Method of modelling of process of gasification of liquid rocket fuel in tank of carrier-rocket and device for its implementation
SUBSTANCE: method for modelling the process of gasification of the liquid component of rocket fuel in the tank of the carrier-rocket, based on the supply of heat to the experimental model set (EMS), carrying out measurements of temperature, pressure at various points of EMS, discharge of vapour-gas mixture (VGM) through the drain line (DL), while the charging of the boost gas and the conductive supply of heat to the EMS are carried out, the amount of which is determined from the condition of equality of the partial pressures of the pressurant gas and liquid vapour in the EM and the fuel tank, and the total pressure corresponds to the beginning of the discharge of the VGM in the DL, the diameter of the DL is determined from the condition for resetting the preset excess pressure for the same time as in the real tank. Actuation pressure of a draining valve is selected first from a predetermined interval, the lower limit of which is the minimum boost pressure in the tank, and the upper one is the maximum pressure at which the strength of the design of the EMS remains, determine the range of parameters of the gasification process, under which condensate appears on the inner surface of the DL and crystallization, perform additional heat supply to the DL to prevent its freezing. It is considered the device for the realization of the method, which includes the EMS in the form of a model tank containing a tray for liquid, a temperature sensor, a pressure sensor, an inlet nozzle, DL, a drain valve, and a gas analyzer, while heating elements for liquid and DL are introduced in the EMS, an equipment for recording condensate and its crystallization is installed in the DL, and EMS and DL are made of a material similar to the material of the fuel tank of the carrier rocket under investigation.
EFFECT: detection of the conditions of condensate in the drainage line with subsequent crystallization during refuelling of the launch vehicle with cryogenic components of fuel or parking in the primed state at the start with the thermal loading of the fuel tank from the environment.
2 cl, 1 dwg
FIELD: engines and pumps.
SUBSTANCE: proposed test bench comprises nozzle case simulator and rear flange unloading device. The latter is arranged at nozzle case simulator and comprises different-diameter cylinders and two pistons with thrust engaged with the test bench load bearing floor. Cylinders and pistons are arranged one by one along the axis. Note here that both cylinders and pistons are interconnected. Smaller-diameter cylinder is engaged with nozzle case simulator. Smaller-diameter is elongated while seals are fitted at its lower cylindrical section. Smaller-diameter piston top cross-section perpendicular to its axis represents a circle with cut-outs at its edges. Note here that three points exist at arcs between said cut-outs that make the vertices of acute triangle.
EFFECT: higher test reliability.
FIELD: engines and pumps.
SUBSTANCE: rack for measurement of axial force of rocket engine traction contains a fixed frame, mobile part with engine fastening assemblies, adapter and force converters. On the adapter the bearing hollow barrel is installed, Inside which one or several hollow pistons are placed, and in each hollow piston on elastic membranes a bushing is installed. Force converters are fixed on the bushing coaxially.
EFFECT: invention allows to increase accuracy of measurement of axial force of draft at bench tests of the rocket engine of solid fuel.
2 cl, 2 dwg
FIELD: testing equipment.
SUBSTANCE: under heat and vacuum tests of thermocatalytic engines within a spacecraft a tight plug is installed onto a chamber of thermocatalytic decomposition of a working medium with a nozzle, a manifold of an interblock pipeline via an inspection neck and a process manifold are communicated with a bench vacuumising facility, a pressure and vacuum gage and a gas station, between which a valve is installed. Upon completion of the spacecraft testing stage with the open cover of the vacuum chamber, they connect circuits of the engine heater to the control unit. After installation of the vacuum chamber cover they pump the vacuum chamber, control generation of information by a control unit according to the fact of closure of pressure alarm contacts, close the valve and vacuumise the manifold of the interblock pipeline to the pressure level that is below the pressure of pressure alarm contacts opening. Commands are sent to switch on the engine valves, information generation by the control unit is monitored by the facts of power supply source connection to appropriate valves and opening of pressure alarm contacts. Commands are sent to switch on the engine heater, information generation by the control unit is monitored according to the facts of connection of the engine heater power supply source, operation of a thermocouple and engine heater is inspected using monitoring of temperature variation rate, corresponding to connection of the engine heater. The engine heater is disconnected, and the pause is maintained to cool down the engine. Vacuumising of the interblock pipeline manifold is completed, the valve is opened, and process gas is supplied from the gas station into the manifold of the interblock pipeline under pressure sufficient to close contacts of the pressure alarm. Then information generation by the control unit is monitored by the fact of closure of the pressure alarm contacts. Commands are sent to disconnect the engine valves, and telemetering information is monitored, which is formed by the control unit by the fact of disconnection of the power supply source of appropriate valves.
