22 results about "Delta" patented technology

An example method for Global Navigation Satellite System (GNSS)-based positioning performed by a GNSS device, the method may include receiving, from at least one satellite, a plurality of signals across a series of consecutive epochs and determining delta-ionosphere errors for the series of consecutive epochs, wherein each delta-ionosphere error indicates a change in ionospheric delay in carrier phase measurements taken on at consecutive epochs. The method may also include determining an ionosphere delay correction based on accumulating the delta-ionosphere errors and obtaining a position of the GNSS device based on the determined ionosphere delay correction.

The invention relates to the technical field of flow guide assemblies, and discloses a stop valve flow guide assembly which comprises a valve seat and a flow guide ring, the flow guide ring is arranged in a valve seat water inlet pipeline and composed of a fixed ring body and a movable ring body, the fixed ring body is fixed to the inner wall of the valve seat water inlet pipeline, and the movable ring body is arranged on the fixed ring body and can freely rotate. Delta wing plates are arranged on the inner side wall face of the movable ring in a circumferential array mode, the plane of each delta wing plate is in a right triangle shape, the joint of each delta wing plate and the movable ring is in large-radius fillet transition, the front edge of each delta wing plate is sharp, and the rear edge of each delta wing plate is thin and sharp.

The invention discloses a circumferential contour prediction method based on cutting force and deformation field space-time correlation, and relates to the field of aerospace, and the method comprises the following steps: 1, carrying out the spatial discretization of a circumferential contour region of a target workpiece; step 2, establishing a lightweight structure stiffness matrix; 3, constructing an elastic mechanism model and solving delta-F mapping; 4, introducing the time sequence correlation of the cutting load into the elastic mechanism model; step 5, constructing a Boolean relationship between the tool path and the workpiece envelope; step 6, fusing actual measurement process signals, and performing state correction; step 7, iterating the elastic model again by using the corrected load; and step 8, reconstructing full-field deformation based on spatial correlation. According to the method, online prediction of the station-level circumferential deformation field is achieved, the method can be used in the multi-procedure machining process, deformation in-situ compensation of the procedure is achieved, a clamping positioning strategy of the next procedure is output, and the manufacturing quality and the production efficiency are improved.

Systems, methods, devices, and means described herein relate to partial or delta updated flight path reporting. A wireless transmit / receive unit (WTRU) may transmit a first message indicating that flight path information is available. The WTRU may receive a second message including a flight path reporting configuration. The flight path reporting configuration may request the flight path information. The WTRU may transmit an initial flight path message in a third message. Based on information associated with a change from a previous flight path to an updated flight path, the WTRU may detect that the previous flight path needs to be updated with the updated flight path. The WTRU may transmit an indication that an updated flight path is available. The WTRU may receive a request for partial or delta updated flight path information and then transmit the partial or delta updated flight path information.

Some embodiments of the invention relate to methods carried out by a navigation satellite system (NSS) receiver for estimating parameters useful to determine a position. The NSS receiver observes NSS signals from a plurality of NSS satellites. The method comprises performing a delta-carrier-phase-observables-computation procedure. It is further determined whether a criterion, indicative of (i) continuity of carrier phase measurements from an anchor epoch to a terminus epoch, and / or (ii) stability of NSS satellite measurement geometry from the anchor epoch to the terminus epoch, is satisfied, and, if not, the anchor epoch is moved forward. Then, after determining whether the criterion is satisfied and after either moving the anchor epoch or not, the delta-carrier-phase-observables-computation procedure is performed again for a new terminus epoch. Systems and vehicles using such a method are also disclosed.

The invention relates to an orbit-protecting and resistance-reducing ultra-low orbit satellite and a control method thereof. The orbit-protecting and resistance-reducing ultra-low orbit satellite comprises a fuselage, a folding wing system, a lifting empennage system, a head distributed electric propulsion system, a tail vector electric propulsion system and the like. The fuselage adopts a conical head, and the wings adopt delta wing structures, so that resistance is reduced, and lift force is improved. During launching, the wings can be folded, and the empennage can be retracted into the fuselage, so that the required space is reduced. An obtuse angle is formed between the propulsive force acting direction of the head distributed electric propulsion system and the advancing direction of the satellite, and the tail vector electric propulsion system can do yawing and pitching composite motion. Compared with a traditional satellite, the lift-drag ratio is remarkably improved, and the orbit attenuation speed is greatly reduced; a composite control method of a pair of horizontal control surfaces, a vertical control surface, head distributed electric propulsion and tail vector electric propulsion is adopted, and sufficient control force and control torque are provided for attitude and orbit adjustment, so that the problem of difficulty in ultralow orbit dynamics control is solved.

