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: To solve the navigation tasks (determination of the apparent accelerations and angular velocities and calculation of rocket orientation angles) in the rocket engineering, the data from the sensing elements (angular velocity sensors and accelerometers) is used. Accuracy of reference attitude determination of the rocket in the steady mode (at lift-off) has great influence on accuracy of the received navigation data during the flight. Gimballess inertial navigation system, built on the basis of inertial MEMS-sensors of Industry class (three angular velocity sensors and three accelerometers), is taken as the navigation device. In the classical version, the integration of data from angular velocity sensors and from accelerometers is the basis of gimballess inertial navigation system operation. It results in accumulation of errors when solving the navigation task (in particular due to the integration of data from angular velocity sensors). Taking it into consideration, the alternative method of rocket’s reference attitude determination during the pre-launch processing is offered. This method does not use mathematical operations of integration and is autonomous. Initial data, received from the gimballess inertial navigation system, is used as the output data. This data is used to determine the rocket’s reference attitude (orientation of object-centered coordinates in the geographical reference system) in the steady mode. Orientation angles are determined without the integration of data picked up from the angular velocity sensors. The comparative analysis to define the processing efficiency of the navigation device initial data was held during the determination of the rocket’s orientation angles in the steady mode, using the proposed method and Runge-Kutta method. The received results showed that accuracy of the reference attitude determination with the proposed method is higher. Thus, the proposed method will help reduce the errors in determination of the rocket’s reference attitude in the steady mode that in the future will improve the accuracy in determination of navigational parameters during the rocket’s flight.

Annotation: Pyrotechnic devices are important elements in rocket and space technology, which to a large degree determine the flight success of the launch vehicles, since they enable instantaneous operations to separate spent stages, change configurations, ensure safety, etc. Pyrotechnic devices are subject to strict requirements for reliability, safety, security and efficiency. The article presents an experimental study of the performance of a linear shaped charge of a launch vehicle stage separation system. This type of linear shaped charge is one of the most common types of linear shaped charge, which are used in launch vehicle separation systems being developed in Ukraine. One of the main characteristics of the linear shaped charge, which determines the efficiency and reliability of the separation process, is the depth of penetration of the cumulative jet into the obstacle. The work studied the effect of a cumulative jet of a linear shaped charge with a semi-cylindrical cumulative part. An experimental confirmation of the performance of this type of linear shaped charge is presented, using the example of a linear shaped charge with a diameter of 5 mm, acting on an obstacle made of aluminum alloy grade 2219. The research methodology, experimental scenario, in particular, a description of the research object and a scheme for measuring test results are presented. Depth of cumulative jet penetration into the obstacle was measured in 60 points along the cut line of the samples under study. A statistical analysis of the experimental results was carried out, in particular, the average penetration depth was determined. An improved formula is proposed for the practical calculation of the penetration depth of a cumulative jet for a linear shaped charge with a semi-cylindrical cumulative part, using an additional correction factor. It is noted that the depth of penetration of a cumulative jet into an obstacle is significantly influenced by technological aspects. Taking into account this influence, the lower limit of the one-sided tolerance interval was determined. Recommendations are provided to improve future experimental procedures. Based on the obtained results, it was established that the linear shaped charges under study are operational and meet the requirements for linear shaped charges, installed on launch vehicle separation systems.

Annotation: Separation of the spent LV stages is one of the important problems of the rocket technology, which requires the comprehensive analysis of different types of systems, evaluation of their parameters and structural layouts. Basic requirements are specified that need to be taken into account when engineering the separation system: reliable and safe separation, minimal losses in payload capability, keeping sufficient distance between the stages at the moment of the propulsion system start. Detailed classification of their types («cold», «warm», «hot», «cold-launched» separation) is given and their technical substance with advantages and drawbacks is described. Certain types of «cold» and «warm» separation of the spent stages of such rockets as Dnepr, Zenit, Antares, Falcon-9 with different operating principle are introduced – braking with the spent stage and pushing apart two stages. Brief characteristics of these systems are given, based on the gas-reactive nozzle thrust, braking with solid-propellant rocket engines, separating with spring or pneumatic pushers. Development of the separation system for the advanced Cyclone-4M ILV is taken as an example and design sequence of stage separation is suggested: determination of the necessary separation velocity and capability of the separation units, determination of the number of active units, calculation of design and energy parameters of the separation units, analysis of the obtained results followed by the selection of the separation system. Use of empirical dependences is shown, based on the great scope of experimental and theoretical activities in the process of design, functional testing and flight operation of similar systems in such rockets as Cyclone, Dnepr and Zenit. According to the comparative analysis results, pneumatic separation system to separate Cyclone-4M Stages 1 and 2 was selected as the most effective one. Its basic characteristics, composition, overall view and configuration are specified. Stated materials are of methodological nature and can be used to engineer the separation systems for LV stages, payload fairings, spacecraft etc.

