışma alanından çıkarır ve bunu gerçekleştiriken mevcut kaynakların kullanımını minize eder ve görevin devamlılığına katkıda bulunur. SGP4 yörünge tahmin metodu ile diferansiyel doğrulama metodu ile aktif uydu çarpışma anından manevra kararı verilecek zamana geriye doğru yörünge ötelenir. Hesaplanan başlangıç noktasından itibaren hız vektöründe değişiklik yapılarak tekrar çarpışma anına ileri ötelenir ve çarpışma olasılığında yaratılan değişim gözlenir. Bu çalışma yer istasyonlarında görev alan uydu operatörleri için otonom optimum manevra tavsiyeleri yapabilir. Bunun yanında uydu sahipleri kendi uydularının hassas verilerini de kullanıp bu uygulamanın doğruluğunu arttırabilir. ABSTRACT Collision risk is increasing at Low-Earth Orbit (LEO) altitude for satellites as many new countries can build and launch satellites in orbit. There are active satellites, spent rocket stages and dead satellites in space. US Air Force (USAF) detects, tracks and catalogs these space objects. USAF uses radars and optical telescopes located at different locations on Earth and an analytical orbit determination method called Simplified General Perturbations (SGP4). Moreover, USAF calculates collision probabilities for all active satellites and warns satellite operators using predefined criteria for the sake of Space Situational Awareness (SSA). However, maneuver decision and responsibility belong to the satellite owners. In this paper, a method for calculating optimal maneuver direction, magnitude and time is investigated using the data within Conjunction Summary Message (CSM) published by USAF. Optimal maneuver decision not only avoids the collision by minimizing collision probability but also minimizes the fuel requirement. Because the smallest velocity requirement for the collision avoidance is calculated, the satellite operations will be minimally interrupted. Method is initiated by propagating the orbit of maneuverable satellite from close approach time to maneuver time. Optimal maneuver direction for velocity vector is calculated by defining the problem as a constraint optimization problem and the orbit is propagated to close approach time with the new velocity vector. The probability of collision is calculated for many different maneuver times and magnitudes is calculated and stored for analysis. Satellite controllers can use the method to calculate optimal maneuver for a given collision warning. In addition, satellite owners can use their precise orbit data to increase the accuracy of the method.

Two-line elements are widely used for space operations to predict the orbit with a moderate accuracy for 2-3 days. Local optimization methods, such as the nonlinear least squares method with differential corrections, can estimate a TLE as long as there exists an initial estimate that provides the desired precision. Global optimization methods to estimate TLEs are computationally intensive, and estimating a large number of them is prohibitive. In this paper, the feasibility of estimating TLEs using machine learning methods is investigated. First, a Monte-Carlo approach to estimate a TLE, when there are no initial estimates that provide the desired precision, is introduced. The proposed Monte-Carlo method is shown to estimate TLEs with root mean square errors below 1 km for space objects with varying area-to-mass ratios and orbital characteristics. Second, gradient boosting decision trees and fully-connected neural networks are trained to map orbital evolution of space objects to the associated TLEs using 8 million publicly available TLEs from the US space catalog. The desired precision in the mapping to estimate a TLE is achieved for one of the three test cases, which is a low area-to-mass ratio space object.

One of the factors that determines the lifetime of low-earth orbit satellites is their ballistic coefficient. This study compared the total atmospheric density obtained from data processing of two-line elements (TLE) for LAPAN-TUBSAT satellite with the density based on a standard empirical model MSISE-90. This is accomplished by varying the value of invers ballistic coefficient which is used as an input in calculating the atmospheric density from TLE. Iteration is performed until the atmospheric density value obtained from TLE is considered similar with the value based on MSISE-90. Starting with the inverse ballistic coefficient value obtained from a previous study, we find the range of value of the ballistic coefficient is 71.4 until 100 kg/m2. This value is much smaller than the result of the previous study which is 115.2263 - 230.4527 kg/m2.

