Annotation: The use of servo drives on flying vehicles determines the requirements to their dynamic characteristics. The problems of dynamics of drive with jet motor are not practically covered in publications. The task arises of selection of structure and parameters of devices consisting of several subsystems whose dynamic characteristics must be brought into agreement with each other in optimal way. The purpose of this work is to develop mathematical dependences for calculation of dynamic characteristics. The functional arrangement of the drive is considered consisting of jet motor based on Segner wheel with de Laval nozzle, mechanical transmission, pneumatic distributing device – jet pipe controlled by electromechanical converter. The layout is presented of mechanical segment of servo drive with jet motor with screw-nut transmission. The dynamic model is presented and the algebraic relations to determine natural frequencies of the drive are given. The motion equations of output rod at full composition of load are given. Using Lagrange transformation as applied to ball screw transmission, the expression for reduced mass of output element was derived. The reduced mass of load depends on the jet motor design and exerts basic influence on the drive’s natural frequencies. The evaluation is given of reduced mass change from the jet motor moment of inertia and reducer transmission coefficient. Based on the proposed algorithms, the dynamic characteristics of servo drive were constructed: transient process and amplitude-frequency characteristic. The drive has relatively low pass band, which is explained by the value of reduced mass of load.

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.

Annotation: In the pneumohydraulic systems of liquid rocket engines and propulsion systems, electromagnetic valves that allow making the pneumohydraulic systems more simple and ensuring multiple ignition of liquid rocket engines have found wide application. The Yuzhnoye-developed electromagnetic valves are designed according to two schemes – of direct and indirect action. In the direct-action electromagnetic valves, the shutting-off device opens (closes) the throat with the force developed by electric magnet. They have gained acceptance in the pneumohydraulic systems with the working medium pressure of ~8.5 MPa, they are of simple design and have high operating speed (0.001…0.05 s). In the electromagnetic valves with amplification, the electromagnet armature is connected with control valve and the main shutting-off device moves due to the force from working medium pressure drop on it. They are used in the operating pressure range of 0.5…56 MPa, at that, the action time is 0.025…0.15 s. For the European Vega launch vehicle fourth stage main engine assembly that has pressure propellant feeding system, the electrohydraulic valve with amplification and drainage was developed. The dependence of this electrohydraulic valve high speed from the line’s output length is decreased to the maximum due to the installation of Venturi nozzle at the output connecting branch. This electrohydraulic valve is operable at the pressure below 8 MPa, the action time is 0.08…0.12 s. The present-day spacecraft gas-jet orientation and stabilization systems use as propulsion devices the electromagnetic valves with nozzles whose thrust is, as a rule, not more than 30 N and the working medium pressure is up to 24 MPa. Yuzhnoye State Design Office developed for 15B36 gas-jet system the electropneumatic valve with amplification and nozzle, which is operable at the pressure below 45 MPa, ensures the action frequency of up to 10 Hz and is capable of creating the thrust of 100 N on gaseous argon. To solve the task of decreasing the dependence of operability and high speed of electromagnetic valves with drainage and amplification on geometry of lines in which a valve is installed, the electropneumatic valve was developed that has spool elements ensuring reliable and quick action with long input lines of 0.004 m diameter. Its mass is 2…2.5 times lower than the mass of analogs. Recently, Yuzhnoye State Design Office develops the apogee RD840 LRE with 400 N thrust, for the conditions of which the direct-action electrohydraulic valve was developed and tested with the following characteristics: pressure – up to 2.15 MPa, consumed power in operation mode – less than 7.1 W, action time – not more than 0.02 s, mass – 0.19 kg. The presented electromagnetic valves by their technical and operational characteristics meet the highest world requirements and have found wide utility in liquid rocket engines and propulsion systems.

