Analysis of the Accuracy of Ballistic Descent from a Circular Circumterrestrial Orbit

The problem of the transportation of the results of experiments and observations to Earth every so often appears in space research. Its simplest and low-cost solution is the employment of a small ballistic reentry spacecraft. Such a spacecraft has no system of control of the descent trajectory in the atmosphere. This can result in a large spread of landing points, which make it difficult to search for the spacecraft and very often a safe landing. In this work, a choice of a compromise scheme of the flight is considered, which includes the optimum braking maneuver, adequate conditions of the entry into the atmosphere with limited heating and overload, and also the possibility of landing within the limits of a circle with a radius of 12.5 km. The following disturbing factors were taken into account in the analysis of the accuracy of landing: the errors of the braking impulse execution, the variations of the atmosphere density and the wind, the error of the specification of the ballistic coefficient of the reentry spacecraft, and a displacement of its center of mass from the symmetry axis. It is demonstrated that the optimum maneuver assures the maximum absolute value of the reentry angle and the insensitivity of the trajectory of descent with respect to small errors of orientation of the braking engine in the plane of the orbit. It is also demonstrated that the possible error of the landing point due to the error of specification of the ballistic coefficient does not depend (in the linear approximation) upon its value and depends only upon the reentry angle and the accuracy of specification of this coefficient. A guided parachute with an aerodynamic efficiency of about two should be used at the last leg of the reentry trajectory. This will allow one to land in a prescribed range and to produce adequate conditions for the interception of the reentry spacecraft by a helicopter in order to prevent a rough landing.     

A Concept of Guidance for an Air-Launch Vehicle with Compensation of Initial Downrange and Launch Time Errors at Direct Insertion into a Rendezvous Point in the Orbit

A flying launcher (airplane carrier) can generate initial errors in position and time of launch. In order to compensate for these errors, one should have two control parameters in addition to those that provide for a spacecraft's injection into a preset orbit. We suggest the concept of controlling the trajectory of injection by choosing thrust values (within allowable regions of control) of second-stage engines or/and of a space booster of the Polyot carrier launcher. As an example, a rendezvous of the spacecraft at the end of its boost phase with the International Space Station (ISS) is considered. The methodology of the suggested approach can be extended to other mobile systems of launch to rendezvous orbits.     

A robust guidance algorithm for the moon landing

The paper deals with a choice of the rational trajectory of motion of a landing module designed for the Moon landing, from the moment of its de-orbiting from the near-lunar orbit up to landing. An integrated conceptual basis is used to develop multistep terminal algorithms for guidance for the three segments of the descent.     

Robust algorithm of control for a maneuver in the Martian atmosphere to insert a spacecraft into the planet’s satellite orbit

When inserting a satellite into an orbit around Mars with the use of aerodynamic drag, it is required to apply a robust algorithm which is capable of being adapted to the actual conditions of the planet’s atmosphere. We suggest a method of adaptation taking into account the specific features of the maneuver including descending and ascending legs of the trajectory. It is demonstrated that the algorithm is efficient when disturbances of the density of the Martian atmosphere increase by a factor of 2–3.     

An Algorithm for Safe Landing of a Spacecraft during Descent from a Circumlunar Orbit

Cosmic Research - The problem of safe landing of a spacecraft with a combined propulsion system from a circumlunar orbit to a given place on the lunar surface is considered. Landing safety is...     

Mars orbiter insertion by use of atmospheric deceleration

The multistep adaptive control algorithms with trajectory prediction for Mars orbiter insertion by use of atmospheric deceleration are described. The navigation task is proposed to be solved simultaneously with defining the control law throughout the entry. The stabilization algorithm ensures a low fuel consumption and acceptability of transient process, even under a considerable disturbing aerodynamic torque. Digital computer simulation has been used for the algorithm development and evaluation. The simulation results have shown high effectiveness of the proposed control methods.