43 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 provides a rocket aircraft flow field correction method based on multi-fidelity data fusion, and the method comprises the following steps: 1, collecting low-fidelity data, and carrying out the construction of a low-fidelity data set: simulating the change process of an unsteady flow field around an object in a finite time step through a Reynolds time-average simulation method or a large vortex simulation method; 2, data correlation analysis is carried out, wherein a linear relation model yH = rho (x) yL + delta (x) of a non-viscous flow field yL and a viscous flow field yH of the rocket aircraft is established; step 3, neural network architecture design: constructing a low-fidelity data approximation network NNL; 4, performing hyper-parameter learning and optimization; defining a loss function; 5, performing data acquisition and preprocessing: performing high-fidelity viscous flow field data acquisition, and preprocessing the acquired high-fidelity and low-fidelity flow field data; and step 6, model training and verification.

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 spacecraft game control method and device under incomplete information based on adaptive dynamic programming, and belongs to the technical field of spaceflight. The method comprises the following steps: establishing a spacecraft pursuit game relative motion model, and estimating an unknown strategy and spatial disturbance of a non-cooperative target through a fixed time expansion state observer; constructing an incomplete information pursuit game model based on the estimation state, designing an adaptive dynamic programming game strategy, and approaching an optimal game strategy by using a neural network; an Ada-delta method is introduced to adaptively adjust the learning rate of the neural network, so that the convergence efficiency is improved; an escape spacecraft shape model is constructed based on a super-quadric surface, and collision avoidance is ensured in combination with a potential function; a time constraint problem is solved through a fixed time compensation item, and an approximate optimal game strategy in fixed time is realized. According to the method, the problems of time constraint and strategy optimization of spacecraft game control under incomplete information are effectively solved while the safety is ensured.

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.

A UAS may use camera-captured ground feature data to iteratively estimate pose change through a Kalman filter. To avoid excessive drift, the UAS may periodically perform a reset by publishing delta pose and covariance as an edge in a global pose graph, by zeroing out the delta pose and covariance. To further improve global pose estimation, two UASs may share range information and associated pose graph information, thereby providing an additional constraint for each UAS to use in its pose graph optimization.

1. The name of the product of this design: UAV (Sunflower200s delta wing). 2. The purpose of this design product: used for transportation, logistics, surveying and inspection. 3. The design key point of this appearance design product lies in the shape. 4. The picture or photo that best illustrates the design points: Stereoscopic diagram 1. 5. The bottom surface of the product of this design is not easy to see or cannot be seen when in use, so the top view is omitted; the top view has been clearly expressed in other views, so the top view is omitted.

The present invention discloses a flight target recognition system based on a photonic neural network and a construction method. The ultrafast flight target recognition system based on a photonic neural network is designed and constructed for high-speed moving aircraft, including but not limited to airplanes, drones, helicopters, gliders, delta wings and other aircraft, as well as missiles, to achieve rapid recognition of high-speed moving targets. The present invention provides a flight target recognition system based on a photonic neural network, including an imaging and filtering module, a light field scaling and coupling module, a photonic neural network module, a mode control module, and an output light field coupling module; a construction method for a flight target recognition system, including artificial neural network design and photonic implementation of the artificial neural network; the present invention designs a photonic neural network transmission and control model based on all-optical passive components, trains for small sample learning of high-speed flying targets, builds static and dynamic target recognition systems, and realizes recognition of high-speed flying moving targets.

The present invention discloses a fault-tolerant control method for a delta-wing aircraft. In this method, by establishing a concurrent learning neural network and introducing a fault-tolerant mechanism on the basis of a basic controller, after a fault occurs in the unmanned aircraft, the aircraft can process and solve it online, so as to continue to work in a certain acceptable state. At the same time, concurrent learning can record historical data, ensure the convergence of the neural network weights without continuous excitation, and thus can perform tasks with higher requirements.

