96 results about "Orbiter" patented technology

A rotating stream sprinkler of the type having a rotatable deflector with spiral vanes engaged by one or more water jets for imparting a rotary drive torque to the deflector, which converts the jets into a plurality of relatively small outwardly projected water streams swept over the surrounding terrain to irrigate adjacent vegetation. The sprinkler includes a speed governor having a turbine driven orbiter with first and second gears meshed respectively with a stator gear and a driven gear having different numbers of gear teeth, wherein the driven gear is carried by the deflector for rotation therewith. The turbine rotatably drives the orbiter on an eccentric axis, to react against the stator gear and thereby rotatably drive the driven gear and deflector with a substantial speed reduction. The speed governor thus regulates deflector speed for slowly sweeping the projected water streams over the adjacent landscape.

The invention discloses a reentry flight attack angle guiding method of a sub-orbital vehicle. The method comprises the following steps of: acquiring an attack angle design predicted value through analogue simulation at time intervals; finding the attack angle design predicted value, which keeps a normal overload value of the vehicle always in a fluctuation region of expected normal overload dynamic balance, by using a vehicle homomorphic prediction model in each time interval, so as to realize normal overload dynamic balance at the time interval; acquiring a practical reentry flight attack angle design parameter of the vehicle through the attack angle design predicted value by a fitting method, and determining a reentry flight attack angle design value of the vehicle; and introducing resistance acceleration integration ratio correction at each flight time to make a practical reentry flight resistance acceleration integral value tend to be the same as a predicted resistance acceleration integral value under the attack angle design value, so the reentry flight normal overload of the sub-orbital vehicle is kept to fluctuate in a predetermined region in a dynamic balance section consisting of each time interval and then the aim of reducing the reentry flight normal overload peak value of the sub-orbital vehicle is fulfilled.

The invention relates to an autonomous navigation method for a planet in a final approaching section, which belongs to the field of deep space exploration. The method comprises the following steps: establishing a state model of the planet in the final approaching section; carrying out radio measurement and communication through a detector and planetary orbiters determined by n positions, observing m pulsars and establishing an autonomous navigation measurement model of the planet in the final approaching section; and calculating the state of the detector through navigation filtering based on the two models. According to the invention, measurement characteristics of pulsar navigation and radio navigation are combined in the autonomous navigation method, a non-linear filtering method is employed, so autonomous navigation of the planet in the final approaching section is realized, precision and instantaneity of autonomous navigation of the planet in the final approaching section are improved, and the autonomous navigation method has good exploitativeness and high measurement precision and can satisfy the requirement for instantaneity of autonomous navigation.

The invention discloses a Mars final approaching section autonomous navigation method based on relative measurement information, and belongs to the technical field of deep space exploration. For achieving Mars final approaching section autonomous navigation, a final approaching and surrounding dynamic model is established. According to differential information of X-ray pulsar arrival time and Doppler velocity measurement information of a Mars orbiter, based on nonlinear filtering algorithm, the states of an approaching detector and the Mars orbiter undergo joint estimation, absolute navigation is achieved through utilization of two kinds of relative measurement information including the differential information of X-ray pulsar arrival time and the Doppler velocity measurement information of the Mars orbiter, and the accuracy of Mars final approaching section autonomous navigation and the accuracy of entry point state estimation are improved. The method has the following advantages: introduction of a planet ephemeris error is avoided, the adverse effect on navigation performance due to nondeterminacy of pulsar parameters is limited, the problem of state divergence resulting from long observation time of pulsars is overcome, the navigation performance of two detectors is improved at the same time, and the navigation filter accuracy and the convergence speed are increased.

The invention discloses a method for obtaining a design attack angle for reentry flight of a suborbital vehicle. The method comprises the following steps: dividing suborbital reentry flight into a plurality of time periods, and establishing and utilizing a homomorphic prediction model for suborbital reentry flight so as to obtain time values of the time periods; by virtue of iteration of the homomorphic prediction model, obtaining design attack angle values which enables the normal acceleration for vehicle reentry to be maintained in a preset fluctuation range within the time periods; and maintaining the normal acceleration for reentry flight of the suborbital vehicle to fluctuate in the preset range in a dynamic balancing segment composed of the time periods, thus achieving the purpose of lowering the normal acceleration peak value for suborbital reentry flight.

