17 results about "Slipstream" patented technology

The application discloses a kind of based on interference speed model's coaxial dual-rotor hover aerodynamic characteristic analysis method, belong to helicopter aerodynamics field.The application is by establishing the speed model of mutual interference of upper and lower rotors, the iteration procedure is used to solve the spanwise distribution of interference speed.For the interference of upper rotor to lower rotor, the non-uniform interference speed is calculated using improved Landgrebe slipstream model;For the interference of lower rotor to upper rotor, the uniform interference speed is calculated using McAlister slipstream model.The application is by comparing slipstream radius and rotor radius, determines interference range, and utilizes piecewise function to describe the non-uniformity of interference speed, can simultaneously process equal radius and unequal radius coaxial dual-rotor configuration, fully considers the mutual aerodynamic interference between upper and lower rotors, overcomes the limitations of existing theoretical analysis method in slipstream model and interference speed calculation, realizes the fast, accurate analysis of coaxial dual-rotor hover aerodynamic characteristics.

In some embodiments, a lift augmentation system for a blown lift aircraft includes a blown lift tailplane operatively coupled to the blown lift aircraft. The blown lift tailplane may include a leading edge and a trailing edge, an upper surface and a lower surface, and a first side and a second side. The lift augmentation system may include one or more tailplane thrust-producing devices on the first side and the second side of the blown lift tailplane operatively coupled to the leading edge of the blown lift tailplane. The one or more tailplane thrust-producing devices on the first side and the second side of the blown lift tailplane may produce a plurality of slipstreams corresponding to each of the tailplane thrust-producing devices. The plurality of slipstreams corresponding to each of the tailplane thrust-producing devices may blow over the upper surface and the lower surface of the blown lift tailplane.

The present application relates to the field of aircraft flow control, and particularly discloses a propeller aircraft and a high-speed jet flow control method under the influence of slipstream, the high-speed jet flow control method comprising: obtaining a current rotating speed r of the propeller; based on the current rotating speed r, adjusting a jet flow momentum coefficient of a slipstream influence area to C μ1 , adjusting a jet flow momentum coefficient of a non-slipstream influence area to C μ2 , wherein C μ1 +C μ2 =C μt , and C μ1 >C μ2 , C μt is a total momentum coefficient; by adjusting the jet flow momentum coefficients of the slipstream influence area and the non-slipstream influence area, the jet flow momentum coefficient C μ1 of the slipstream influence area is higher than the jet flow momentum coefficient C μ2 of the non-slipstream influence area, the slipstream influence area adopts a larger momentum coefficient, the non-slipstream influence area adopts a smaller momentum coefficient, the jet flow momentum coefficients in the two areas are reasonably distributed by considering the slipstream influence of the propeller, and efficient utilization of the jet flow momentum is realized.

The embodiment of the invention provides a simulation method and device for aileron faults of a medium-sized propeller aircraft, computer equipment and a medium, and the method comprises the following steps: setting fault modes of ailerons, and setting a deflection rate proportion and control surface target skewness for each fault mode of the ailerons; the control surface skewness of each aileron is output through a control surface actuator model; estimating six-component aerodynamic parameters of a single aileron through the three-component aerodynamic parameters of the whole aileron; performing propeller slip flow asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters; according to the control surface skewness and the asymmetric six-component aerodynamic parameters of each aileron, the state parameters of the airplane at the next moment are obtained through simulation calculation of the state parameters of the airplane at the previous moment by means of the six-degree-of-freedom simulation calculation model of the airplane. According to the scheme, through analog simulation of the aileron fault, the problems that an existing aileron fault model is rough, and consistency of theoretical simulation and a real airplane model is poor are solved.

