8 results about "Rocket engine nozzle" patented technology

The application provides a solid rocket engine nozzle thrust contrast test device, which comprises a front head, a combustion chamber, a rear head, a nozzle A, a nozzle B, an upper clamp, a lower clamp, a six-component force measuring platform, a hexagonal bolt, a hexagonal nut, an internal hexagonal cylindrical head screw and a stud. The two nozzles A and B with different aerodynamic profiles are fixed on the end face of the rear head through the internal hexagonal cylindrical head screw; the combustion chamber is fixed by limiting with the upper clamp, the lower clamp, the hexagonal bolt and the hexagonal nut; the base of the lower clamp is fixed on the working plane of the six-component force measuring platform through the hexagonal bolt, the hexagonal nut and the stud; during the test, the high-temperature and high-pressure gas generated by the combustion chamber is simultaneously sprayed out from the nozzles A and B, and the generated thrust is transmitted to the six-component force measuring platform through the upper clamp, the lower clamp and the connecting piece. The application can realize the accurate test and contrast of the thrust performance of the two nozzles with different aerodynamic profiles under the same engine working condition.

The present application relates to the field of rocket engine nozzle testing, and particularly relates to a flexible rocket engine nozzle cold swing test simulation test system and a test method thereof, the system comprising a pressure vessel, a servo mechanism, an inclinometer, a linear displacement sensor and a real-time simulator, the pressure vessel provides a cold pressurized environment, the servo mechanism drives the nozzle to swing, the inclinometer calibrates the swing angle, and the linear displacement sensor detects the swing and outputs an analog signal. During testing, the swing angle-voltage relationship is calibrated through the inclinometer and the linear displacement sensor; then the real-time simulator sends a sine instruction to drive the nozzle to swing, and synchronously collects the displacement signal, analyzes the dynamic characteristics of the servo mechanism and the coordination with the nozzle. Through the swing angle-voltage equation obtained by calibrating the swing, combined with continuous attitude simulation controlled by the simulator, the nozzle swing angle data is calculated, and the dynamic response of the servo mechanism under the pressurized condition is evaluated, the present application can truly reflect the continuous swing working condition of the nozzle, and improve the reliability and authenticity of the test data.

The application discloses a method for improving hydrogen embrittlement resistance of iron-based superalloy by grain boundary regulation and application thereof. The application uses iron-based superalloy powder containing trace elements as raw material, and regulates hydrogen embrittlement resistance of the iron-based superalloy by laser powder bed melting process combined with vacuum sealing tube treatment and five-step heat treatment process. The first three steps are high-temperature heat treatment, and the last two steps are low-temperature heat treatment. The hydrogen embrittlement resistant iron-based superalloy obtained by the process has high strength and plasticity matching at room temperature and 600 DEG C, and has small elongation loss after hydrogen charging, and the hydrogen embrittlement resistance is significantly better than that of Inconel 718 nickel-based alloy. The process regulates the grain boundary type and element distribution in the alloy through multi-stage synergistic regulation, not only significantly improves the comprehensive mechanical properties of the material, but also endows the material with excellent hydrogen embrittlement resistance, can effectively meet the strict requirements of liquid hydrogen rocket engine nozzle and key parts of aviation system on high reliability, long service life and extreme environment adaptability, and has wide application prospect.

The application provides a rocket engine tail nozzle closure, the pressure bearing surface of the closure is a spherical surface with an inner convex and outer concave shape, and reinforcing ribs are uniformly distributed in a radial manner on the concave surface; the thickness of the pressure bearing surface changes in a gradient manner, the thinnest part is in the center of the circle, and the thickness gradually increases towards the edge; the closure is integrally formed by injection molding; the injection molding material is short carbon fiber reinforced polyether ether ketone, which contains 10-30% short carbon fibers in terms of mass fraction. The application can take into account both low pressure sealing and high pressure reliable rupture: under the action of an internal and external pressure difference of 0.2±0.02 MPa, the adhesive interface is not detached, the closure is not ruptured or excessively deformed, and excellent air and liquid tightness is exhibited, while reliable rupture can occur in a pressure range of 0.20-1.5 MPa, and the number of fragments generated is 8-10 pieces and the size is small.