The invention discloses a safety protection design method for a self-adaptive augmentation controller of a carrier rocket, and the method comprises the steps: simplifying an error term of the self-adaptive augmentation controller, and obtaining a single-input-output AAC controller; amplitude limiting design is carried out on parameter angular rate deviation delta omega, an error term er, a carrier rocket attitude angle # imgabs0 # and a damping term ys signal of an error term loop and a damping term loop of the single-input-output AAC controller; carrying out amplitude limiting on an output channel of the output loop, and carrying out amplitude limiting on the output change rate of the AAC controller; and setting a trigger access mechanism for the output change rate of the single-input-output AAC controller. According to the safety protection design method for the self-adaptive augmentation controller of the carrier rocket, the safety protection design problem of the self-adaptive augmentation controller adopted by the carrier rocket is solved, and the engineering application effects of being beneficial, harmless, easy and convenient to design, flexible and reliable are achieved.

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.

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.

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.

A method for determining the fuel supply of an aeroengine, comprising: calculating the direct change in thrust ΔF1 = f1(ΔW) caused by the change in fuel supply ΔW under different intake air temperatures T1 of the aeroengine fa ; calculating the change in outlet temperature ΔT = f2(ΔW) caused by the change in fuel supply ΔW under different intake air temperatures T1 of the aeroengine fa ; calculating the change in intake air temperature ΔT1 = f3(ΔT) caused by the change in outlet temperature ΔT, and the indirect change in thrust ΔF2 = f4(ΔT1) caused by the change in intake air temperature ΔT1; calculating the total change in thrust ΔF = ΔF1 + ΔF2 = f5(ΔW) caused by the change in fuel supply ΔW under different intake air temperatures T1 of the aeroengine fa ; selecting the change in fuel supply ΔW corresponding to the maximum change in thrust ΔF among them as the optimal change in fuel quantity ΔW 8a ; calculating the maximum fuel supply W under different intake air temperatures T1 of the aeroengine fa ; calculating the difference between the maximum fuel supply W and the optimal change in fuel quantity ΔW under different intake air temperatures T1 of the aeroengine as the fuel supply W = W - ΔW 8a ; calculating the change in intake air temperature ΔT1 = f3(ΔT) caused by the change in outlet temperature ΔT 8a ; calculating the indirect change in thrust ΔF2 = f4(ΔT1) caused by the change in intake air temperature ΔT1; calculating the total change in thrust ΔF = ΔF1 + ΔF2 = f5(ΔW) caused by the change in fuel supply ΔW under different intake air temperatures T1 of the aeroengine fa ; selecting the change in fuel supply ΔW corresponding to the maximum change in thrust ΔF among them as the optimal change in fuel quantity ΔW fa ; calculating the maximum fuel supply W under different intake air temperatures T1 of the aeroengine fa ; calculating the difference between the maximum fuel supply W and the optimal change in fuel quantity ΔW under different intake air temperatures T1 of the aeroengine as the fuel supply W = W - ΔW fa.opt ; calculating the maximum fuel supply W under different intake air temperatures T1 of the aeroengine fa.max ; calculating the difference between the maximum fuel supply W and the optimal change in fuel quantity ΔW under different intake air temperatures T1 of the aeroengine as the fuel supply W = W - ΔW fa.max ; calculating the difference between the maximum fuel supply W and the optimal change in fuel quantity ΔW as the fuel supply W = W - ΔW fa.max ; calculating the difference between the maximum fuel supply W and the optimal change in fuel quantity ΔW as the fuel supply W = W - ΔW fa.dem = W fa.max - ΔW fa.opt .

In some implementations, a global navigation satellite system (GNSS) device may determine a plurality of position estimates of the GNSS device for a plurality of epochs spanning a period of time beginning with a first epoch and ending with a second epoch. The GNSS device may determine position change values comprising a position change value of the GNSS device for each successive epoch of the plurality of epochs based at least in part on delta carrier phase (DCP) values for the plurality of epochs. The GNSS device may perform backward smoothing on the plurality of position estimates, modifying at least a portion of the plurality of position estimates based at least in part on the position estimate of the second epoch and the position change values. The GNSS device may output information indicative of at least a portion of the modified position estimates of the GNSS device.

