Annotation: The scientific and methodological propositions for the designing single-stage guided missiles with the solid rocket motors for advanced multiple launch rocket systems are defined. The guided missiles of multiple launch rocket system are intended for delivering munitions to the given spatial point with required and specified kinematic motion parameters at the end of flight. The aim of the article is an analysis of the development trends of the guided missiles with the solid rocket motors for the multiple launch rocket systems, identifying the characteristics and requirements for the flight trajectories, design parameters, control programs, overall dimensions and mass characteristics, structural layout and aerodynamic schemes of missiles. The formalization of the complex task to optimize design parameters, trajectory parameters and motion control programs for the guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories is given. The complex task belongs to a problem of the optimal control theory with limitations in form of equa lity, inequality and differential constraints. To simplify the problem, an approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as a material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of the criterion function that was used for determination of the optimal parameters, programs and characteristics. The mathematical model of the guided missile provides adequate accuracy for design study to determine depending on the main design parameters: overall dimensions and mass characteristics of the guided missile in general and its structural comp onents and subsystems; power, thrust and consumption characteristics of the rocket motor; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by determining design and trajectory parameters, overall dimensions and mass characteristics, power and ballistic characteristics of two guided missiles with wings for advanced multiple launch rocket systems produced by the People’s Republic of China, using the limited amount of information available in the product catalog.

Annotation: The article deals with the problems of applying new approaches to formation and improvement of operation system. Turning of space hardware and services into marketable commodity requires their new qualities that determine competitiveness. The main task of presented works was approbation of new approaches to improvement of space launch systems operation quality and operation process effectiveness by the example of prospective Cyclone-4M space rocket complex. The works to form and improve its operation system were performed using the methods based on general theory of space systems operation and the pocedures based on the results of research work conducted by Yuzhnoye SDO in 2015 for analytical evaluation of launch services costs. The topicality of the article is confirmed by the results of practical application of new approaches in main directions of Cyclone-4M space rocket complex operation system improvement, which allowed increasing commercial attractibility of Yuzhnoye SDO-developed systems due to reduction of direct recurring costs and annual expenses. The article describes the course of development of operation model of a created object; based on investigation of the processes of this model, the object’s performance characteristics are detemined. The basis of the article are the organizational-and-technical decisions used herewith and the results obtained for Cyclone-4M space rocket complex. The article is of practical interest for specialists involved in creation of space rocket complexes and other sophisticated systems where the operation system is a multi-level organizational-technical system.

Annotation: The use of Langrangian multipliers at solution of optimal control problems in linear statement with qua dratic quality criterion leads to the necessity of solving boundary value problem with conditions for multipliers at the right end of control interval. Solution of the obtained equations for the purpose of regulation synthesis in forward time in this case does not produce stabilizing effect, as a rule. For regulation synthesis, the met hod is widely used of analytical construction of optimal regulator based on stabilizing matrix, which is obtained by solution of algebraic Riccati equation. However, in this case, there are some difficulties ‒ the necessity of calculating the stabilizing matrix, impossibility of calculating this matrix in non-stationary problem. The article proposes the regulation synthesis method by way of solving boundary value problem on regulation cycle i nterval. For this purpose, the differential equations for state parameters and Langrangian multipliers are expressed in the form of finite-difference linear relations. Taking into account that the state parameters and Langrangian multipliers are equal to zero at the end of cycle, the Langrangian multipliers at the beginning of cycle are determined by known values of state parameters for the same moment through solving the above linear system. The obtained values form the regulation law. In consequence of small duration of regulation cycle, an amplifying coefficient is introduced in the regulation law. Its value is determined based on results of preliminary modeling. Efficiency of the proposed method was verified by the example of adopted dynamic system, including non-stationary. The amplifying coefficient is fairly simply selected by the type of stabilization process. The proposed method may be used in the control systems of rockets of various purpose for motion parameters regulation.