Annotation: During lift-off of integrated launch vehicles, the propulsion system jet generates acoustic field. Therewith, the loads can be created that are critical for the launching equipment, rocket body and especially for the spacecraft, which are under the fairing. To take into account the effects on these elements, it is necessary to determine the characteristics of generated acoustic field. The method was developed that allows modeling the acoustic fields during integrated launch vehicle lift-off based on determination of acoustic sources type. In particular, modeling of Cyclone-4M ILV jet acoustic radiation by bulky source was performed. This provided the possibility to calculate acoustic pressure amplitudes in ILV ambient medium and to evaluate acoustic effect on the rocket body at certain points. The method is expected to be used to investigate kR wave parameter. The modeling of integrated launch vehicle propulsion system (ILV PS) jet acoustic field as bulky radiation source was performed in the rocket flight leg where ILV ascent altitude does not exceed ~ 25 m. In this case, one should be based on the value of boundary frequency fb =150 Hz which separates two types of acoustic field: fb ˂ 150 Hz – front of acoustic wave of spherical type, fb > 150 Hz – front of acoustic wave of flat type. The algorithm and program of calculation of sound pressure levels were developed in JAVA language. The characteristics of acoustic fields sound pressure levels were calculated depending on radiation frequency taking into account environmental temperature. The maximal acoustic pressure level in 150 Hz frequency in the payload area outside the fairing – 155 dB, in the instrumentation bay area – 157 dB, in the intertank bay area – 172 dB, in the aft bay area – 182 dB. In the frequencies lower than 150 Hz, the sound pressure levels are lower. The calculation data are presented graphically.

Annotation: This paper describes the configuration of the rig, designed to fasten the solid-propellant rocket engine (hereunder referred to engine) during their testing, with engine thrust measurements by force transducers. It is determined that rig is the important link in the measurement chain, which directly impacts the thrust measurements, values of which will be used for calculation of the most important engine characteristics (for example, total burn). It was observed that when engine starts the basic mode of operation the thrust-measuring system of the rig is impacted by the dynamic loading caused by the engine thrust, which results in the occurrence of the oscillating processes in the rig, therefore it is important to know the behavior of these oscillating processes to estimate their impact on the thrust measurement during the design of the rigs and their thrustmeasuring systems. It is shown that if there is no procedure in the design practice to define the mode of oscillations in the rig it is rational to do it analyzing experimental thrust – time curves when different engines start the basic mode of operation. Number of experimental thrust (displacement of the engine/rig moving part) – time curves are provided for analysis when four different engines with different ballistic characteristic and capabilities start the basic mode of operation. Oscillating processes in the rig distort the physical behavior of the thrust, which requires additional analysis of the data following the tests, but since the oscillations subside rapidly, they will not significantly affect the calculation of the total burn. It is also observed that frequency of oscillations of the moving part of the rig, registered by the motion sensor, coincides with oscillation frequency, registered by the force transducer, moreover stable behavior of the oscillating processes, registered by the force transducer of the same engines, when they started the basic mode of operation, indicate the stable behavior of those engines and stable performance of the rig.

Annotation: The procedure and program were developed for calculation of rotor external characteristic (rotor rotation frequency as a function of load moment and wind speed) for autonomous wind power plants operating according to the scheme of power regulation by rotor withdrawal from under wind.

Annotation: The article considers the procedure for evaluation of acoustic stressing parameters at the observation point nearby the launch vehicle nose cone when passing the sectors with maximum velocity heads and close to 1 Mach numbers. And the problem is set to determine the overall sound pressure level and the corresponding levels in octave and 1/3-octave frequency bands. Procedure under consideration is based on the semi-empirical dependency of characteristics of the wideband aerodynamic noise, which occurs during the launch vehicle flight at high velocities due to the turbulent pressure fluctuations and dimensionless aerodynamic parameters of the main stream. General idea of this approach is to establish relation of the velocity heads with wall pressure fluctuations in the boundary layer, calculating shear stress (friction) on the shell surface based on relationships applicable in the boundary layer theory and engineering experience. Attempts of development of similar calculation models go back to the early efforts, dedicated to the study of the aeroacoustics of the launch vehicle in flight. Main advantages of the procedure are its simplicity and versatility since it can be used to determine the acoustic loads around the payload fairings of launch vehicles of different sizes and shapes within the wide range of flight velocities and altitudes.