The invention discloses a coaxial dual-rotor hovering aerodynamic characteristic analysis method based on an interference speed model, and belongs to the field of helicopter aerodynamics. According to the method, the velocity model of mutual interference of the upper and lower rotors is established, and the spanwise distribution of the interference velocity is solved by adopting an iterative process. Aiming at the interference of the upper rotor wing to the lower rotor wing, an improved Landgrebe slipstream model is adopted to calculate a non-uniform interference speed; and aiming at the interference of the lower rotor wing to the upper rotor wing, a McAlister slip flow model is adopted to calculate a uniform interference speed. According to the method, the slip flow radius and the rotor radius are compared, the interference range is judged, the non-uniformity of the interference speed is described through the piecewise function, the equal-radius and unequal-radius coaxial dual-rotor configuration can be processed at the same time, mutual aerodynamic interference between the upper rotor and the lower rotor is fully considered, and the method is suitable for large-scale popularization and application. The limitation of an existing theoretical analysis method on slip flow model and interference velocity calculation is overcome, and rapid and accurate analysis of coaxial dual-rotor hovering aerodynamic characteristics is achieved.

The invention belongs to the technical field of airplane wind tunnel test model design, and discloses a middle fuselage structure of a large-size low-speed wind tunnel turboprop airplane test model and a design method thereof, the middle fuselage structure adopts a composite design of combining an internal steel skeleton and an external aluminum alloy cover plate, and the internal steel skeleton is made of 30CrMnSiA alloy steel and serves as a main bearing structure; the outer aluminum alloy cover plate is made of 7075 aluminum alloy, and the aerodynamic shape of the middle fuselage is formed. A structure for installing wind tunnel test equipment such as a slip flow test air bridge is arranged on the internal steel framework, and the problems that the internal space of a traditional large-size middle fuselage is small, and the equipment is difficult to install are solved. Meanwhile, the method further comprises the optimal design of detachable upper and lower cover plates, a balance, a supporting rod connecting piece, a split type transition rectification bag and the like, and on the premise that the structural rigidity and strength and the test precision are guaranteed, the machining difficulty is greatly reduced, and the machining period and cost are greatly reduced. The method is suitable for manufacturing and testing the large-size low-speed wind tunnel turboprop aircraft test model.

The invention relates to the technical field of aviation aircrafts, and discloses a lifting body aircraft and a method for keeping constant static stability thereof.The lifting body aircraft comprises a lifting body fuselage, the two sides of the rear portion of the fuselage extend backwards to be provided with side end plates, and a fan slip flow plate assembly capable of rotating upwards and forwards is connected between the two side end plates; two variable-sweep-angle wings are symmetrically arranged at the front upper part of the fuselage, the tops of wing tips of the variable-sweep-angle wings are rotationally connected with rotors, and the variable-sweep-angle wings can horizontally rotate around vertical rotating shafts at the roots; horizontal tails are symmetrically arranged on the outer sides of the two side end plates on the rear portion of the machine body, slidably connected with the side end plates and capable of moving front and back. Through the rotor wings, the wings and the fan slip flow plate assemblies, pulling (lifting) force and control torque for vertical and horizontal flight of the aircraft are provided; by adjusting the swept angles of the wings and the front and back positions of the horizontal tails, the resultant action point of the two rotors and the fan and the pneumatic focus of the whole aircraft are changed, so that the stability of the aircraft is kept constant under the condition of different use gravity centers, and the safety is ensured under the condition of not increasing the balance weight or reducing the load.

The present application proposes a high-efficiency and high-precision propeller slipstream numerical simulation method: firstly, the isolated propeller under actual working conditions is numerically simulated by using the frozen rotor method, the axial force and the circumferential force of the blade element are extracted by integrating the blade surface pressure and friction force at different radial positions, and the size and phase of the blade load are corrected by a semi-empirical method; then, the axial force and the circumferential force of the blade at the same radial position and different phases are taken as a group of fitting points, the fitting parameters based on the blade element momentum theory are obtained, and the non-uniform load distribution of the disc is predicted; next, the disc load is time-averaged, and the work of the propeller is converted into the axial and circumferential pressure increments of the excitation disc; finally, the pressure increments of the excitation disc are added to the momentum equation in the form of volume force source term, and then the slipstream effect is simulated. The method has both the calculation accuracy of the unsteady method and the calculation efficiency of the quasi-steady method, and has good engineering practicability.