EFFECT: invention makes it possible to simplify a scheme of tests for thermocatalytic engines and to reduce their duration.
FIELD: weapons and ammunition.
SUBSTANCE: complex comprises an electronic computer, data transfer equipment with communications means, automation means, interface unit of operational and tactical and radar information, which is connected to the switching device of connection and two data transfer instruments with communications means. One of the instruments is designed for information exchange of radar information with the reconnaissance asset, and the other - for information exchange of operational and tactical information with superior, subordinate and interacting control bodies. The interface unit comprises a control unit which enables to handle radar information from the reconnaissance assets at that the ability to process in it the operational and tactical information coming from the superior, subordinate and interacting control bodies is maintained. For the target accepted for destruction the complex provides an opportunity to detect the proofs of shooting settings and data of flight mission to serve the shooting in real time.
EFFECT: increased efficiency of target destruction by giving rocket formations and formations of multiple artillery rocket systems of heavy calibre the properties of reconnaissance-strike complex operating on the principle of reconnoitred-struck.
FIELD: machine building.
SUBSTANCE: test rig contains movable horizontal platform with drive, drain tank with consumption pipeline, drain pipeline with continuity transmitter and flexible link. The platform is installed on the rig frame by means of the several parallel pendulum posts. On the platform the tested tank with intake device and drain pipeline with continuity transmitter are rigidly secured. In the consumption pipeline the flowmeter, shutoff valve, flowrate controller, hydraulic pump are installed. Pump inlet is connected to the drain tank by loop main with installed valve. The drain pipeline is rigidly secured on the platform, connected to tested tank and via the flexible link is connected with the consumption pipeline. The flexible link is made in form of a pipe with tight spherical hinges at ends and is located parallel to the posts. The flexible link length is equal to the posts height.
EFFECT: increased accuracy of hydraulic residues determination in the tested tank of the missile, and exclusion of the loads on the drain pipeline of the tested tank.
FIELD: machine building.
SUBSTANCE: test rig contains a drain tank, consumption pipeline with installed continuity transmitters, flowmeter, hydraulic pump, shutoff valve, and device for fuelling and drainage, to which the dosing unit for water topping is connected. The water dosing unit is set to the work volume equal to the volume of the expected hydraulic residues of tested fuel tank connected to the consumption pipeline. The top part of the drain tank is made in form of vertical tapering upward cone nozzle with taper 15°, on which the second continuity transmitter and overflow tank are installed. The rig contains loop main with shutoff valve connected in the consumption pipeline at the pump inlet, and fuelling pipeline with valve connected in the consumption pipeline at the pump output, which second end is connected to the consumption pipeline upstream the shutoff valve. Prior to the tested tank fuelling the consumption pipeline and drain tank are completely filled with water, then the hydraulic system is topped with dosed water volume equal to the expected hydraulic residues. Then testing is performed. Upon activation of both continuity transmitters in any sequence the shutoff valve is closed, moment of gas blow-by to the consumption pipeline and the drain tank complete filling are recorded. Then based on the known flow rate and time values, as well as topping volume ensured by the dosing unit the hydraulic residues are calculated.
