Rocket Space Corporation Energia

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Rocket Space Corporation Energia

Rocket, Russia

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Murtazin R.,Rocket Space Corporation Energia
Proceedings of the International Astronautical Congress, IAC | Year: 2016

The paper describes the reduction of the vehicle autonomous flight duration before docking to the ISS. Due to the limited volume inside Soyuz the reduction of time till docking to the ISS is very important, since the long stay of the cosmonauts in the limited volume adds to the strain of the space flight. In the previous papers of the author it was shown that the existing capabilities of Soyuz, the ISS and the ground control loop make it possible to transfer to the four-orbit rendezvous profile. Since 2012 Russian spacecrafts as Soyuz and Progress began to use this new short rendezvous profile. To date more than 20 spacecrafts have successfully performed quick docking. Despite the success achieved more fast flight from insertion to docking with ISS remains relevant due to cosmonauts' schedule on the launch day which shows that its duration is at the limit of allowable. Over the past years Russian specialists got considerable experience in the preparation of ballistics conditions for quick rendezvous despite a busy schedule of flight operations on the ISS. These ballistics conditions are determined by the allowable angular distance (phase angle) between chaser and target at the chaser's launch time. Also now Soyuz and Progress have autonomous satellite navigation and for their insertion used more accurate launch vehicle. Together, this allows a faster profile. The paper describes new approach which called as "quasi-coplanar insertion", when orbit plane of chaser a little bit differ from orbit plane of target, but with permissible phase angle. This approach is not requires additional changes in spacecraft software and allows to decrease flight duration till docking to 2 orbit or about 3 hour. Before using this profile for Soyuz flight offered to perform several demonstration flights on the Progress for confirmation the correctness of the embedded principles. The paper considers the possible improvements of the proposed approach and recovery from the contingencies. Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved.


Ulybyshev Y.,Rocket Space Corporation Energia
Journal of Guidance, Control, and Dynamics | Year: 2014

A new geometric approach for discontinuous coverage constellations in circular orbits is presented. Satellite constellations for zonal coverage with revisit timesofmore than onetotwo orbital periods are considered. The method is based on two-dimensional maps of visibility properties. The space dimensions are the right ascension of the ascending node (in the Earth-centered inertial frame) and the time. A new geometric pattern named a coverage belt is introduced. The coverage belt with sawtooth boundaries is formed by visibility intervals for a satellite at consecutive revolutions and the two-dimensional space is divided into two parts, with the revisit time near to and more than one orbital period, respectively. The revisit times for the satellite constellations are analytically computed as the sum of two terms: 1) the revisit time between straight-line envelopes of adjacent sawtooth coverage belts, and 2) the relative shift of the teeth in the belts. The last term is more than one and less than two orbital periods. Various types of interfaces between coverage belts are considered. Qualitative aspects of the solutions for well-known Walker-type constellations are discussed. Results are validated using extensive computer simulation and known solutions. Numerical examples of constellations with three to four satellites for discontinuous coverage of a specified latitude band are presented. Copyright © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


Ulybyshev Y.,Rocket Space Corporation Energia
Advances in the Astronautical Sciences | Year: 2012

For almost continuous coverage of light Earth's areas can be used new pseudo-sun-synchronous, highly elliptic orbits using the critical inclination and orbit apogee in the Earth's hemisphere with coverage areas. In opposite to the classical Molniya-type orbits, the new orbits are synchronized with the solar time (or the solar day) that the apogee time is coincident with the local noon of a coverage area or reference meridian. Qualitative analysis and long-term simulation results for one satellite and constellations with 2-4 satellites in the orbits are presented. For such coverage, the constellations required less satellites than highly elliptical, Molniya-type orbit constellations for continuous zonal or regional coverage.


