13 results about "Body axis" patented technology

The wide-speed-range tailless variable-structure aircraft comprises an aircraft body, the aircraft body comprises an aircraft body-fixed wing fusion body and variable-structure wing rudders, the variable-structure wing rudders are symmetrically arranged on the two sides of the aircraft body-fixed wing fusion body, and a wing body fusion body layout is integrally formed; the fuselage-fixed wing fusion body is connected with the variable-structure wing rudder through a driving mechanism, and the variable-structure wing rudder can deflect in the axial direction of the fuselage; the tail of the fuselage-fixed wing fusion body is provided with a trailing edge rudder, the trailing edge rudder of the fuselage-fixed wing fusion body achieves pitching channel control through same-direction deflection, and under the condition that the channel coupling coefficient is smaller than 1.5, rolling channel control is achieved through differential deflection. The problem that the yaw channel operation stability characteristic is insufficient under the tailless layout condition is solved, and the wide-speed-range operation performance of the aircraft is effectively improved; compared with other variable-structure aircrafts in which additional variables are introduced, the combination of the variable-structure wing rudder and the trailing edge rudder is simple and efficient, and the burden on a control system is small.

The application provides a kind of jet flight control wind tunnel rudder effect test system and test method, belong to wind tunnel test technical field.Test system includes aircraft model and be set in the engine and air induction system of aircraft model interior, fuel tank, jet actuation system, flight control computer and balance;Establish the body axis coordinate system corresponding to aircraft model, the origin of body axis coordinate system is model moment reference point, balance is located in the lower region of aircraft model, the axis of balance is parallel with the x axis of body axis coordinate system, and the heart of balance is located on the z axis of body axis coordinate system, engine is located in the tail region of aircraft model, the thrust line of engine overlaps with the x axis of body axis coordinate system, and air induction system is connected with jet actuation system.In the application, airborne air supply and control are realized, jet is taken as a part of aircraft system energy, and the accuracy of the result obtained by test on the influence of jet on the comprehensive performance of aircraft is improved.

The application discloses a normal vector-based follow-sighting attitude maneuver path planning method, and belongs to the technical field of spacecraft attitude control, and comprises the following steps: at any time during the follow-sighting maneuver, a camera optical axis is aligned with a target, and a satellite body axis perpendicular to the camera optical axis is set to point to a set normal vector direction, the set normal vector direction is set according to maneuver capability and pointing requirement, and the setting basis is that the body axis is in the set normal vector direction; the set normal vector direction is determined according to the follow-sighting task, and includes four cases, namely, keeping data transmission, keeping energy, keeping maneuver and keeping imaging. The normal vector of the plane formed by the target pointing and the ground pointing is applied, and the target attitude setting during the antenna data transmission to the ground while the follow-sighting is realized.

The invention provides a power system of a single-person jet aircraft, the power system comprises a plurality of fixed engines and a plurality of tilting engines, the fixed engines are fixedly arranged on a force bearing frame, and the tilting engines are rotatably arranged on the force bearing frame; the fixed engine nozzle is obliquely arranged outwards relative to the Z axis of the body axis coordinate system of the aircraft; a nozzle of the tilting engine in a neutral state is obliquely arranged relative to a Z axis of a body axis coordinate system of the aircraft, and component forces of thrust of the tilting engine relative to all directions of the aircraft are changed by changing a tilting angle of the tilting engine relative to the aircraft, so that the attitude of the aircraft is adjusted; therefore, the rolling, pitching and yawing three-axis angular motion and the vertical axis total thrust of the aircraft can be controlled without adjusting the speed / thrust of the tilting engine, and then the aircraft can be controlled to move along the three axes; when the attitude of the aircraft is changed, the phenomenon of falling down or decelerating is avoided.