EFFECT: increased accuracy of the hydraulic residues determination in the tested missile tank, and reduced labour intensity of tests.
2 cl, 1 dwg
FIELD: power industry.
SUBSTANCE: group of inventions can be used in physical modelling of gasification of residual liquid fuel in tanks of detachable parts (DP) of carrier rockets (CR) in low gravitation, using experimental simulation aggregates on the ground and in pilot CR launches with gasification systems. Gasification modelling method for residual liquid rocket fuel component (RFC) in DP tanks of CR is based on addition of heat carrier (HC) with given parameters to experimental aggregate, maintenance of defined interaction conditions in contact area of HC with the surface of liquid RFC to be gasified, temperature and pressure measurement in different points of experimental installation, additional HC flow rate measurement in different points of experimental installation, and calculation of total heat transferred to experimental installation volume through the whole experiment on the basis of measurements made.
EFFECT: improved reliability of experimental study results, reduced cost of experiments upon detection of unreliable measurements or equipment failure by experiment termination, improved measurement reliability.
SUBSTANCE: invention can be used in making components from carbon-carbon composite material operating in a high-temperature oxidative environment on the surface of rocket equipment. The apparatus for determining oxidation resistance of a carbon-carbon composite material, including with a protective coating, which includes a chamber made of refractory material for holding a sample of test material and a nozzle for feeding a gas stream into the chamber on the front wall of the apparatus, is provided with a set of detachable front walls of different thickness, wherein the nozzle is located at different angles to the longitudinal axis of the chamber of the apparatus, wherein the chamber of the apparatus is placed in a metal housing with a heat-protective cover, the heat-protective cover and the chamber being detachable.
EFFECT: invention enables simulation of the effect of a high-temperature gas stream on a component of rocket equipment in conditions close to real conditions, and determination of oxidation resistance of a carbon-carbon composite material with action of a high-temperature gas stream at different angles and at different distances.
FIELD: test equipment.
SUBSTANCE: invention relates to simulative testing of nozzles. Device comprises feed pipeline connected with receiver for plug-in joint in two mutually perpendicular planes with nozzle to be tested by means of detachable flange straps to allow measuring means to rest on receiver casing wherein feed pipeline has a resilient insert. Besides, said receiver has two holes, one being made at its and while another in its lateral surface. Note here that holes necks feature identical section and are provided with detachable flange straps to fix the nozzle to be tested in two mutually perpendicular positions. Note also that used instruments represent single-component force transducers secured at receiver casing Measuring bars of said instruments are arranged three mutually perpendicular directions while their ends rest on receive casing to hold it back.
EFFECT: higher efficiency and precision, lower labour input in fabrication and operation.
SUBSTANCE: invention relates to aerospace engineering and can be used in gas-hydraulic lines of liquid-fuel rocket engines. In the method of liquid-fuel rocket engine chamber geometrical axis installation in rated position based on excluding influence of process tolerances during manufacturing aggregates, parts and assembling units, as well as influence of material shrinkage in joint welds of gas lines between auxiliary pump-drive assembly and chamber heads for angular deviation of chamber geometrical axes from rated position. According to invention, measurement of compensating closing device actual parameters, its manufacturing, adjustment and welding are performed at the final stage of line assembling after executing all welds of jointed aggregates, parts and assembling units. The method is implemented by gas line compensating closing device containing equalising sub with cavities at its butt ends for backing rings installation. According to invention, in the device, cavities for backing rings installation are made of length equal to length of backing rings, and above cavities, through holes are drilled where fixtures for moving backing rings to the area of parts and assembling units joint welds; cavities in fixtures for screw-driver are turned perpendicular to chamfer projection plane; fixtures are installed along perimeter at every 120°; in holes of equalising sub and fixture heads bevels are made to exclude poor penetrations in weld roots.
EFFECT: invention provides higher accuracy of its installation and lowering thrust vector losses of engine working in flight or on stand.