The application discloses a kind of liquid rocket engine nozzle brazing defect laser shearing speckle interferometry detection methods, suitable for liquid rocket engine nozzle brazing surface microdeformation and internal defect detection. Controllable internal pressure is applied to nozzle inner cavity by water pressure or air pressure active loading mode, which is different from traditional thermal loading, vacuum loading, vibration loading. By uniform speed pressurization to preset detection pressure value and stabilize, reference phase shift diagram is collected by shearing speckle interferometry optical path, and reference phase of reference state is obtained by multi-step TPS method reconstruction. After stopping loading, target phase of target state is obtained by the same method in the reconstruction recovery process. Full-field phase difference is calculated by phase reconstruction device and phase difference image is generated. The abnormal stripe area presented in the image is the nozzle brazing defect position. The application can stimulate microdeformation of defect area by internal pressure loading, which can adapt to complex structure of nozzle, improve detection sensitivity and eliminate background interference.

This invention provides a fixing device for rocket engine nozzles, including a ring frame. Several connecting arms are evenly distributed around the center of the ring frame at its lower end. A support arm is provided at the upper end of the ring frame corresponding to each connecting arm. Each connecting arm includes an integrally formed docking portion and a connecting portion. The docking portion is vertically arranged, and the connecting portion extends outward from the ring frame, with a connecting corner piece installed at its free end. The docking portion is fixed to the ring frame by a connecting assembly. The support arm includes a fixing frame, with an arc-shaped support plate at one end extending outward from the ring frame. The support plate is inclined, gradually tilting towards the center of the ring frame from bottom to top, and has a heat insulation structure on its outer side. The ring frame includes a paper honeycomb core material, with a carbon fiber composite panel covering the outer wall of the paper honeycomb core material. This invention effectively achieves lightweight and high specific stiffness, integrating structural load-bearing and thermal protection, avoiding the weight increase caused by external heat insulation devices.

The present application relates to the technical field of solid rocket engine nozzle, in particular to a long-time small-thrust nozzle diffuser section. It comprises a nozzle shell, a back wall heat insulation layer, a front heat sink ring, a middle heat sink ring, a back heat sink ring, a baffle ring, a back heat insulation layer, a diffuser section and a throat liner. The back wall heat insulation layer is located inside the nozzle shell. The front heat sink ring, the middle heat sink ring and the back heat insulation layer are located adjacent to the back wall heat insulation layer. The throat liner is located inside the front heat sink ring and the middle heat sink ring. The diffuser section is located adjacent to the middle heat sink ring, the back heat sink ring and the throat liner along the flow direction of the gas. The back heat sink ring is located adjacent to the back heat insulation layer and the diffuser section. The baffle ring is located at the tail end of the nozzle. The present application is suitable for solid rocket engine with working time over 100 seconds, improves the specific impulse of the solid rocket engine, improves the ablation condition of the diffuser section and improves the overall reliability of the solid rocket engine.

This application discloses a threaded foam plug and its molding method, belonging to the technical field of solid rocket engine nozzle plugs. Installed at the tail end of a rocket engine nozzle, it includes: a plug body with a first trapezoidal thread structure on its outer side; and a retaining ring with a second trapezoidal thread structure on its inner side. The first and second trapezoidal thread structures are mutually adapted and screwed together, and the mating surfaces of the first and second trapezoidal thread structures are configured to generate mechanical self-locking and radial clamping force during screwing and tightening, thereby achieving a sealed connection between the plug body and the retaining ring, thus sealing the rocket engine nozzle. Both the plug body and the retaining ring are made of expandable polystyrene foam and molded in a single step. This invention effectively resists vibration and loosening, ensuring a durable connection and structural stability under operating conditions.