The application discloses a distributed electric propulsion propeller slipstream wing lift measuring device, which comprises a base, a free flow power mechanism and a front-rear moving mechanism arranged on the base, a lifting mechanism arranged on the front-rear moving mechanism, a measuring mechanism arranged on the lifting mechanism, and an attack angle adjusting mechanism arranged on the measuring mechanism and used for fixing a wing to be detected; the power output directions of a plurality of free flow propellers of the free flow power mechanism are all directed to a fairing; the front-rear moving mechanism is used for adjusting the distance between the wing and the fairing; the lifting mechanism is used for adjusting the placing height of the wing; the measuring mechanism is used for monitoring the stress condition of the wing; and the attack angle adjusting mechanism is used for adjusting the placing angle of the wing; thereby, a propeller and wing aerodynamic coupling test device is built, the propeller slipstream wing lift increment can be more accurately obtained compared with the computational fluid dynamics method, and a large amount of effective data can be quickly and cheaply obtained compared with the wind tunnel test.

An aerodynamic system for lift augmentation and thrust recovery in propeller-driven aircraft. The system includes at least one auxiliary wing positioned within a slipstream of a propeller and coupled to a propulsion assembly via at least one structural pylon. The auxiliary wing is configured to generate supplemental lift from the high-velocity airflow of the slipstream, particularly during high-thrust operational phases. The structural pylon is configured as a flow-rectifying vane to deswirl a rotational component of the slipstream, thereby increasing net axial thrust. The system facilitates energy conservation and accelerates the transition between vertical and horizontal flight modes by utilizing recovered slipstream energy to augment fixed-wing lift generation.

The application discloses a modular reconfigurable and self-adaptive control tunnel unmanned plane and belongs to the field of unmanned planes. Specifically, the application comprises a minimum module unit of a single-rotor attached to a slipstream rudder, wherein the minimum module unit comprises a fuselage, a power assembly, a control assembly, a connecting assembly and a take-off and landing assembly; two module units are spliced into a minimum flight unit through connecting rods; a plurality of flight units can be connected in a head-to-tail mode or a side-by-side mode; the self-rotation torque is offset through the reverse rotation of the rotors on the two sides of the connecting rods; the pitch of the flight unit is controlled through the differential speed of the rotors; and the lateral change of the flight unit is controlled by the slipstream rudder surface. The application can assemble, disassemble and scale the aircraft according to the weight and size requirements of the load, and has the ability to complete the reconnaissance and transportation through small windows and long and narrow tunnels.

A system includes a bluff body having a tail end and a wake region and slipstream direction behind the tail end. The system further includes a propeller including a hub and a plurality of blades extending outward from the hub. The hub is mounted to the tail end of the bluff body for rotation about an axis. The axis is angled relative to the slipstream direction in both pitch and yaw such that when fluid is flowing over the bluff body, at least one of the blades is in the slipstream and extracts energy to rotate the hub, while at least one of the blades is out of the slipstream and redirects fluid from the slipstream into the wake region. Such autorotation of the hub causes the blades to continuously move in and out of the slipstream to continuously extract energy and redirect slipstream fluid into the wake region.

The invention provides a multi-propeller-wing coupling aerodynamic optimization design method based on an excitation disc model, and relates to the field of aerodynamic configuration optimization design, and the method comprises the steps: taking a multi-propeller-wing configuration as a reference configuration to be subjected to aerodynamic optimization design, and generating corresponding surface grids and volume grids; performing FFD parameterization processing on the surface grid of the component to obtain a corresponding FFD control body; iteratively updating the surface grid, and quickly updating the volume grid; solving a quasi-steady RANS equation based on the excitation disk model and the updated volume grid, and obtaining a flow field under slip flow interference; solving an adjoint equation of the original flow field equation, and solving gradient values of the target function to all design variables; and iteratively updating the design variable of the aerodynamic characteristics of the wing, and iterating the FFD control body corresponding to the surface grid until the function value of the aerodynamic optimization objective reaches a preset convergence criterion. According to the method, the aerodynamic configuration optimization design of the multi-propeller aircraft wing with relatively high efficiency and precision is realized.