A system comprises a GNSS receiver in a vehicle that receives satellite signals from GNSS satellites and produces multiple satellite measurements. An INS in the vehicle communicates with the GNSS receiver and includes inertial sensors that produce inertial measurements for the vehicle. A processor, in communication with the GNSS receiver and the INS, hosts a navigation filter and a spoof detector module in communication with the navigation filter. The spoof detector module includes at least a first spoof detection monitor comprising a delta range monitor. The processor is operative to process the satellite measurements and inertial measurements in the navigation filter to produce a navigation solution. The delta range monitor processes the satellite measurements to detect whether there is a spoof event of the satellite signals received by the GNSS receiver. When a spoof event is detected, GNSS aiding of the INS is disabled and the spoof event is announced.

The invention relates to a Delta-type rocket recovery mechanism based on an empty ring buffer moving platform. The Delta-type rocket recovery mechanism is mainly used for meeting the actual requirements of heavy carrier rocket recovery. The device comprises a tower system, a force bearing column, a novel moving pair, a parallelogram connecting rod mechanism and an empty ring moving platform. Two-dimensional capturing motion in the horizontal plane and lowering motion in the vertical direction of the rocket are achieved through the space symmetric three-dimensional motion characteristic of the Delta mechanism, the requirement for rocket control precision is lowered, and capturing stability is effectively improved. By means of one degree of freedom of the air ring moving platform, buffering and energy absorption at the capturing moment of the rocket can be achieved, and the collision force between the rocket and the air ring moving platform is effectively reduced. The passive locking mechanism is embedded in the novel sliding pair, so that the dynamic and static states of driving can be quickly switched at the moment of impact of the rocket, and the transmission of impact load to the driver is blocked.

The invention relates to the technical field of power transmission line maintenance, in particular to a yoke plate, device and method for replacing an insulator adjacent to a triangular yoke plate side, and the yoke plate comprises a first yoke plate which comprises a triangular yoke plate connecting part used for being connected with the triangular yoke plate and a clamp connecting part used for being connected with an insulator replacement clamp. The clamp connecting part comprises an end clamp connecting part and a middle clamp connecting part; the triangular yoke plate connecting part comprises two triangular yoke plate connecting holes, the end clamp connecting part is provided with an end clamp connecting hole, and the middle clamp connecting part is provided with a middle clamp connecting hole. A stable anchoring point is provided for an existing insulator replacement clamp, and the safety and portability of high-altitude operation are improved.

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.

The present application relates to a method for designing a transfer orbit from a DRO orbit to a low lunar circular orbit. The method includes: establishing a mapping data table and a three-dimensional data pair of the delta-v of orbit transfer according to the preset DRO orbit size parameters and the phase angle of the DRO vehicle at the initial moment; dividing the three-dimensional data pair of the delta-v of orbit transfer into two groups according to the positive and negative conditions of the high and low angles of the delta-v of orbit transfer, obtaining the right ascension of the ascending node and the transfer flight time corresponding to the data pair with the minimum delta-v during the transfer process in each group to calculate the orbit transfer moment, and selecting the minimum value of the two orbit transfer moments as the selected orbit transfer moment; correcting the phase of the selected orbit transfer moment, and calculating the projection of the delta-v of orbit transfer according to the corrected phase; using the Newton iteration method to perform iterative solution with the selected orbit transfer moment and the projection of the delta-v of orbit transfer as the initial iteration values to obtain the transfer orbit transfer moment and the transfer delta-v of orbit transfer. Using this method can improve the design accuracy of the transfer orbit.

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.

1. Name of the product under this design: Chair (DELTA‑10430‑A). 2. Purpose of this design product: for people to rest. 3. The key point of the design of this product lies in its shape. 4. The picture or photo that best illustrates the design points: three-dimensional picture.