Annotation: Despite stringent environmental requirements, modern launch vehicles/integrated launch vehicles (LV/ILV) burn toxic propellants such as NTO and UDMH. Typically, such propellants are used in the LV/ILV upper stages, where a small amount of propellant is contained; however, some LV/ILV still use such fuel in all sustainer propulsion stages. For launch vehicles containing toxic rocket propellants, flight accidents may result in the failed launch vehicle falling to the Earth’s surface, forming large zones of chemical damage to people (the zones may exceed blast and fire zones). This is typical for accidents occurring in the first stage flight segment, when an intact launch vehicle or its components (usually individual stages) with rocket propellants will reach the Earth’s surface. An explosion and fire following such an impact will most likely lead to a massive release of toxicant and contamination of the surface air. An accident during the flight segment of the LV/ILV first stage with toxic rocket propellants, equipped with a flight termination system that implements emergency engine shutdown in case of detection of an emergency situation, has been considered. To assess the risk of toxic damage to a person located at a certain point, it is necessary to mathematically describe the zone within which a potential impact of the failed LV/ILV will entail toxic damage to the person (the so-called zone of dangerous impact of the failed LV/ILV). The complexity of this lies in the need to take into account the characteristics of the atmosphere, primarily the wind. Using the zone of toxic damage to people during the fall of the failed launch vehicle, which is proposed to be represented by a combination of two figures: a semicircle and a half-ellipse, the corresponding zone of dangerous impact of the failed LV/ILV is constructed. Taking into account the difficulties of writing the analytical expressions for these figures during the transition to the launch coordinate system and further integration when identifying the risk, in practical calculations we propose to approximate the zone of dangerous impact of the failed LV/ILV using a polygon. This allows using a known procedure to identify risks. A generalization of the developed model for identifying the risk of toxic damage to people involves taking into account various types of critical failures that can lead to the fall of the failed LV/ILV, and blocking emergency engine shutdown during the initial flight phase. A zone dangerous for people was constructed using the proposed model for the case of the failure of the Dnepr launch vehicle, where the risks of toxic damage exceed the permissible level (10–6). The resulting danger zone significantly exceeds the danger zone caused by the damaging effect of the blast wave. Directions for further improvement of the model are shown, related to taking into account the real distribution of the toxicant in the atmosphere and a person’s exposure to a certain toxic dose.

Annotation: A promising experimental bench – a shock wind tunnel was put into operation at Yuzhnoye SDO. The shock wind tunnel is designed to simulate the incident flow during rocket flight at high supersonic and hypersonic velocities. To solve actual design problems facing Yuzhnoye SDO, it was necessary to expand the range of velocities under investigation in a shock wind tunnel by low supersonic Mach numbers (Mа=1.5; 2; 3). As a result of this work, a modernized configuration of the shock wind tunnel was developed, which allows simulating flow parameters at low supersonic velocities. The results of aerodynamic experiment performed in the modernized shock wind tunnel, which are close to full scale ones, can be obtained using as much data on peculiarities of supersonic flow formation in it as possible. Therefore, the study of the distribution of Mach numbers profiles in the working section of the modernized shock wind tunnel at low and high supersonic velocity was chosen as the main line of research. The results of the research presented in the article are based on the use of numerical simulation methods, as well as data obtained experimentally. As a result of gasdynamic simulation of a supersonic flow conducted for the nozzle Mа=4 and the nozzle Mа=2, the calculated and experimental data on the distribution pattern and field values of Mach numbers in the working section of the tunnel were obtained. A comparative analysis was carried out. The boundaries of the region of equal velocities, within which the condition of quasistatic supersonic flow is satisfied, and the lifetime of the operating mode for the selected nozzle type were determined. At the flow from the nozzle Mа=2, a peculiarity was revealed in the distribution pattern of Mach numbers fields associated with the appearance of “blocking” effect of the supersonic flow. The methods for eliminating the effect of flow “blocking” at low supersonic velocities are proposed.