The invention discloses an aircraft high lift device load calculation method considering propeller slip flow influence. The method comprises the following steps: S1, calculating a correction coefficient Ks of the influence of the slip flow of the propeller on the load of the high lift device based on the geometric parameters of the high lift device of the aircraft, the engine parameters of the turbine propeller and the flight state; s2, acquiring initial pressure distribution of the surface of the high lift device of the aircraft without being influenced by the slip flow; and S3, combining the correction coefficient with the initial pressure distribution, performing correction according to the slip flow influence area to obtain the pressure distribution of the surface of the airplane high lift device after slip flow influence, and calculating the aerodynamic load of the airplane high lift device according to the pressure distribution. According to the method, the slipstream influence of the turboprop engine is accurately calculated, the effect of the propeller effect on the high lift device is quantified, and the accuracy and design safety of load evaluation of the high lift device of the airplane are improved.

The application provides a propeller aircraft polar curve determination method, introduces a tension coefficient term in a general propeller aircraft polar curve model, represents the influence of propeller slipstream on the resistance coefficient, establishes a general polar curve model which can be simultaneously applied to the cruising, climbing and descending states of the propeller aircraft, and then determines the expression of the general polar curve model by using the cruising, climbing and descending test flight data of the propeller aircraft, so that the same configuration can be simultaneously applied to the flight performance expansion calculation of the propeller aircraft in different flight states, the efficiency and accuracy of the propeller flight performance calculation can be improved, and effective guidance can be provided for the rapid iteration of the design and improvement of the propeller aircraft.

The invention provides a high lift propeller propulsion unit design method suitable for distributed propulsion layout, and belongs to the technical field of aviation new energy power. According to the method, the mathematical relationship between the axial slip flow speed increment of the propeller and the blade element tension is established, and the distribution of the axial slip flow speed is controlled through the distribution of the tension along the radial direction of the blade, so that the lift augmentation effect of a downstream wing is strengthened. By adopting the design method provided by the invention, a novel high-lift propeller propulsion unit with a given flight height H of 0-500m, a flight speed of 33m / s, a propeller design tension of 230N, a design rotation speed of 5000rpm and a propeller disc radius of 280mm is completed. According to the novel propeller design method, the pulling force / pushing force requirements of the aircraft and the lift augmentation effect of the slipstream on the downstream wings can be considered, on the premise that the pulling force / pushing force requirements of the aircraft are met, the lift augmentation effect of the downstream wings is enhanced by changing the slipstream form of the propeller, and then propelling-pneumatic comprehensive benefits are obtained.

The invention relates to the field of aircraft flow control, and particularly discloses a propeller aircraft and a high-speed jet flow control method under the influence of slip flow, and the high-speed jet flow control method comprises the following steps: obtaining the current rotating speed r of a propeller; on the basis of the current rotating speed r, the jet flow momentum coefficient of the slipstream influence area is adjusted to be C [mu] 1, the jet flow momentum coefficient of the non-slipstream influence area is adjusted to be C [mu] 2, C [mu] 1 + C [mu] 2 = C [mu] t, C [mu] 1 is larger than C [mu] 2, and C [mu] t is the total momentum coefficient; by adjusting the jet flow momentum coefficients of the slipstream influence area and the non-slipstream influence area, the jet flow momentum coefficient C [mu] 1 in the slipstream influence area is higher than the jet flow momentum coefficient C [mu] 2 in the non-slipstream influence area, the slipstream influence area adopts a large momentum coefficient, and the non-slipstream influence area adopts a small momentum coefficient. The jet flow momentum coefficients in the two areas are reasonably distributed by considering the slip flow influence of the propeller, and efficient utilization of the jet flow momentum is achieved.