This invention discloses a thrust vector control modeling method for a multi-engine parallel aircraft, relating to the field of launch vehicle flight control technology. The method includes: for each engine of the multi-engine parallel aircraft, obtaining the rotation angle from the engine's local yaw coordinate system to its orthogonal inertial coordinate system; combining the vertical yaw angle command and the lateral yaw angle command to determine the transformation matrix corresponding to each engine; obtaining the first thrust of each engine; using the transformation matrix to calculate the second thrust of each engine after yaw in the aircraft's body axis system; obtaining the coordinates of the yaw point of each engine in the aircraft's body axis system; and determining the control force and control torque generated by the yaw motion of all engines based on the second thrust and the yaw point coordinates. This invention simplifies the six-degree-of-freedom dynamics modeling and servo command allocation modeling of rockets, improving model readability and versatility.

A support head (400) for a blank (80) for use in helical flight forming apparatus (210), the blank having opposed main faces, an outer peripheral edge, a central hole having an inner peripheral edge, and a split extending from the outer peripheral edge to the inner peripheral edge so as to provide for opposed side edge sections. In one embodiment the support head (400) comprises: a main body part (410) having a main body axis (C-C) extending in a direction from one end thereof, a holder carriage part (416) for use with or including a blank holder having a blank receiving slot with a slot axis and gripper sections which contact the opposed main faces in the region of the side edge sections, when the blank is in an installed position, the holder carriage part (416) being operatively connectable to the main body part (410) in an adjustable fashion so that slot axis (D-D) can be inclined relative to the main body axis (C-C) in a number of different angular positions. In another embodiment the support head (500) comprises: a main body (512) which includes a first part (515) and a second part (516), the first and second parts being operatively connected together so is to the movable between an open position and a closed position in the direction of a displacement axis.

The application relates to a method for realizing variable-rate rotary motion of a tethered support airplane model, belonging to the technical field of wind tunnel testing. First, an 8-tether redundant constraint tethered support airplane model is adopted; second, in a three-dimensional rotary space, a relationship between a roll angle velocity and a pitch angle velocity is established, so that a angle of attack change rate is zero; third, a relationship between the roll angle velocity and a yaw angle velocity is further established, so that a side slip angle change rate is zero; subsequently, given an initial angle of attack and an initial side slip angle, a three-rotary-degree-of-freedom fixed coupling ratio motion is completed; then, three rotary angle velocity components around the body axis are synthesized to obtain a total rotary angle velocity, that is, an angle velocity of variable-rate rotation around a velocity axis; finally, according to a kinematic relationship, inverse solution of the airplane model position is obtained to obtain a tether length, which is used for motion control, so as to realize variable-rate rotary motion of the airplane model around the velocity axis. The method can realize rotary motion of the tethered support airplane model.

The invention discloses a rotor rotational inertia estimation method based on autorotation gliding deceleration pull control, and relates to the technical field of rotor aircrafts, and the method comprises the steps: simplifying a helicopter into a two-dimensional mass point model, and carrying out the stress analysis of a ground shafting helicopter, so as to obtain a helicopter motion equation; obtaining the relation between the helicopter ground axis speed and the body axis speed; calculating stress of a helicopter body based on panel hypothesis, and substituting the stress into a helicopter motion equation to solve rotor pulling force and an included angle between the pulling force and a vertical plane; the rotor pulling force is expressed as a dimensionless form, and an expression of a torque coefficient required by the rotor is obtained; the total inflow is solved through an algebraic method, the total inflow is substituted into the rotor wing required torque coefficient expression, and the rotor wing required torque coefficient is calculated; and calculating the rotational inertia estimated value of the rotor system based on the moment of momentum theorem. According to the method, the rotational inertia of the rotor system can be estimated by virtue of existing airspeed and ground speed measuring equipment, an angle gauge and a rotating speed indicator or a rotating speed sensor on the helicopter without dismounting the blades.

The application discloses a design method of a novel pre-compressor wing of a hypersonic air-breathing aircraft, and relates to the technical field of aerodynamic shape design of the hypersonic air-breathing aircraft. The method is used to solve the problem that the existing aircraft internal and external flow integration research still focuses on local structure and flow field simulation optimization, and the wing aerodynamic shape development through internal flow index simulation data optimization is still less. The wing pre-compression angle includes the variation characteristics of the internal flow performance with the angle of attack and the reference lift-drag ratio, and the wing pre-compression angle is expressed as θ pre‑c , θ pre‑c is the angle difference between the wing chord line and the body axis on the longitudinal reference XOY plane, the wing shape includes a near-body side curve, a far-body side curve and a trailing edge curve, the wing section includes a thickness distribution and a blunt leading edge section shape, the reference lift-drag ratio includes an initial lift-drag ratio and a target lift-drag ratio, and the blunt leading edge section shape includes a blunt leading edge section profile line and a prediction of the compensation effect of the blunt leading edge section profile on the lift-drag ratio performance.

