Annotation: The article is devoted to optimization of a trajectory of the antiaircraft guided missile performed in design phase. The review of existing solutions on this issue confirmed the topicality of the problem. The analytical solution cannot be obtained, therefore, according to modern tendencies, optimization by numerical method of original development was performed. The basis of the method is two-level optimization which is carried out, in turn, by two different numerical methods and for two different criteria functions. At the top level, by method of random search and as a variant, by method of coordinate descent, the search was carried out for a fixed set of intermediate for the specified flight range trajectory points which co-ordinates in aggregate provide the necessary optimum. At the bottom level, for each pair of consecutive intermediate points, the boundary problem of falling into distant point by one-dimensional optimization was solved. The coordinate descent method was used for search for the simplified flight program. As optimization criteria for top level, minimum flight time or maximum final speed, for bottom  terminal criterion were used. The control program selected the angle of attack  program. As a result, the optimum and suboptimum (additionally ensuring minimum calculation time) trajectories and flight programs to maximum range and different altitudes were obtained. The analysis of results showed practical proximity of trajectories of minimum flight time and maximum final speed.

Annotation: For monitoring rocket flight and determining accuracy of spacecraft injection into the planned orbit, it is necessary to ensure the reception of telemetry data from the launch vehicle. Telemetry data receiving stations may be located either on land or on shipboard. When the antenna system of such station is placed on shipboard, ship roll and ship drift have the most considerable impact on the antenna guidance accuracy. To ensure the guidance accuracy of the telemetry receiving antenna set, placed on shipboard, the control algorithm was designed. It was offered to use triaxial rotary support with axis of reflector inclination angle to meet the requirements specified. In the article, the connection between kinematic parameters of the antenna rotary support drives and parameters of the space launch vehicle motion were identified, rotation angles of the antenna drives along the three axes were determined, and the law of angular velocity variation along the azimuthal axis, including the maximum feasible angular velocity provided by the azimuthal axis drive, was chosen. Numerical simulation of antenna guidance algorithms that provide stable signal receiving under conditions of ship roll was carried out in the visual development environment of Embarcadero RAD Studio XE6. Several variants for operation of the rotary support drives of the antenna set were chosen for mathematical simulation; disturbing conditions of ship roll and ship drift were analyzed and chosen for ships with small displacement. The simulation validated the designed antenna control algorithm and showed that the requirements for the cinematic parameters of the antenna drives were reduced under conditions of ship roll when the axis of reflector inclination angle was introduced; and accelerometer unit or GPS receiver installed in the antenna structure additionally increased the accuracy of target designation of the antenna and improved its guidance accuracy

Annotation: The main task of the radiotransparent radome design is to develop a structure that reliably protects the aircraft components, sensitive instrumentation of the guidance system from incoming air flow and heating in all phases of operation, meeting all the requirements of minimum weight, allowable temperature under the radome, strength, centre of mass and radio characteristics. Development of the radome solves the complex task of coupling and optimization of the geometric aspects of design with physical-mechanical, radio and thermal properties of materials. The article dwells on the aspects of development of the radiotransparent radomes depending on the aircraft flight speed; basic requirements for the selection of outside perimeter configuration; structural materials, providing the required radio properties in combination with minimum mass and necessary strength, acceptable temperature inside the fairing. Yuzhnoye-developed radome for the up to 5 M aircraft was considered, consisting of ogive shell made of АФ-10ПО fiberglass, heat-resistant tip and structural ring made of aluminum alloy. Methods of thermal and strength analyses of the shell are considered, results of calculations confirming the fitness for work of the structure are presented. Options of configuration with ceramic shells and methods of their coupling with rings are presented for the aircraft that reaches up to 5 M speed. Basic principles of radiotransparent radomes development and aspects of ceramic shells coupling with structural rings are considered, as well as the results of strength analyses and thermal calculations that ensure the performance of the structure and equipment in the area under the radome.

Annotation: This article considers the problem of determination of propulsion system solid fuel burn-out time in the extraatmospheric flight segment taking the apparent acceleration and apparent speed measured by the inertial navigation system. Correlation analysis of the realized and nominal dependencies of the apparent acceleration and apparent speed of the launch vehicle on relative operating time of the propulsion system is suggested to be used to forecast the fuel burn-out time. In order to improve the accuracy of the forecast, and to decrease the amplitude and vibration rate of its results several channels simultaneously are suggested to be used for calculations with subsequent majority voting and digital filtration. As a result of the study, the procedure to forecast the time of solid fuel burn-out in the launch vehicle propulsion system in flight has been developed. Operability of the suggested procedure has been verified using the mathematical simulation of the launch vehicle flight for two operating modes of the propulsion system different from the nominal ones. Based on the statistical processing of the deviations of the predicted time of solid fuel burn-out versus the realized one it was determined that the forecast based on the results of apparent acceleration measurement has the greatest accuracy with the minimal number of operations. Suggested procedure is easily realized as the multistage adaptive algorithm and can be used in the guidance system of the solid-propellant launch vehicle in the extra-atmospheric flight segment for the numerical forecast of the reachable terminal parameters of flight, definition of command vector and development of the relevant thrust vector control commands.