The application provides a spacecraft attitude jet control method and system based on a delta-sigma modulator, which comprises the following steps: collecting the attitude angle and the attitude angular rate of a spacecraft, and feeding back the attitude angle and the attitude angular rate to an outer loop and an inner loop, respectively; in the outer loop, calculating an angular rate command according to the feedback attitude angle and an attitude angle command input into the outer loop, and inputting the angular rate command into the inner loop; in the inner loop, obtaining an angular rate deviation signal according to the angular rate command and the feedback attitude angular rate; in the inner loop, converting the angular rate deviation signal into a jet pipe relative thrust command, modulating the jet pipe relative thrust command into a jet pipe switch command with adjustable duty ratio, and converting the jet pipe switch command into an electromagnetic valve on-off signal of the jet pipe; and controlling the forward jet pipe and the reverse jet pipe to perform jet actions according to the electromagnetic valve on-off signal. The application realizes continuous control of the spacecraft attitude angle and the angular rate and improves the control precision of the spacecraft attitude under the premise of avoiding an increase in the number of jet pipe jet actions.

The present technology pertains to a system for optimizing space logistics for multi-part space missions. Space logistics are optimized by selecting a combination of space vehicles and space maneuvers performed by those space vehicles that maximize the destination mass while simultaneously minimizing the initial mass, as well as minimize mission delta-v and costs. This optimization can be performed using an objective function as a weighted combination of the destination mass and the corresponding initial mass. This optimization can be subject to various constraints, including system compatibility, operational feasibility, lower and upper bounds for the variables, mass balance at each node, inflow-outflow relationship, rocket capacity constraints, and non-negativity applied to the flow.

A method to detect time errors in a global navigation satellite system (GNSS) receiver can include determining a carrier-to-noise (CNo) delta between a CNo ratio of a first type of GNSS signal and a CNo ratio of a second type of GNSS signal. A time error may be detected based on the first CNo delta and a first threshold CNo delta, the time error indicating a use of an erroneous time uncertainty (TUNC) by the GNSS receiver. An error mitigation operation may be performed based on detecting the time error. The first type of GNSS signal may be selected based on determining that a first level of susceptibility to errors associated with the first type of GNSS signal is lower than a second level of susceptibility to errors associated with the second type of GNSS signal.

A method to detect time errors in a global navigation satellite system (GNSS) receiver can include determining a carrier-to-noise (CNo) delta between a CNo ratio of a first type of GNSS signal and a CNo ratio of a second type of GNSS signal. A time error may be detected based on the first CNo delta and a first threshold CNo delta, the time error indicating a use of an erroneous time uncertainty (TUNC) by the GNSS receiver. An error mitigation operation may be performed based on detecting the time error. The first type of GNSS signal may be selected based on determining that a first level of susceptibility to errors associated with the first type of GNSS signal is lower than a second level of susceptibility to errors associated with the second type of GNSS signal.

An example method for Global Navigation Satellite System (GNSS)-based positioning performed by a GNSS device, the method may include receiving, from at least one satellite, a plurality of signals across a series of consecutive epochs and determining delta-ionosphere errors for the series of consecutive epochs, wherein each delta-ionosphere error indicates a change in ionospheric delay in carrier phase measurements taken on at consecutive epochs. The method may also include determining an ionosphere delay correction based on accumulating the delta-ionosphere errors and obtaining a position of the GNSS device based on the determined ionosphere delay correction.

The utility model belongs to the technical field of condenser leakage detection, and discloses a detection probe for condenser leakage detection, which comprises a delta wing main body, a vacuum pumping mechanism, a driving mechanism and a slide rail, the slide rail is arranged in the condenser to be detected and is arranged along a preset leak detection route of the condenser to be detected; the delta wing main body is movably arranged on the sliding rail through the driving mechanism, and the vacuum pumping mechanism is arranged below the delta wing main body; wherein an outlet of the vacuum pumping mechanism is used for being connected with a helium mass spectrometer leak detector so as to pump water vapor at a preset sampling position in the to-be-tested condenser to the helium mass spectrometer leak detector; the driving mechanism is used for driving the delta wing main body to move along the sliding rail; by introducing the delta wing main body, the resistance of the detection probe can be obviously reduced, the detection probe is ensured to move more smoothly in a medium, the leak detection efficiency is improved, and the accuracy of a leak detection result is effectively ensured.

The invention discloses a method for determining an advanced turning parameter of a path for executing required navigation performance by a transport plane flight management system, and the method comprises the steps: determining an index requirement of an FMS executing an RNP path in a system demand confirmation process by an FMS supplier, and designing and calculating a t phi function in a system design process according to a turning rate; the function is composed of a scheduling frequency of the rolling instruction calculation task, a gain ks of time overhead of the rolling instruction calculation task, a time gain k [delta] psi of a task calculation time-consuming turning rate of the rolling instruction calculation task, and a rolling angle variable quantity [delta] phi for tracking a nonlinear path; and taking the k delta psi as a configurable parameter, and if the FMS executes turning guidance of the RNP path tracking nonlinear path and enters a steady state after leg switching and does not meet the index requirement, adjusting the k delta psi. The problems that a traditional method is high in cost, the iteration process is low in efficiency, and under some conditions, results are unstable or poor due to the fact that experiment conditions are non-repeatable are solved.