With the growing number of resident space objects (RSOs) precise orbit determination and prediction plays an increasingly important role. In the scope of protecting operational spacecraft very accurate orbit prediction is a prerequisite. Predicting close conjunctions and assessing collision probabilities well in advance on a reliable confidence level enables planning and performing of collision avoidance manoeuvres. With sophisticated post-processing methods, when all information is available accuracies in the order of centimeters can be achieved. Post-processing of surveillance data (optical, radar) is still assumed to be in the order of meters. The accuracy of orbit prediction rather than a post-fit of data of common operational products is inferior by orders of magnitudes. This paper addresses the task of orbit propagation based on angle-only observations of RSOs in geostationary and high eccentricity orbits. The focus is on objects which do not perform manoeuvres, to clearly separate the effects of manoeuvre data -which is probably unavailable and/or inaccurate -from modeling the natural forces. The angle-only observations stem from the one-meter ZIMmerwald Laser and Astrometry Telescope (ZIMLAT), the 18-cm Zimmerwald SMall Robotic Telescope (ZimSMART), both located in Zimmerwald, close to Bern, Switzerland, and from the one-meter ESA Space Debris Telescope (ESASDT) located on Tenerife, Spain. The one-meter telescopes have been used for RSO surveys and for so-called follow-up observations over the past decade by the Astronomical Institute of the University of Bern, which is a prime contractor of the European Space Agency (ESA) in the fields of Surveillance and Tracking. Supplementary observations are provided by the courtesy of the International Scientific Optical Network (ISON). Orbit determination and propagation is performed with an enhanced version of the CelMech tool (Beutler, Methods of Celestial Mechanics, Springer 2001). The force model used includes Earth's potential coefficients up to order and degree 12, perturbations due to earth tides, and corrections due to general relativity. Earth shadow passages are modeled. In addition, a model for estimating the direct radiation pressure (DRP) is used, which allows an estimate of the area to mass ratio as a scaling factor, as well as the osculating elements. Optionally, biases, which account inter alia for asymmetries in the observed object, e.g. misalignment of solar panels, can be estimated. Orbit propagation is performed using the determined osculating elements, the estimated area to mass ratio, and the reflection coefficient within the force model mentioned above. The estimated biases are not available for orbit propagation. The ephemerides of the propagated orbits are compared to observations of the same object, which were not used for orbit determination. Those additional angle-only observations serve as ground truth. The additional observations that belong to the same object are validated by a further orbit determination including all observations. The influence of the temporal spacing of observations and the effect of fusing data of different observation sites is investigated. As a reference, two line element data (TLEs) of the US Strategic Command (USSTRATCOM) catalogue are propagated with the SDP4 propagator and the resulting ephemerides are compared to the angle-only observations. Covariances for the TLE data are then estimated.

This paper presents the two element U- slot loaded circular microstrip array antenna for dual band operation. The antenna operates between 2.6 to 7.00 GHz. The antenna has been fabricated with a volume of 9 X 5 X 0.16 cm3 . The maximum bandwidth of 44.90% is achieved. With a peak gain of 5.24 dB is obtained in primary band. The antenna exhibits a broadside and linear radiation characteristics. The results are presented and discussed. This antenna may find its applications in WLAN communication system.

With the growing number of resident space objects (RSOs) precise orbit determination and prediction plays an increasingly important role. In the scope of protecting operational spacecraft very accurate orbit prediction is a prerequisite. Predicting close conjunctions and assessing collision probabilities well in advance on a reliable confidence level enables planning and performing of collision avoidance manoeuvres. With sophisticated post-processing methods, when all information is available accuracies in the order of centimeters can be achieved. Post-processing of surveillance data (optical, radar) is still assumed to be in the order of meters. The accuracy of orbit prediction rather than a post-fit of data of common operational products is inferior by orders of magnitudes. This paper addresses the task of orbit propagation based on angle-only observations of RSOs in geostationary and high eccentricity orbits. The focus is on objects which do not perform manoeuvres, to clearly separate the effects of manoeuvre data -which is probably unavailable and/or inaccurate -from modeling the natural forces. The angle-only observations stem from the one-meter ZIMmerwald Laser and Astrometry Telescope (ZIMLAT), the 18-cm Zimmerwald SMall Robotic Telescope (ZimSMART), both located in Zimmerwald, close to Bern, Switzerland, and from the one-meter ESA Space Debris Telescope (ESASDT) located on Tenerife, Spain. The one-meter telescopes have been used for RSO surveys and for so-called follow-up observations over the past decade by the Astronomical Institute of the University of Bern, which is a prime contractor of the European Space Agency (ESA) in the fields of Surveillance and Tracking. Supplementary observations are provided by the courtesy of the International Scientific Optical Network (ISON). Orbit determination and propagation is performed with an enhanced version of the CelMech tool (Beutler, Methods of Celestial Mechanics, Springer 2001). The force model used includes Earth's potential coefficients up to order and degree 12, perturbations due to earth tides, and corrections due to general relativity. Earth shadow passages are modeled. In addition, a model for estimating the direct radiation pressure (DRP) is used, which allows an estimate of the area to mass ratio as a scaling factor, as well as the osculating elements. Optionally, biases, which account inter alia for asymmetries in the observed object, e.g. misalignment of solar panels, can be estimated. Orbit propagation is performed using the determined osculating elements, the estimated area to mass ratio, and the reflection coefficient within the force model mentioned above. The estimated biases are not available for orbit propagation. The ephemerides of the propagated orbits are compared to observations of the same object, which were not used for orbit determination. Those additional angle-only observations serve as ground truth. The additional observations that belong to the same object are validated by a further orbit determination including all observations. The influence of the temporal spacing of observations and the effect of fusing data of different observation sites is investigated. As a reference, two line element data (TLEs) of the US Strategic Command (USSTRATCOM) catalogue are propagated with the SDP4 propagator and the resulting ephemerides are compared to the angle-only observations. Covariances for the TLE data are then estimated.

This paper presents the two element U-slot loaded circular microstrip array antenna for dual band operation. The antenna operates between 2.6 to 7.00 GHz. The antenna has been fabricated with a volume of 9 X 5 X 0.16 cm 3 . The maximum bandwidth of 44.90% is achieved. With a peak gain of 5.24 dB is obtained in primary band. The antenna exhibits a broadside and linear radiation characteristics. The results are presented and discussed. This antenna may find its applications in WLAN communication system.