A reusable module is affixed atop a reusable orbital vehicle (OV). Various configurations of the reusable module have identical external dimensions in the region of attachment to the OV to permit interchangeability. Different configurations can accommodate a variety of missions of different type and duration. A variety of cargo modules of different configurations allow cargo to be uplifted into orbit. In one embodiment, the cargo module is an unpressurized cargo module in which the cargo is exposed to the environment of space during the unloading process. The cargo module may also be a pressurized cargo module. In an alternative embodiment, the cargo module may include both a pressurized cargo module and unpressurized cargo module.

The invention relates to a layout optimization method of a planet navigation orbiter based on a prediction track, belonging to the technical field of deep space detection. The layout optimization method provided by the invention has the advantages that a dynamic moving track of a navigation orbiter and a detector in an inlet section in a process that the initial layout of the planet navigation orbiter is optimized is considered, the observability of a Fisher information matrix derivation navigation system is utilized for representing a navigation property, the initial layout of the planet navigation orbiter is optimized based on the integral of the observability to a navigation system, the maximization of the observability of the navigation system is realized, the estimated accuracy of the state when a detector enters is improved, and the optimum of a navigation property is guaranteed.

A gyroscopic orbiter with vertical takeoff and vertical landing capabilities can transition between different functional modes while in-flight. The orbiter typically includes a fuselage, a front boom, a front propulsion unit, a rear boom, and a rear propulsion unit. The front boom is mounted at two pivot points to a bow of the fuselage by the front boom. The rear boom is mounted at two pivot points to a stern of the fuselage by the rear boom. One functional mode is the vertical takeoff and landing mode, wherein the propulsion units are oriented parallel to each other and are directed upward. Another functional mode is the shuttle mode, wherein the propulsion units are oriented at an angle with each other, and the front propulsion unit is directed forward. Another functional mode is the high speed mode, wherein the propulsion units are oriented collinear with a roll axis of the fuselage.

The transmission includes an orbital gear complex in combination with a variable hydraulic pump and motor. The input to the transmission is increased in speed by the orbital gearing such that, when the pump and motor are not operating, the orbiter is stationary, and the orbital gearing produces an overdrive condition. A gear reduction is accomplished by rotating the web with the web-rotating device, providing a high gear reduction. The pump and motor are preferably long-piston hydraulic machines with infinitely variable swash plates. The hydraulic machines preferably have wobblers stabilized by full gimbals and hold-down plates with elongated holes for the long pistons to eliminate possible impacts between the hold-down plates and the head ends of the long pistons when the swash-plates are at or near their maximum angle of inclination.

The method for lightening the weight of fuel stowed onboard during an interplanetary mission is characterized in that it consists:

• • in launching (10a) a first orbiter spacecraft (1) from the Earth on a first interplanetary trajectory (31, 33) towards a target planet to be explored,
  • in launching (10b) a second orbiter spacecraft (2) from the Earth on a second interplanetary trajectory (32, 34) towards a rendezvous spot (38), the second interplanetary trajectory not comprising any phase of placing in orbit around the target planet,
  • in recovering a load to be transported and in loading it onto the first orbiter spacecraft (1),
  • in returning the first orbiter spacecraft (1) and the load from the target planet to the rendezvous spot (38),
  • in effecting a docking (14) of the two orbiter spacecraft (1, 2),
  • in returning at least the second orbiter spacecraft (2) and the load, from the rendezvous spot (38) to a terrestrial orbit (41).

The invention discloses a cross adjustment method for reentry flight process of a suborbital vehicle. The method comprises the following steps: dividing suborbital reentry flight into a plurality of time periods, and establishing and utilizing a homomorphic prediction model for suborbital reentry flight so as to obtain time values of the time periods; by virtue of iteration of the homomorphic prediction model, obtaining design attack angle values which enables normal acceleration for vehicle reentry to fluctuate near a normal acceleration constraint value within the time periods; maintaining the normal acceleration for reentry flight of the suborbital vehicle to fluctuate near the normal acceleration constraint value in a dynamic balancing segment composed of the time periods so that the suborbital vehicle reaches the maximum velocity attenuation under normal acceleration restriction within the time periods, and obtaining the earliest cross adjustment time; and adding a velocity heeling angle after the earliest cross adjustment time so as to achieve the purpose of improving the capability for cross adjustment of suborbital reentry flight.