The application discloses a triangular wing unmanned aerial vehicle box type double-launching rocket boost launching system, which comprises a triangular wing unmanned aerial vehicle, a first boost unit, a second boost unit, a first blocking piece, a second blocking piece and a launching box; the first boost unit and the second boost unit are symmetrically arranged on two wings of the triangular wing unmanned aerial vehicle and each comprises a containing cabin, an ignition cable, an ignition socket, a first ball head spring pin, a second ball head spring pin, a push type gas operating cylinder, a first fixed ear, a second fixed ear, a booster sleeve, a flange, a flexible rope and a rocket booster. The application provides a rocket booster mounting and launching scheme which is compact in structure and highly integrated with the aerodynamic shape of a carrier aircraft and has the characteristics of simple operation and high preparation efficiency.

The invention relates to a satellite launching process key event image real-time acquisition method, and belongs to the technical field of satellite launching process monitoring. Calculating a shooting duration a and a time starting point range of separation of a carrying fairing, a shooting duration b and a time starting point range of separation of a satellite and a rocket, and a shooting duration c and a time starting point range of unfolding of a solar wing; setting an instruction uploading time error as T; performing precision processing on the ranges of the duration a and the duration b; the cover throwing shooting countdown time A, the satellite-rocket separation shooting countdown time B and the solar wing unfolding shooting countdown time C are determined; the whole satellite launching process image real-time acquisition process design is completed; the method realizes real-time acquisition of images in the whole process of key events including carrier rocket fairing separation, satellite-rocket separation, satellite solar wing unfolding and the like, and is particularly suitable for monitoring the launching process of a satellite equipped with a satellite-borne monitoring small camera.

This invention relates to the field of unmanned aerial vehicle (UAV) control technology, proposing a landing method and system for delta-wing UAVs to cope with ground effect nonlinear interference. The method includes: using acquired flight status and sensor data, and based on a spatial gradient dynamic residual verification mechanism, determining whether the UAV has entered a strong ground effect zone; if it has, cutting off the altitude closed-loop feedback loop and smoothly switching to the attitude impedance holding mode; in the attitude impedance holding mode, constructing a second-order impedance dynamic model of the UAV's pitch axis, generating compensating rudder deflection commands to absorb ground effect disturbance torque; during deceleration and landing, running a total energy control system algorithm, monitoring the specific total energy decay rate in real time and predicting the takeoff distance, and performing go-around redundancy protection determination; monitoring the main landing gear load sensor in real time, and triggering a feedforward control strategy to actively unload residual aerodynamic lift and complete landing at the moment of confirmed ground contact. This invention solves the problems of high-frequency oscillations in flight control and extended drift distance caused by ground effect.

The application discloses a six-degree-of-freedom flexible adaptive mechanism with a recovery reset function, which comprises a delta parallel structure, the delta parallel structure comprising a moving platform and a static platform, three connecting rod groups being hingedly connected between the moving platform and the static platform, three sliders being slidably arranged on the static platform, the three sliders being arranged in one-to-one correspondence with the three connecting rod groups, one end of each connecting rod group being hingedly connected to the moving platform, the other end of each connecting rod group being hingedly connected to a corresponding slider, a support plate being connected to the moving platform through a locking mechanism, a conical part being fixedly arranged on the support plate, an opening being arranged on the moving platform, an adapter plate being connected to the support plate through a spherical hinge, and an elastic support part being fixedly arranged on the support plate. The six-degree-of-freedom flexible adaptive mechanism with the recovery reset function can realize flexible adaptation to six-direction following errors of a docking device, can realize recovery of a connector outside a safe distance of a rocket, and has the characteristics of simple operation, safe use and strong universality.

In one embodiment, the method includes determining an initial cost volume associated with a plurality of potential trajectories of a vehicle within an environment based on a set of travel constraints for the vehicle; generating a delta cost volume using the initial cost volume and environmental data associated with the environment, the delta cost volume being generated by determining an adjustment to the initial cost volume that incorporates observed driving behavior; and scoring one of the plurality of potential trajectories for the vehicle based on the initial cost volume and the delta cost volume.

The invention discloses a nonsingular predetermined time aircraft incremental attitude control method, which comprises the following steps: establishing an attitude dynamics model, and converting the attitude dynamics model into an angular rate equation in a strict feedback nonlinear form; performing first-order Taylor expansion on the angular rate equation to obtain an incremental kinetic equation in a nonlinear form; designing a preset time attitude angle controller and an attitude angle virtual controller of the aircraft; designing a predetermined time incremental angular rate controller of the aircraft, and designing an angular rate virtual controller in a non-singular predefined time convergence form; and transforming the controller by using the output of the filter to obtain a singular-point-free predetermined time incremental angular rate controller and an angular rate virtual controller. The controller is designed to be non-singular, the stability of a closed-loop control system is guaranteed, and the robustness and the anti-jamming capability of the system can be effectively improved through the incremental controller; the convergence time of the attitude angle controller can be customized and designed by a user, and the convergence time is not influenced by the initial state of the system.