5 cl, 12 dwg
FIELD: testing hybrid rocket engines.
SUBSTANCE: proposed device has housing accommodating barrier ozonizer and precipitation condenser installed in tandem along gaseous emission flow and incorporating, respectively, parallel plate electrodes of two types, that is, charging and precipitation ones, with electric field built up in plate-to-plate gaps. Charging plates are connected to ac current supply rated at voltage higher than breakdown value; they are provided with barriers made in the form of insulating coatings and discharge gap between adjacent plates for passing gas emissions and forming corona discharge. Precipitation plates are connected to dc current supply rated at voltage lower than breakdown value.
EFFECT: facilitated permeability of gas jet for oxygen, enhanced corona discharge power and aerosol particle removal capability.
3 cl, 3 dwg
FIELD: rocketry; test facilities.
SUBSTANCE: proposed test set designed for testing small-size solid propellant rocket engines contains storage reservoir and combustion products cleaning system interconnected by gas-dynamic duct. Storage reservoir is installed directly after engine and is hermetically connected with engine nozzle. Gas cooler is installed in gas-dynamic duct between storage reservoir and cleaning system.
EFFECT: no adverse action on environment at combustion of solid propellant charge in the open.
SUBSTANCE: proposed method includes fitting of two fusible wire warning units forming two independent electric circuits in solid propellant restricted rod specimen at a distance over length of specimen. Then specimen is burnt. In process of burning continuous recording of relative resistances for each circuit is carried out found from ratio of current resistance in circuit to resistance in circuit before beginning of burning. All time moments corresponding to moments of reaching of some preset empiric value equal to 1.05-1.15 of relative resistance are determined for each circuit separately. Then fixing fusing time for each warning unit is set equal to maximum time chosen from fixed moments. Then turning rate of solid propellant is found from ratio between distance at which fusible wire warning units are installed in specimen and difference in time of fusing of first and second warning units.
EFFECT: possibility of fixing fusing time of warning unit at slow change of resistance, prevention of error in determining moments of fusing at electric interferences.
FIELD: experimental determination of interballistic and power characteristics of charges for multi-regime rocket engines.
SUBSTANCE: a sleeve with a spring valve in the bottom part is joined in the bench test set to the combustion chamber, an adjusting piston is positioned in the valve neck, exit ports are made on the side surface of the neck, the necks are closed with an outer shutter kinematically linked with a drive, control panel and a pressure transducer in the combustion chamber, and a nozzle unit with an insert of a definite critical section is joined to the sleeve.
EFFECT: enhanced accuracy of determination of the charge burning rate for each regime of the engine and specific impulse of the solid propellant using a real charge in the tests.
FIELD: ground testing facilities.
SUBSTANCE: invention is designed for determining parameters of thrust vector of engines with skewed nozzle. In proposed test stand provision is made for fixing horizontal and vertical components of thrust vectors by force meters with minimum distortion owing to mounting of force meters on carriages and using pair of force meters both in horizontal and vertical directions to determine point of intersection of vector of thrust with axis of nozzle.
EFFECT: improved accuracy of measurement of engine thrust vector, namely, direction and coordinate of point of vector passing relative to nozzle axis.
FIELD: aircraft industry; aircraft engines; testing facilities.
SUBSTANCE: invention is designed for testing of aircraft vectored thrust engines on test stands with measuring of thrust components ±Rx, ±Ry, ±Rz at high accuracy and stability of measurements at direct and reverse operation of engine. According to invention, all force sensors for measuring components of thrust force ±Rx, ±Ry, ±Rz are furnished with additional leverages and are loaded before testing with designed weights Gx, Gy, Gz providing playless operation of force measuring leverages when determining values of components of thrust force -Rx, -Ry, -Rz at reversing of engine under testing and equality of gear ratios of two pairs of force receiving levers operating in parallel and taking up load from dynamometric platform, and two pairs of force measuring levers interacting with force sensors and provides equalizing of forces of connecting hinge joints from moments relative to axes X, Y, Z which prevents relative influence of forces on measured components of thrust forces ±Rx, ±Ry, ±Rz .