Ulybyshev Y.,Rocket Space Corporation Energia
AIAA Guidance, Navigation, and Control Conference 2011 | Year: 2011

Trajectory optimization methods for low-thrust spacecraft proximity maneuvering at near-circular orbits with interior-points constrains are presented. The methods use discretization of the spacecraft trajectory on segments and sets of pseudoimpulses for each segment. A matrix inequality on the sum of the characteristic velocities for the pseudoimpulses is used to transform the problem into a large-scale linear programming form. Terminal boundary conditions are presented as a linear matrix equation. For the interior-points constraints the matrices must be corresponding extensions. This approach gives flexible possibilities for computation of trajectories with operational constraints. An optimal number of the maneuvers is automatically determined. As an application example, planar trajectories for fly-around of a target spacecraft with a constant range are considered. Specified relative motion trajectories can be presented by sets of interior-points constraints in a form of equalities or double-sided inequalities. The second example is optimization of a spatial collision avoidance trajectory with a minimum allowable range and return to the initial trajectory. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


Ulybyshev Y.,Rocket Space Corporation Energia
Journal of Guidance, Control, and Dynamics | Year: 2010

A method of optimal control problem solutions based on a new concept of pseudocontrol sets is presented. This approach combines large-scale linear programming algorithms with the well-known discretization of the continuous system dynamics on small segments and uses discrete pseudocontrol sets, which are considered independently for each segment. Every set is expressed as a mesh approximation of an admissible control space. The method is associated with significant increases in the number of decision variables and requires introducing artificial variables or pseudovariables. Terminal conditions are presented as a linear matrix equation. An extension of the matrix equation for the sums of the pseudocontrols is used to transform the problem into a linear programming form. Interior-point inequality constraints are represented as a linear matrix inequality. The resulting linear programming form is characterized by matrices that are very large and sparse. The number of decision variables is on the order of tens of thousands. In modern linear programming, there are effective interior-point algorithms to solve such problems.Aminimum path-planning problem with nonlinear constraints, reentry trajectory optimization with maximum cross range, and maximum-radius orbit transfer are considered as application examples. The results of the last example are almost coincident with known solutions using other methods. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.


Zaborsky S.,Rocket Space Corporation Energia
Journal of Guidance, Control, and Dynamics | Year: 2014

An optimization solution of the two-impulse coplanar transfer of a spacecraft in Earth's gravitational field has been obtained on the basis of Lawden's primer-vector theory development. This analytical solution minimizes the total velocity impulse based on the transfer semilatus rectum and on the transfer angle between fixed points of the initial and final orbits. The primer vector satisfies the necessary and sufficient conditions for minimum on the transfer semilatus rectum and on the transfer angle. The global minimum of the total velocity impulse was obtained. It is shown that setting one of Lawden's constants to zero in the primer-vector expression complies with the necessary optimality condition of the transfer angle. It follows that the direction of the optimal velocity impulses lies between the normal in-plane and cotangential directions.


Ulybyshev Y.,Rocket Space Corporation Energia
Journal of Guidance, Control, and Dynamics | Year: 2015

A more universal method for a discontinuous coverage analysis that will be suitable from satellite constellations to arbitrary sets of satellites in circular orbits. The approximate approach is based on two-dimensional maps of visibility properties and can be used for any type of satellite constellations and is applicable to arbitrary sets of satellites. The presented method is based on the sequential unifications of the visibility longitude ranges from revolution to revolution, subject to the ground-track shift. For each satellite, the visibility times are calculated for all of the revolutions at an analysis interval as a list including the times, satellite numbers, and revolution numbers. In the analysis, the evolution of orbital elements, and/or the discrete changes due to orbital maneuvers, can also be considered.