The present invention relates to a microrobot (10) configured to move along a propulsion direction within a viscoelastic anatomical environment, the microrobot (10) comprising: a body (12) extending along a body axis (X); a navigation head (14) extending along a head axis (H), the navigation head (14) assuming a given configuration relative to the body; a propulsion element (16) presenting an outer surface (18) having a helical external thread (20) extending along the propulsion axis (A) and assuming a given configuration relative to the body; and a drive means (22) configured to rotationally drive the propulsion element about the propulsion axis (A). The microrobot (10) further comprises an orientation device (24) configured to change at least one of the given configuration of the navigation head and the given configuration of the external thread relative to the body to change the propulsion direction.

This invention discloses a method for reducing the nutation axis deflection and nutation angle of a space-borne spinning object based on a solid rocket motor. The method includes: real-time measurement of the pitch angle, yaw angle, and roll angular velocity of the space load; integration after coordinate transformation to obtain the attitude angles of the space load in the ECI coordinate system, obtaining the transformation matrix; calculating possible nutation axis direction vectors; averaging the results to obtain the optimal nutation axis direction vector; and calculating the plane... YOUR With plane NOZ The clockwise angle, the angle between the nutation axis and the target axis, the angle between the body axis and the target axis, and the nutation angle are calculated; it is determined whether to enter the deflection step, the delayed ignition time and the expected ignition direction of the solid rocket motor are calculated; the corresponding solid rocket motor is started for deflection; the plane is calculated. YOUR With plane NOZ The clockwise angle is used to calculate the delayed ignition timing and the expected ignition direction of the solid rocket motor. This invention improves the controllability and maneuverability of spin-stabilized space payloads, laying the technological foundation for modern deep space exploration.

The application discloses a high-universality fixed-wing unmanned aerial vehicle air-based launching and recovering device, which is integrally installed on an air-based platform, and unmanned aerial vehicles are guided to complete docking through a four-cone-rod guiding and centering mechanism, a funnel-shaped butt joint area formed by four carbon tubes, and electromagnets fixed by supports arranged at front and back positions of the four-cone-rod guiding and centering mechanism and matched with electromagnets installed on the unmanned aerial vehicles to realize adsorption and fixation of the unmanned aerial vehicles during butt joint. A mechanical claw is designed in front of the front support, the mechanical claw clamps the nose of the unmanned aerial vehicle after fixation of the unmanned aerial vehicle, the unmanned aerial vehicle is further fixed, deviation of the unmanned aerial vehicle is corrected, and the unmanned aerial vehicle and the air-based platform have their body axes coplanar. The application can guarantee the accuracy and stability of recovery under certain wind disturbance conditions, the whole device is modularly designed, different locking mechanisms can be arranged according to types of carried sub-machines, and the device is convenient and fast to iterate, and has good universality and expansibility.

The invention provides a jet flight control wind tunnel steerage test system and test method, and belongs to the technical field of wind tunnel tests. The test system comprises an airplane model, an engine, an air entraining system, an oil tank, a jet flow actuating system, a flight control computer and a balance, a body axis coordinate system corresponding to the airplane model is established, the original point of the body axis coordinate system is a model torque reference point, the balance is located in the lower area of the airplane model, the axis of the balance is parallel to the x axis of the body axis coordinate system, the calibration center of the balance is located on the z axis of the body axis coordinate system, and the engine is located in the tail area of the airplane model; the thrust line of the engine is overlapped with the x axis of the body axis coordinate system, and the air entraining system is connected with the jet flow actuating system. Airborne air supply and control are achieved, the jet flow serves as a part of aircraft system energy, and the result accuracy of the influence of the jet flow obtained through the test on the comprehensive performance of the aircraft is improved.