Annotation: Contemporary trends in developing space-rocket hardware indicate the increased demand for light and ultra-light rockets. The first trend in developing the up-to-date light and ultra-light rocket hardware includes improving accuracy of cargo delivery to the specified area; the second trend covers the enhancement of energetic properties and the reduction of production and operational costs. Spinning about the longitudinal axis of symmetry may be one of the ways to improve the light and ultra-light rocket hardware in these trends. Spinning significantly increases stability of a moving object and partially evens out the negative impact of external and internal disturbing factors (skewness and eccentricities of propulsion system and control elements, wind). Refusal to use systems that provide stabilization about the longitudinal axis of symmetry leads to reduction in mass of the control system equipment, thus increasing energetic perfection of the rocket hardware. Hence, rotation of the rocket about the longitudinal axis may be caused by the spinning elements on purpose as well as by disturbing impacts in case of control failure in the roll channel. This article considers suggestions on algorithmic realization of light rocket control methods under conditions of rapid rotation about the longitudinal axis for each of the options mentioned above. This article offers control methods for the rocket, rotating about the longitudinal axis, that provide angular stabilization, improve the transient quality, and determine the angle of roll after program stop of rotation about the longitudinal axis.

Annotation: The main scientific and methodological propositions for designing single-stage guided missiles with main solid rocket motors that are intended for delivering payload to the given spatial point with required and specified kinematic motion parameters are defined. The aim of the article is to develop methodology for the early design phase to improve the basic characteristics of guided missiles, including formalization of complex problem to optimize design parameters, trajectory parameters and motion control programs for guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories. The task is defined as a problem of the optimal control theory with limitations in form of equality, inequality and differential constraints. An approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of criterion functional that was used for optimization of design parameters, control programs and basic characteristics of the guided missile. The mathematical model of the guided missile provides adequate accuracy for design study to determine: overall dimensions and mass characteristics of the guided missile in general and its structural components and subsystems; power, thrust and consumption characteristics of the main engine; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by solving design problems. Applications of the developed program were studied to present the research results in a user-friendly form.

Annotation: During 60 years of existence of specialized Liquid Rocket Engines Design Office – DO-4 as a part of Yuzhnoye Design Office, extensive experience was accumulated in development of liquid rocket engines of various purpose on storable and cryogenic propellant components. The required test benches and production base were created. When developing the engines, the DO-4 specialists widely use the experience accumulated during manufacturing and testing of the engines developed by the other design offices for Yuzhnoye SDO LVs that were manufactured by SE PA Yuzhny Machine-Building Plant and tested at Yuzhnoye SDO’s and Plant’s test benches. Along with the conventional ones, new original engine designs were developed to achieve high energy-mass characteristics, reliability and quality. Among them we should mention the RD858 and RD859 engines for the soviet lunar take-off-and –landing module of Block E, the unique RD857 and RD862 engines with afterburning of reducing generator gas and gas dynamic method of thrust vector control, the RD866 multifunctional engine of space tug ensuring multiple ignition in flight, and many others. At present, Yuzhnoye SDO jointly with SE PA Yuzhny Machine-Building Plant deliver the engine for the European Vega LV forth stage propulsion system under the contract with Avio company (Italy). Based on Yuzhnoye SDO–created engines, propulsions systems for ballistic missiles and space rockets that are unique by their characteristics and scope of functions, the engines, propulsions systems for spacecraft, LV upper stages and transfer orbit stages can be developed in short terms and at minimal costs.

Annotation: The article presents the main results obtained when solving the standardization tasks in Yuzhnoye SDO practice. The specific ways are presented of reducing the periods of work performance and increasing the accuracy of results due to the use of automated methods. The article also presents the recommendations in respect of sequence and methods of creating the standard data arrays, which allows optimizing the work process performance.

Annotation: The paper deals with the options of solving the task of constructing an inertial navigation system in the conditions of considerable g-load and angular velocity in identified direction by method of setting the sensitive elements at some angle to the identified direction, which allows making measurements in it without loss of measurement quality in the other directions. The paper describes the technique of calculating the angle of sensitive elements setting to the identified direction. The scheme of constructing an inertial navigation system with incomplete set of sensitive elements is considered for the cases when in entire operation leg, rotation around the identified direction is executed. The analysis is given of measurement vector error due to incompleteness of the sensitive elements set.

Annotation: The paper deals with the issues that arise when solving the tasks connected with the possibility of suppressing the onboard radio electronic means of small-size unmanned aerial vehicles by focused powerful offfrequency emissions of microwave band electromagnetic fields.

Annotation: Some main results of scientific support of development of launch vehicle head module composite loadbearing bays are presented. The methodology is proposed for developing these units. By the example of payload fairing and interstage bay of Cyclone-4 launch vehicle, high efficiency is shown of proposed methodology implementation when selecting their rational design and technological parameters.