Murtazin R.,Rocket Space Corporation Energia | Petrov N.,Rocket Space Corporation Energia
Acta Astronautica | Year: 2012

Reduction of flight duration after insertion till docking to the ISS is considered. In the beginning of the human flight era both the USSR and the USA used short mission profiles due to limited life support resources. A rendezvous during these missions was usually achieved in 1-5 revolutions. The short-term rendezvous were made possible by the coordinated launch profiles of both rendezvousing spacecraft, which provided specific relative position of the spacecraft or phase angle conditions. After the beginning of regular flights to the orbital stations these requirements became difficult to fulfill. That is why it was decided to transfer to 1- or 2-day rendezvous profile. The long stay of a crew in a limited habitation volume of the Soyuz-TMA spacecraft before docking to the ISS is one of the most strained parts of the flight and naturally cosmonauts wish to dock to the ISS as soon as possible. As a result of previous studies the short four-burn rendezvous mission profile with docking in a few orbits was developed. It is shown that the current capabilities of the Soyuz-FG launch vehicle and the Soyuz-TMA spacecraft are sufficient to provide for that. The first test of the short rendezvous mission during Progress cargo vehicle flight to the ISS is planned for 2012. Possible contingencies pertinent to this profile are described. In particular, in the majority of the emergency cases there is a possibility of an urgent transfer to the present 2-day rendezvous profile. Thus, the short mission will be very flexible and will not influence the ISS mission plan. Fuel consumption for the nominal and emergency cases is defined by statistical simulation of the rendezvous mission. The qualitative analysis of the short-term and current 2-day rendezvous missions is performed. © 2012 Elsevier Ltd.


Ulybyshev Y.,Rocket Space Corporation Energia
Journal of Propulsion and Power | Year: 2015

Optimization methods are presented for long-termstation keeping of a low-thrust space station in lunar halo orbits for unstable collinear libration points. The methods use discretization of each half-revolution into segments and a set of pseudoimpulses for each segment. A matrix inequality of the sum of the characteristic velocities for the pseudoimpulses is used to transform the problem into a large-scale linear programming form. Terminal constraints are presented as a linear matrix equation using numerical partial derivatives along a reference orbit. Possible stationkeeping strategies for quasi-periodic orbits in the vicinity of the Earth-moon collinear libration point L2 are considered. An iterative shooting algorithm, based on the full ephemeris model and differential correction, is used. Preliminary simulation and comparative results of long-term station-keeping strategies for one year are presented. As application examples, two quasi-periodic orbits are considered. The first is similar to a planar Lyapunov orbit. The second is a halo orbit, with continuous visibility of the space station from the Earth. An algorithm for very low-thrust station keeping with long-duration burns is also presented. The algorithm can be used to estimate the required thrustto-weight ratio of the space station. Copyright © 2013 by the American Institute of Aeronautics and Astronautics,Inc. All rights reserved.


Murtazin R.,Rocket Space Corporation Energia | Petrov N.,Rocket Space Corporation Energia
Acta Astronautica | Year: 2014

The paper describes the reduction of the vehicle autonomous flight duration before docking to the ISS. The Russian Soyuz-TMA spacecraft dock to the ISS two days after launch. Due to the limited volume inside Soyuz-TMA the reduction of time until docking to the ISS is very important, since the long stay of the cosmonauts in the limited volume adds to the strain of the space flight. In the previous papers of the authors it was shown that the existing capabilities of Soyuz-TMA, the ISS and the ground control loop make it possible to transfer to the five-orbit rendezvous profile. However, the analysis of the cosmonauts' schedule on the launch day shows that its duration is at the allowable limit and that is why it is necessary to find a way to further reduce the flight duration of Soyuz-TMA before docking to less than five orbits. In a traditional rendezvous profile, the calculation of rendezvous burns begins only after determination of the actual vehicle insertion orbit. The paper describes an approach in which the first two rendezvous burns are performed as soon as the spacecraft reaches the reference orbit and the values of the burns are calculated prior to the launch based on the pre-flight data for the nominal insertion. This approach decreases the duration of the rendezvous by one orbit. The demonstration flight of a Progress vehicle using the proposed profile was implemented on August 1, 2012 and completely confirmed the correctness of the imbedded principles. The paper considers the possible improvements of the proposed approach and recovery from the contingencies. © 2013 Elsevier Ltd. All rights reserved.

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