EFFECT: improved accuracy of measurements.
FIELD: rocketry, applicable for unbiased measurement of static pressures of gas flow in the duct of charges of the solid-propellant rocket engine.
SUBSTANCE: the method for determination of radial pressure differential in the duct charge of the solid-propellant engine by measurement of static pressures at the surface of the charge duct and in the clearance between the charge and the combustion chamber of the rocket engine is accomplished as follows. Pressure transducers are installed in the adapters located in the measuring sections of the combustion chamber, and a firing rig test of the rocket engine is performed. The pressure near the surface of the charge duct is measured through a steel intake tube having an extension piece of polymeric material. The extension piece has the shape of a truncated cone with a thermodecomposition temperature within 200 to 240C and a height equal to the thickness of the charge arch in the measuring section. The extension piece is pasted-in in the charge arch in the radial direction. One end of the intake tube is mounted in the extension piece to a depth of 0.1 to 0.5 of the extension piece height. The opposite end of the intake tube is jointed to the adapter of the pressure transducer for reciprocating motion in it. The radial pressure differential is estimated by the difference of pressures measured in the clearance and duct.
EFFECT: enhanced precision of pressure measurement in the measuring section, prevented distortion of the gas flow parameters at measurement.
3 cl, 1 dwg
FIELD: rocket manufacturing.
SUBSTANCE: method consists in producing an additional source of combustion product depletion by introducing a simulating charge with a low susceptibility of the combustion rate to pressure into the pilot engine chamber. The simulating charge features the combustion product temperature equal to that of the rocket solid-propellant charge combustion products and the exponent in the combustion rate law meeting the requirement to be protected by this invention.
EFFECT: simplifying the pressure stabilisation in the pilot engine chamber with a rocket solid-propellant charge combustion rate susceptible to pressure, and reducing the costs of the equipment required for pressure stabilisation.
FIELD: motors and pumps.
SUBSTANCE: inventions are referred to rocket engineering, namely to elimination of rocket engine solid fuel charge on a test bench equipped with localisation, cooling and neutralisation chamber for combustion products (LCNC). Determination and monitoring of various process parameters including combined steam gas flow rate measurement in combustion products is required to solve a number of problems related to charge elimination in solid fuel-operating rocket engine on test benches equipped with LCNC sections. The inventions for combined steam gas flow rate measurement in combustion products ensure reliability and validity of data to be acquired under complicated conditions of semi-closed LCNC spaces, where high-speed, high temperature steam gas flow of combustion products is remotely monitored. The initial rate of combined steam and gas flow as well as rate dynamics during the whole process is monitored and measured by means of this method and telescopic module. It allows for determining optimal process modes (especially amount of the supplied cooling agent, hydraulic parameters, changing conditions of combustion products outflow rate, etc) ensuring rocket engine charge elimination safety on test benches equipped with LCNC sections.
EFFECT: process safety of rocket engine charge elimination on test benches equipped with LCNC sections.
2 cl, 6 dwg
FIELD: engines and pumps.
SUBSTANCE: invention relates to test stands intended for testing high-output liquid propellant rocket engines. The ejector prechamber communicates with the intake line, while its nozzle is connected to the multi-chamber pump outlet. Note that the ejector outlet is coupled with the pump inlet sucking water in from the reservoir, while the suck-in pipeline intake represents a branch pipe with its cylindrical surface furnished with a number of orifices. The said branch pipe end face is stopped with a plug, while the branch pipe is arranged in the pit made at the reservoir bottom center. There is a tapered cover fitted right above the said pit. Note that there is a gap between the tapered cover face and the reservoir bottom arranged inclined to the center at an acute angle.
EFFECT: higher efficiency of intake device removing deposits from reservoir bottom using ejector without draining water from test stand.
6 cl, 5 dwg