The invention discloses an effective load space cabin structure applicable to an orbiter, comprising a cabin body and cabin doors; the cabin doors are connected to the cabin body; the cabin doors can open and close the cabin body; the cabin doors are two symmetrical opposite-opening structures, and every cabin door comprises an inner skin, an outer skin, main force bearing beams, honeycomb cores, a reinforcing rib and a heat protecting structure; the honeycomb cores are located between the inner skin and the outer skin; a guide flow slot is arranged on a contact position of the honeycomb cores and the outer skin, the reinforcing rib is arranged between the honeycomb cores, and connected to the main force bearing beams; three cabin door locks are evenly distributed on the main force bearing beams, the two cabin doors are locked to each other in the process of closing; the thermal protecting structure is arranged on the outer skin, and arranged in a staggering manner; the vapor return absorption of the surface of the thermal protecting structure under the natural environment is not more than 0.15%, thus the effective load cabin door structure is high in structural rigidity, good in structure compatibility and sealing property; meanwhile, the effective load space cabin structure can meet integration and lightweight requirements at the same time.

The invention relates to a super-high-dynamic spread spectrum signal fast-capturing method suitable for a reusable orbiter. The method comprises the steps that fast-capturing parameters of spread spectrum signals are calculated at first, then multiple channels of pseudo codes based on masks are generated, and fast capturing of the spread spectrum signals is achieved by conducting FFT analysis by turns at last. The specific steps are that a carrier frequency is divided into subsections, and signal capturing is conducted in each frequency subsection; multiple channels of parallel pseudo codes are generated in a mask mode, partial correlation operation is conducted on each channel of pseudo codes and input signals, and FFT conversion is conducted on multi-channel parallel correlation operation results by turns to save resources; after all frequency bands are traversed, precise values of carrier Doppler and pseudo code phases are obtained. The super-high-dynamic spread spectrum signal fast-capturing method suitable for the reusable orbiter effectively solves the spread spectrum signal fast-capturing problem of the reusable orbiter, and achieves fast capturing of the spread spectrum signals with the huge Doppler frequency offset and the dynamic range.

A launch system comprises a nose section comprising a nose coupling surface, a tail section comprising a tail coupling surface facing the nose coupling surface and a mast coupling the nose and tail sections. The mast is configured to expand and retract to displace the nose and tail sections within a range of distances from one another. In a retracted state, the nose and tail sections are either structurally coupled to one another at the nose and tail coupling surfaces or structurally coupled to at least one integrated module located between the nose and the tail sections.

A relative attitude and orbit coupling control method for a rendezvous and docking final approximation section aims at particularity of a space rendezvous and docking task, inherits advantages of a preset performance control method, and utilizes approximation capability of a radial basis function neural network (RBFNN) to compensate saturation nonlinearity of an execution mechanism in an actual rendezvous and docking task. The whole controller design process does not need model priori knowledge such as saturation nonlinear structural characteristics of the execution mechanism, external interference and orbiter mass rotational inertia, the calculation complexity can be greatly reduced, meanwhile, the control saturation constraint of the execution mechanism is considered while the preset steady-state and transient-state performance is considered, and the method has high engineering practicability.

The invention relates to a six-degree-of-freedom air bearing table disturbance torque determining method. The method is used for measuring inertia features and disturbance torques of an orbiter and ascender simulation device in a moon weak impact rendezvous and docking ground full-physical simulation test. According to the invention, a method used for measuring disturbance torques at various kinds of directions of a six-degree-of-freedom air bearing table is provided for the first time. According to the method, an existing rotational inertia measuring device can be improved, and rotational inertia measuring time can be shortened; and a mature attitude angle measuring means is used in combination, so that a complete disturbance torque measuring method can be constructed; the essential characteristics of an air bearing ball are utilized, and iterative calculation is performed on actual measurement data, and therefore, the rotational inertia of a system can be accurately measured, and the accuracy of spacecraft ground dynamics simulation can be improved.

A moving platform position measurement method is used for dynamic measurement of horizontal inclination angles of orbiter and ascender simulating devices in an all-physic simulation test of weak compact intersection abutting between a moon and the ground. The invention brings forward highly dynamic measurement of accurate position and attitude information of a moving platform on the basis of a CCD scan principle based on a linear array.