Annotation: Due to an increase of power of rocket engines, the high intensity sound field generated by the exhaust jets have become an important factor, which determines the success rate of a rocket launch. Ensuring a successful launch of a rocket system became harder due to new engineering problems. Identification and definition of acoustic sources structure within a complex supersonic jet, being a one of the most important scientific problems, which have to be solved to find the ways to control accoustic radiation. A three components of acoustic sources can be defined here – broadband signals from large and small components of of turbulent jet and tonal signals which usually being overlooked during the estimation of overall sound pressure level. The paper considers various aspects of acoustics of the launch of rocket systems, which includes characteristics of acoustic sources in supersonic jets, possibilities and physical limitation factors, under which it is possible to control the sound radiation. Among the possible ways to control the process of sound generation by a jet, a method of water injection in a jet is being studied. While saving the general thrust of the engine this method can not greatly reduce the sound radiation by a jet. It is recommended to use big amounts of water-air mix to protect the launch pad from damage. Significant progress on the topic of understanding the process of sound generation by supersonic jets can be achieved via mathematical modeling of sound radiation. The latest achievements of mathematical modeling of sound generation by supersonic jets being presented.

Annotation: Being a final executive element of rocket control systems, a hydraulic actuator is at the same time the main source of various non-linear dependencies in rocket dynamic design whose availability dramatically com plicates theoretical analysis of their dynamics and control systems synthesis. The required accuracy and complexity of mathematical models of hydraulic servo mechanisms are different for different design phases of guided rockets. The paper deals with the simplest models of hydraulic servo actuators intended to calculate rocket controllability and to define requirements to response and power characteristics of the actuators. To calculate the rocket stability regions and to evaluate own stability of servo actuators, a linearized mathematical model of hydraulic servo actuator is used that takes into account the most important parameters having impact on stability of the servo actuator itself and on that of the rocket: hardness of working fluid, stiffness of elastic suspension of the actuator and control element, slope of mechanical characteristic of the actuator in the area of small control signals, which, as full mathematical model analysis showed, is conditioned only by dimensions of initial axial clearances of slide’s throats. The full mathematical model constructed based on accurate calculations of the balance of fluid flow rate through the slide’s throats allows, as early as at designing phase, determining the values of most important static and dynamic characteristics of a future hydraulic actuator, selecting optimal characteristics of slides based on specified degree of stability and response of servo actuator and conducting final modeling of rocket flight on the integrated control system test benches without using real actuators and loading stands. It is correct and universal for all phases of rockets and their control systems designing and testing. Using this mathematical model, the powerful actuators of a line of intercontinental ballistic missiles with swinging reentry vehicle and the main engines actuators of Zenit launch vehicle first stage were developed. The results of their testing separately and in rockets practically fully comply with the data of theoretical calculations.

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 measurement errors upon conducting flight tests for launch vehicles are evaluated by considering the interferences and uncertainties in the measurement system procedure. Formal use of this approach can lead to unpredictable consequences. More reliable evaluation of errors upon conducted measurements can be achieved if the measurement process is regarded as a procedure of successive activities for designing, manufacturing, and testing the measurement system and the rocket including measurements and their processing during the after-flight analysis of the received data. The sampling rates of the main controlled parameters are three to ten times higher than the frequency range of their changing. Therefore, it is possible to determine the characteristics of the random error components directly on the basis of registered data. The unrevealed systematic components create the basic uncertainty in the evaluation of the examined parameter’s total measurement error. To evaluate the precision and measurement accuracy of a particular launch, the article suggests specifying the preliminary data on measurement error components determined during prelaunch processing and launch. Basic structures of algorithms for evaluation of precision and measurement accuracy for certain mathematical models that form the measured parameters were considered along with the practical case when static correlation existed among the measured parameters.